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: Aircraft flight path changes commanded by a pilot via the wheel, column and pedal (12) are converted to position sensor signals (14) and passed as the input to primary flight computer (26). The primary flight computer (26) converts these pilot inputs and the inputs from autopilot (25) into desired surface actuator commands and then transmits them to actuator control electronics (18). The actuator control electronics (18) also receives a position feedback signal, representative of the position of the aircraft control surface (42). The actuator control electronics (18) produces a control signal which is fed to the input of an actuator (32) which includes a hydraulic system (34). The actuator responds to control input signals to drive linkage (40) which then positions the control surface (42).

Abstract: A flight control system according to the present invention includes a vertical acceleration control device for calculating a pitch axis steering angle command to make a difference between a vertical acceleration of an airplane and a target turn acceleration to zero, and transmitting it as a variable to a pitch axis control device, a reference bank angle device for calculating a reference bank angle from the target turn acceleration, an altitude control device for calculating a bank angle correcting quantity from a difference between an altitude of the airplane and a target altitude and obtaining a bank angle command by correcting the reference bank angle, and a roll axis control device for calculating a roll axis steering angle command for make a difference between a real bank angle and the bank angle command to zero and for transmitting it as a variable to a roll axis control device.

Abstract: Disclosed is a pitch-axis stability and command augmentation system (19) in which a pilot column input (.delta..sub.C) is provided to a pitch command processor (26), the output of which is a C*U feedforward command (C*U.sub.FFC) that is supplied as an additive input to a combining unit (20). The combining unit (20) receives a second additive input of an augmented feedback command signal (AFB.sub.COM). The resulting output is filtered and generates an elevator command signal (.delta..sub.e,FILT). The command processor (26) additionally supplies a corrected column position signal (.delta..sub.C,COR) to a pitch command C*U processor (26) that converts the corrected column position signal (.delta..sub.C,COR) into a C*U pitch command (C*U.sub.PilotCmd), representative of the pilot's requested elevator pitch which is generated by movement of the control column. A computed C*U processor (32) forms a computed C*U signal (C*U.sub.Computed) that is based on the current state of the aircraft.

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

Abstract: The present invention relates to a device for controlling the thrust of an aircraft, with several engines (M1, M4), comprising calculation units (UC1, UC4) that formulate first commands for controlling the engine speeds. According to the invention, said device (1) additionally comprises means (CONT1, CONT4) of controlling the two outer engines (M1, M4), calculation means that formulate second commands for controlling the speed of the outer engines (M1, M4) making it possible to obtain a reduction in thrust, and switching means which, during normal operation of the outer engines (M1, M4) send the first commands to these engines and, when one outer engine (M1) fails, send the second commands to the opposite outer engine (M4).

Abstract: A system for controlling the tab (6) of an aircraft control surface (3) having a pair of position sensors (12 and 15) that supply information regarding the turning of the control surface (3) and the position of a jack (13), respectively, on the basis of signals generated by the sensors (12 and 15) as well as possibly on the basis of at least one parameter (p) originating from the aircraft, such as airspeed, positions of lift-augmenting devices, forces exerted by the pilot on a control, etc., and which formulates a command for the jack (13).

Abstract: Practical application of real-time (or near real-time) Adaptive Performance Optimization (APO) is provided for a transport aircraft in steady climb, cruise, turn descent or other flight conditions based on measurements and calculations of incremental drag from a forced response maneuver of one or more redundant control effectors defined as those in excess of the minimum set of control effectors required to maintain the steady flight condition in progress. The method comprises the steps of applying excitation in a raised-cosine form over an interval of from 100 to 500 sec. at the rate of 1 to 10 sets/sec of excitation, and data for analysis is gathered in sets of measurements made during the excitation to calculate lift and drag coefficients C.sub.L and C.sub.D from two equations, one for each coefficient. A third equation is an expansion of C.sub.

Abstract: A method and apparatus for use with an aircraft autopilot to prevent an aircraft part from striking the ground during near ground maneuvers by employing a protection circuit between the outer loop and the inner loop of the control chain between the autopilot and the control surface, the prevention circuit producing an output signal when there is a possibility of ground strike which output signal operates to reduce the control surface command signal from the inner loop to the control surface and thus be rapidly responsive to aircraft attitude changes that could produce ground strike.

Abstract: A monitoring system for detecting failures, or impending failures, in subsystems of a fly-by-wire primary flight control system of an aircraft. The system monitor a critical variable of at least two controllers acting in concert, in real time. Differences between the critical variables are calculated and, after filtering to remove insignificant variations, structural fatigue damage is estimated based on significant detected variations. Depending upon the relative severity of the estimated fatigue damage, the aircraft's primary flight control computer is programmed to shut down the malfunctioning subsystem, display a message indicating need for repair, or take other appropriate action.

Abstract: A method and an apparatus are provided for optimizing the aerodynamic effect of the airfoil of an aircraft by defined changes in camber. The method includes the following steps:a. determining the flow for the flight condition caused by the change in camber,b. comparing the ascertained characteristic values with stored nominal reference values for an optimal flow,c. forming differential values between the characteristic values and the stored nominal reference values,d. deriving actuator signals from the differential values, ande. changing the camber by motor, based on the actuator signals, for minimizing the differential values.The optimum wing flow is thereby maintained more exactly. For transonic wings, the position and strength of compression shocks is also effectively controlled, which leads to a reduction of the direct shock induced separation.

Abstract: An aircraft modal suppression system which recognizes that the frequency and phase of the body bending mode varies when the weight of the aircraft differs from the design gross weight. An active damper notch filter which is tabulated as a function of aircraft gross weight is utilized, thereby enabling not only the frequency, but also the width and depth of the notch filter to vary according to the gross weight of the aircraft.

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

Abstract: An aircraft overspeed protection system produces proportional and integral commands for input to an autopilot controlling the aircraft. The proportional and integral commands are produced by comparing actual monitored speed of the aircraft with a target speed from a target speed selector. The target speed selector selects as the target speed a trigger speed above a nominal maximum operating airspeed of the aircraft until the trigger speed is reached by the aircraft. When the trigger speed is reached by the aircraft, the overspeed control goes into Overspeed Protect Command active mode and the target speed selector selects a new target speed below the nominal maximum operating speed of the aircraft. The overspeed protection system remains in Overspeed Protect Command active mode until the pilot takes a positive action by selecting a new autopilot mode or by disengaging the autopilot.

Abstract: An aircraft modal suppression system which recognizes that the frequency and phase of the body bending mode varies when the weight of the aircraft differs from the design gross weight. An active damper notch filter which is tabulated as a function of aircraft gross weight is utilized, thereby enabling not only the frequency, but also the width and depth of the notch filter to vary according to the gross weight of the aircraft.

Abstract: A self adaptive limiter for use in aircraft control systems during approach and landing is disclosed. Estimated flight path angle is continuously computed during the glidepath tracking phase until a predetermined lock altitude above ground when a nominal flight path angle is latched. Nominal vertical speed is continuously computed below the lock altitude as a function of the latched nominal flight path angle and ground speed of the aircraft. A vertical speed limit function is generated as a function of the nominal vertical speed and radio altitude. During approach and landing a pitch limit is computed from the vertical speed limit, vertical speed, ground speed, and pitch. When a pitch command to the autopilot exceeds the pitch limit(i.e. commands excessive pitch down attitude), it is limited to the pitch limit thus preventing the aircraft from descending below a safe altitude.

Abstract: The invention provides a method for controlling the precession of a spinning spacecraft (20) which allows the spacecraft body to respond to an input torque without the nutation normally attendant when an input torque is applied about one transverse axis to accelerate a spinning spacecraft about that one axis. Dynamic decoupling eliminates nutation through the impression of additional derived feedback torques (44,46) to the input torque control of a spinning spacecraft to oppose or cancel the intrinsic cross-coupling terms (34,36) of the spinning spacecraft's gyrodynamics that give rise to the nutation.

Abstract: A fault tolerant actuation system (12) for flight control systems is provided. The fault tolerant actuation system (12) includes a plurality of primary flight computers (14a, 14b, and 14c) with corresponding power control units (24a, 24b, and 24c). Each power control unit includes a remote electronic unit (18a, 18b, and 18c), an electro-hydraulic servo valve (26a, 26b, and 26c), and an actuator (28a, 28b, and 28c). The actuators are linked to a flight control surface (30) to control its position. The electro-hydraulic servo valves (26a, 26b, and 26c) and the actuators (28a, 28b, and 28c) include sensors that monitor their operation. Each RE (18a, 18b, and 18c) generates a control current (i.sub.1, i.sub.2, and i.sub.3) based upon commands of the corresponding primary flight computer as well as feedback data transmitted from the sensors of the corresponding electro-hydraulic servo valve and actuator only. The feedback data is transmitted along separate servo loops having separate compensations (66a, 70a).

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

Abstract: A rudder bar system for a helicopter controlled in yaw by acting on the tail rotor or an equivalent device has a capability, as a function of the parameters representative of the current flight status of the helicopter, of either automatically continuously recentering the forces on the rudder bar, which cancels out the residual static forces and gives the pilot a tactile sensation close to that given by a friction-type rudder bar, or allowing the pilot fully to feel the countering action of an elastic return device such as a spring.

Abstract: A vertical performance limit compensator is provided for a rotary wing aircraft. The compensator modifies a vertical velocity command signal for the aircraft's autopilot. The vertical velocity command signal represents a desired vertical velocity for the aircraft, and results in an associated engine power demand. Existing engine power remaining beyond that required for level flight is excess engine power. An engine performance signal is provided as an indication of an engine power demand that exceeds the excess engine power. A compensated vertical velocity command signal having a magnitude lower than the vertical velocity command signal is provided in response to the vertical velocity command signal and the presence of the engine performance signal.

Abstract: Disclosed is a pitch-axis stability and command augmentation system (19) in which a pilot column input (.delta..sub.C) is provided to a pitch command processor (26), the output of which is a C*U feedforward command (C*U.sub.FFC) that is supplied as an additive input to a combining unit (20). The combining unit (20) receives a second additive input of an augmented feedback command signal (AFB.sub.COM). The resulting output is filtered and generates an elevator command signal (.delta..sub.e,FILT). The command processor (26) additionally supplies a corrected column position signal (.delta..sub.C,COR) to a pitch command C*U processor (26) that converts the corrected column position signal (.delta..sub.C,COR) into a C*U pitch command (C*U.sub.PilotCmd), representative of the pilot's requested elevator pitch which is generated by movement of the control cola. A computed C*U processor (32) forms a computed C*U signal (C*U.sub.Computed) that is based on the current state of the aircraft.

Abstract: Process for piloting an aircraft in order to improve a microgravity state and the corresponding system.Instead of supplying the pilot with an information linked with the instantaneous position of the equipment in free floating form (M), he is supplied with an information relative to the anticipated position of said equipment.Application to microgravity studies.

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 failure of any one of three swashplate actuators of a helicopter rotor blade is detected. Once such a detection is made, the position of this nonfunctional swashplate is locked and measured. The inputted commanded swashplate collective position, commanded swashplate x-axis rotational position, and commanded swashplate y-axis rotational position are then passed to a failure-mode control matrix. The failure-mode control matrix computes swashplate actuator commanded positions for the two operable swashplate actuators so that aircraft attitude control is maintained. These two swashplate actuator commanded positions instruct positional movement of the rotor blade swashplate to thereby meet the commanded swashplate x-axis rotational position and the commanded swashplate y-axis rotational position which will control attitude. A quasi-swashplate-collective-position corrector computes the quasi-swashplate collective position that will occur because of the successful control of the aircraft's attitude.

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

Abstract: An aircraft modal suppression system which recognizes that the frequency and phase of the body bending mode varies when the weight of the aircraft differs from the design gross weight. An active damper notch filter which is tabulated as a function of aircraft gross weight is utilized, thereby enabling not only the frequency, but also the width and depth of the notch filter to vary according to the gross weight of the aircraft.

Abstract: A flight control system that provides energy management for an aircraft. The flight control system includes an energy control system which generates a first error signal that is a function of an actual energy characteristic and a threshold energy characteristic. The energy characteristic may be the bleed rate of the aircraft. The error signal is converted into a control surface command signal which moves the control surface(s) of the aircraft, accordingly. The control system limits the movement of the control surface(s) so that the actual energy characteristic does not exceed the threshold energy characteristic. The energy control system can be coupled to a nominal control system which generates a second error signal that is a function of a pilot input command and a feedback signal which is indicative of the present state of the aircraft. The second error signal can also be converted to a control surface command signal which moves the control surface(s) of the aircraft.

Abstract: Apparatus for detecting and signaling a skewing or misalignment of adjacent aircraft leading edge slats is disclosed. A cable is attached to an actuator having a compression spring system and located in an outboard slat. The cable passes through cable guides in several adjacent slats before being attached to an inboard slat. The compression spring system utilizes a dual concentric pair of compression springs for maintaining a tight cable. When a misaligned condition is detected by a proximity switch, the increased cable load will cause the actuator to lock itself in a position out of the range of the proximity sensor.

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: Segmented spoilers are parallel arrays of individually-extended barrier surfaces, such as rotatable eccentrically-mounted disks. The overlapping surface areas extend through a slot on a aircraft surface. When actuated, a segmented array generates a stiff, extendable, profiled spoiler-barrier. The individual surface areas are power-activated from below the airfoil surface according to motor commands from autopilots, operators, sensors and computers. Disk spoiler systems provide very rapid generation and retreat of controllable height barriers. The management of Bernoulli lift phenomenon with disk spoilers has unique use on an aircraft's nose, along the top of its wings, on the forward surfaces of horizontal and vertical stabilizers and within the intake sections of gas turbine aircraft engines. Disks are rotated by electric rotor-positioning motors and aircraft-powered axial force systems.

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 turn coordination inhibit system (150) inhibits a rotary winged aircraft control system from operating in an automatic turn coordination mode when a pilot desired to perform a sideslip maneuver, e.g., a flat turn. When automatic turn coordination is not engaged (132,212,215), e.g., the aircraft is not in a coordinated turn, and either aircraft bank angle (119) exceeds an inhibit threshold magnitude (210) or a pilot yaw command provided by a pilot sidearm controller (155) exceeds a minimum threshold value (243), e.g., the sidearm controller is out of detent in the yaw axis, automatic turn coordination is inhibited (152). Automatic turn coordination remains inhibited until both aircraft bank angle falls below a reset threshold magnitude (230) and the sidearm controller is back in the detent position for the yaw axis (243).

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 rotatable slotted cylinder (RSC), partially embedded within the contours of a tail surface, such as an aircraft tail airfoil, serves as a force-producing element in a closed-loop active control system for buffet alleviation. A longitudinal axis of the RSC runs spanwise to the airfoil at or near the three-quarter chord location. In a so-called "home" position, the RSC projects as two small spanwise humps out of opposite sides of the tail surface. By active feedback control using a buffet response signal measured by an accelerometer, the RSC rotates up to .+-.45.degree. maximum deflection from the home position, thus allowing free stream air to flow through the airfoil, thereby creating lift forces for the active alleviation of the buffet response on the tail surface. An alternate embodiment of the invention places the RSC and a drive motor assembly outside of and adjacent to the tail surface near the airfoil quarter chord station.

Abstract: According to the invention, attenuator means (8) are provided, making it possible to limit the rate of rotation of the airplane on takeoff, so as to avoid the tail of the airplane striking the ground and to reduce the dispersion of the rotation rates.

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: The present invention relates to an automatic detector for a control device for controlling the flying state of a remote-control toy airplane, which can provide automatic correction to maintain the remote-control toy airplane in a flying state in a level or horizontal manner. The detector is a digital or analogic device located at a proper position within the housing of the airplane. The detector includes a device locating plate, fixing racks, a suspension cord, a movable light-obstructing sheet or a movable light-permeable sheet, light emitting element, and one or more detecting elements the light-obstruction sheet is swingingly mounted on the suspension that in turn is connected to the cord fixing racks. When light emitting element emits light and the airplane maintains a horizontal or level attitude, the effect of gravity on light-obstruction sheet will cause it to obstruct light.

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: An improved autopilot system of the type which uses pitch commands to control airspeed with an improvement of including a temporary pitch hold command being issued if thrust changes would oppose the autopilot pitch command. The temporary pitch hold command would end when the airspeed reaches a calculated airspeed capture point.

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

Abstract: 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: An aircraft modal suppression system which recognizes that the frequency and phase of the body bending mode varies when the weight of the aircraft differs from the design gross weight. An active damper notch filter which is tabulated as a function of aircraft gross weight is utilized, thereby enabling not only the frequency, but also the width and depth of the notch filter to vary according to the gross weight of the aircraft.

Abstract: The attitude of a spinning spacecraft (20) whose spin axis is substantially in the plane of the orbit is controlled without the use of reaction control thrusters. A two-axis gimbal (24) on which a momentun wheel (26) is mounted is secured to a central body (21). Two actuators (40, 42) are used to selectively pivot the gimballed momentum wheel (26) about each gimbal axis (x, y) in order to apply a control moment to change the attitude state of the spacecraft (20).

Abstract: In a stability compensating mechanism of an electro-hydraulic servo system which includes a load-displacement detecting unit for detecting a displacement of a piston of actuator or a displacement of a load connected to the piston, and a control valve for supplying a hadraulic oil to the actuator to drive the load and in which a piston-displacement detection signal of the load-displacement detecting unit is fed as a feedback signal back to a forward circuit and the control valve is operated by a deviation signal between the feedback signal and a command signal, the stability compensating mechanism comprising a valve-displacement detecting unit for detecting a valve displacement of the control valve and outputting a valve-displacement detection signal representative of the valve displacement, and a load-pressure computing unit for computing a load pressure that is acted on the actuator due to the load, on the basis of the piston-displacement detection signal of the load-displacement detecting unit and the valve

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 aircraft autoland flare control system which augments the autopilot's aircraft flare control commands in the event the aircraft exceeds defined flare envelopes, monitors aircraft longitudinal position, altitude, altitude rate, lateral position and lateral position rate during flare, and outputs commands to correct aircraft position thus minimizing touchdown dispersion in both the longitudinal and lateral axes. When the aircraft deviates from the flare envelopes, such as may be due to extreme atmospheric conditions such as wind gusts, a signal commanding pitch or roll correction is output to the pitch or roll control loop which subsequently acts to adjust aircraft position such that the aircraft's flarepath is returned to within the flare envelopes.