United Technologies Corporation

Sikorsky Aircraft Corporation

Phillip J Gold

Donald L Fogler, Jr

Stuart C Wright

James B Dryfoos

Joseph P Skonieczny

Don L Adams

James F Keller

Joseph A Post

James L Richard

Nicholas D Lappos

Joseph T Driscoll

Lorren Stiles

Thomas H Lawrence

Todd M Fanciullo

John Randall Mayo

Roy Brewer

Walter R Faull

Steven L Glusman

A static longitudinal stability (SLS) system provides an unobtrusive airspeed hold function that reacts to the pilot control inputs into the flight control system and the measured states of the aircraft to engage and disengage smoothly without any explicit mode selection by the pilot. The SLS engages airspeed hold when the aircraft is close to trimmed, non-accelerating state. This logic allows the pilot to either fly close to a desired airspeed and let SLS engage, or trim the aircraft in an accelerating/decelerating maneuver and just wait for the SLS to capture a speed when longitudinal acceleration is small. The SLS is not dependant on the pilot specifically selecting this mode, but rather engages and disengages in response to flight control system status and how the aircraft is being maneuvered.

Fly by wire static longitudinal stability compensator system

The flight control system includes a trim augmentation algorithm which advantageously provides two primary functions: direct augmentation of the rotor trim via on-axis stick input; and automatic cross-axis trim transfer as a function of the commanded yaw rate. In a first function of the present invention, when the aircraft is lagging the pilots commanded maneuver state the stick command is multiplied by a gain to direct the actual vehicle response toward said pilot commanded rate signal and improve the model following control laws. In a second function of the present invention, a cross-axis trim transfer function of the trim augmentation algorithm uses the commanded yaw rate to anticipate the trim change when the aircraft is rotating in winds. When a yaw maneuver is performed, the cross-axis trim transfer function automatically &ldquo;stirs&rdquo; the cyclic as the aircraft rotates about the yaw axis.

Trim augmentation system for a rotary wing aircraft

A feedback system for reducing the effects of destabilizing forces in a vehicle (e.g. helicopter) by utilizing acceleration feedback data sensed by an acceleration sensor located in the vehicle. This sensed acceleration signal is passed through a low-pass filter and combined with a command signal to produce an error signal. The error signal is multiplied by a predetermined gain to produce a corrective signal that is utilized by actuators to adjust an adjustable surface, such as a swashplate, on the vehicle.

Model-following control system using acceleration feedback

An integrated fire and flight control system of the type which controls aircraft flight dynamics to referenced values defined by a weapon launch solution to provide optimum aircraft to target orientation, further optimizes the aircraft's angle of attack (AOA) to the target by controlling the aircraft's vertical speed to modify the aircraft's rate of climb or descent as necessary to produce an actual pitch attitude that is within a range of AOA values corresponding to the range of permissive weapon launch vertical speeds recommended by the weapons manufacturer.

Integrated fire and flight control system for controlling the angle of

An integrated fire and flight control (IFFC) system senses the actual engine torque of a rotary wing aircraft in successive real time intervals and compares the sensed interval values with one or more reference maximum torque limits for each of one or more prescribed maximum time intervals and, in the presence of an above limit actual torque for an above limit prescribed time interval, the IFFC automatically reduces the collective axis command.

Integrated fire and flight control system with automatic engine torque

A control system for small unmanned rotorcraft compensates for vehicle pitch control saturation caused by the need for sudden vehicle pitch attitude correction, in turn often caused by wind gusts. The rotorcraft has a pitch-variable rotor system responsive to a vehicle pitch servo command for cyclically controlling rotor pitch and responsive to a collective servo command for collectively controlling rotor pitch. A compensating signal derived from the unlimited vehicle pitch servo command signal is cross-connected to the unlimited collective servo command signal to compensate for pitch control saturation, typically by reducing the magnitude of the resulting collective servo command signal. The compensating signal is derived by passing the unlimited vehicle pitch servo command signal through a dead band which responds as the signal approaches saturation and by preferably also then providing high and low frequency shaping to that signal.

Compensation for rotorcraft pitch axis control saturation

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.

Inertial velocity command system

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).

Turn coordination inhibit for rotary wing aircraft control system

A vertical control system for a rotary winged aircraft receives inputs from a displacement collective stick and a sidearm controller. The system provides a set point for the vertical rate of change of the helicopter as a function of a vertical lift command signal from the sidearm controller. The set point is used as a reference for an altitude rate of change feedback path, and an integrated value of the set point is used for an altitude feedback path. The set point is also input to a feedforward control path having an inverse vehicle model to provide a command signal indicative of the command for aircraft collective pitch necessary to achieve the desired set point. Signals from all three paths (i.e., the altitude rate of change feedback path, the altitude feedback path, and the feedforward path) are summed to provide a signal to backdrive the displacement collective which controls main rotor collective pitch.

Vertical control system for rotary wing aircraft

A transient free synchronizer stores an input trim signal value on command, and provides an output signal value which transitions in a smooth, continuous manner, when the synchronizer is commanded to the store the input signal. Synchronizers are typically used to capture desired aircraft attitude, position, or velocity trim points for use in aircraft autopilot and automatic flight control systems.

Transient free synchronizer for autopilot systems

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

High speed turn coordination for rotary wing aircraft

A helicopter flight control system includes an automatic turn coordination system which provides a coordinating yaw command signal to the tail rotor of the helicopter. The system stores on command a trim attitude (e.g., signals indicative of bank angle, lateral ground speed, and lateral acceleration) about which automatic turn coordination is provided, thus providing the pilot with automatic turn coordination about attitudes other than wings level.

Automatic turn coordination trim control for rotary wing aircraft

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

Low speed turn coordination for rotary wing aircraft

A helicopter fly-by-wire flight control system (21) includes a model following control system architecture which provides the pilot with maneuvering feel. Maneuvering feel is generated by varying the gain of a rate model (52) in the control system to schedule a constant sidearm controller force per g of load factor currently on the helicopter main rotor. This gain variation works in conjunction with control logic which requires the pilot to apply a nose up pitch command during steep banked turns in order to offset a control system induced downward (pitch) bias of the aircraft nose. The control system induced bias is generated only during steep banked turns, thus allowing the pilot to fly shallow banked coordinated turns without having to apply a nose up pitch command.

Maneuver feel system for a rotary wing aircraft

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).

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Shelton, CT, United States of America

Phillip J Gold