United Technologies Corporation

The Boeing Company

Sikorsky Aircraft Corporation

Stuart C Wright

Don L Adams

David J Verzella

William C Fischer, Jr

Phillip J Gold

James B Dryfoos

Nicholas D Lappos

Richard D Murphy

James J Howlett

Charles W Evans

Joseph P Skonieczny

Ari M Novis

Karim Khial

Paul T Weisser, Jr

Arthur Lawrence Sivigny

Byron Graham, Jr

Troy J LaMontagne

Porter D Sherman

Lorren Stiles, Jr

David H Hetzler

A system for sensing a force applied to an aircraft includes a first sensor, a second sensor, and a processor operative to define a first velocity vector as a function of a first velocity due to a rotation motion of the aircraft, define a second velocity vector as a function of a second velocity due to the rotation motion of the aircraft, define an instant axis of rotation of the aircraft as a function of the first velocity vector and the second velocity vector, determine whether a force has been exerted on a first portion of the aircraft, and output an indication that a force has been exerted on the first portion of the aircraft responsive to determining that the force has been exerted on the first portion of the aircraft.

Method and system for detecting forces on aircraft

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.

Helicopter turn coordination and heading hold mode control

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

Adaptive control system input limiting

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 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 includes a model following control system architecture to control helicopter response about the yaw axis. The control system processes information from a variety of helicopter sensors in order to provide an output command signal to the tail rotor. Yaw control is accomplished by applying a force on a sidearm controller which provides a yaw axis command signal that is used to schedule a desired yaw rate of change. The lateral acceleration and airspeed signals are used to compute a synthesized yaw rate signal which is compared against the desired yaw rate of change, thus allowing the pilot to control lateral acceleration when he applies force on the sidearm controller.

High speed yaw control system for rotary wing aircraft

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.

Attitude synchronization for model following control systems

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.

Model following control system

A method of and an arrangement for controlling the operation of a reversible D.C. electric motor including a winding having two ends each of which is connected to both the positive and negative terminals of an electric current source through respective control switches that can be selectively closed and opened in response to command signals to apply voltage differentials across the winding with attendant electric current flow through the winding in one or the opposite direction for operating the electric motor in one or the other of its senses of rotation with intervening rotation sense reversals utilize the original command signals for determining when the sense of rotation reversal would result in an excessive electric current flow through the winding. In such a case, the original command signals are temporarily replaced by substitute command signals of such a character that both of the winding ends are connected through the respective control switches to only one of the electric current source terminals with attendant dynamic braking of the electric motor.

Method of and arrangement for controlling and h-bridge electric motor

Pilot workload is reduced in a helicopter having a force-type, multi-axis sidearm control stick by providing a displacement-type control stick for collective blade pitch control. Either stick may be used by the pilot. When the force-type stick is employed, a trim system causes the displacement-stick to track a collective position command signal which is provided to the blade actuators. Changeover of control to the displacement-type control stick is accomplished either with a switch, or by moving the collective-type control stick. The signals associated with each control stick are alternately faded in and out to assure a smooth transition when collective control is switched over from one to the other.

Collective control system for a helicopter

A heading trim reference system provides beep commands indicative of a desired heading/attitude change to the roll channel of an AFCS as roll commands at slow speed in a conventional manner, and to the yaw channel as yaw commands at cruise speed. At cruise speed, the yaw synchronizing integrator responds to the beep commands and provides an error signal to existing turn coordination logic to turn the aircraft.

Heading reference trim system

The directions of travel of an inner and an outer loop actuator are monitored for movement in opposite directions within selected ranges and the outer loop actuator is disabled under selected conditions. These conditions may include movement of both the inner and outer loop actuators in opposite directions, any one of which has been detected moving at a rate greater than a selected rate or to a position outside a selected range. The invention is particularly suited for an aircraft trim actuator shutdown monitor system.

Aircraft trim actuator shutdown monitor system and method

In a dual actuator system, runaways are identified by comparing the position and rate of one actuator to another. When there is a threshold position discrepancy and a sustained rate discrepancy, a fault is indicated. The faster actuator is identified as the runaway.

Dual-actuator monitor

The sense (direction) of outer-loop commands is compared with the sense (direction) of inner-loop signals in an aircraft AFCS. A disparity in sense indicates an outer-loop failure. To reduce nuisance failures the comparison is made only when inner-loop signal thresholds are exceeded and, in the context of outer-loop commands which are pulses, only at the onset of an outer-loop command and for a limited time thereafter.

Outer-loop monitor for aircraft automatic flight control system

An aircraft automatic flight control system (AFCS) includes a pair of fast, limited authority inner loop actuators responsive to signals indicative of aircraft attitude or other flight parameters such as airspeed, the inner loop being recentered by an outer loop actuator responsive to attitude or other aircraft parameter-indicating signals (54,55). Commands applied to the outer loop are applied in a lagged fashion in opposite direction so as to drive the inner loop actuators back toward the center of their authority. The rate of response of the outer loop is adaptive in response to magnitude of inner loop input (101, FIG. 2). A pitch bias command is provided to the inner loop as a function of airspeed multiplied inversely with collective pitch, and as a function of the rate of change of collective stick position, so as to provide a positive static pitch trim gradient and decouple collective pitch from the longitudinal cyclic pitch channel. When the inner loop runs out of authority, the outer loop may assume a high gain mode to keep pace with the demand of the pitch bias function. When the AFCS is operating single-on, the outer loop response is normally delayed for three seconds to allow a pilot time to recover from a hardover. However, when a collective maneuver is being made, the outer loop response is not delayed so as not to interfere with the pitch bias function, especially with regard to collective decoupling.

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Milford, CT, United States of America

Stuart C Wright