United Technologies Corporation

Sikorsky Aircraft Corporation

USA As Represented by Secretary of the Navy

Don L Adams

Stuart C Wright

William C Fischer, Jr

David J Verzella

Ari M Novis

Phillip J Gold

Richard D Murphy

Albert Sivahop

Troy J LaMontagne

Bryan S Cotton

Nicholas D Lappos

Charles W Evans

Franklin A Tefft

Raymond G Johnson, Jr

Lou S Cotton

Gregory P Wright

Byron Graham, Jr

Arthur Lawrence Sivigny

Thomas J Provenzano

Lorren Stiles, Jr

Ann Crowley

An aircraft flight mode selector assembly utilizes computer-generated menus and soft wired bezel keys to enable the use of a greater number of flight modes to be accessed in a smaller space than is presently possible with a flight mode selector consisting solely of conventional dedicated switches, switch indicator lights, and hard wired selector panel assemblies. Serial bus connections between the flight mode selector assembly and the flight control computer in the aircraft reduce the amount of wiring needed for the flight mode selector assembly, as compared to present protocol which utilizes dedicated wiring for each dedicated switch in the aircraft flight mode selector. The use of computer-generated displays and menus permits easy expansion of the flight mode selector assembly without the need to add additional switches and switch indicator lights. The inclusion of a display information look-up table in the processor controller, which table is separate from the software program in the controller, allows changes to the flight mode selector panel display to be made merely by changing the look-up table. The flight mode selector assembly also includes a pilot message display which is used to convey text messages to the pilot of the aircraft.

Aircraft flight mode selector system

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.

Vertical performance limit compensator

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 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

An arrangement for controlling the position of a movable member, especially an aircraft control surface member, includes an operational amplifier and a capacitor connected in parallel to the amplifier. The control arrangement further includes a detector which detects the restoration of the supply of electric power following its interruption and generates a limited-duration pulse. This pulse is then used to close a switch, which is interposed between an input terminal of the capacitor and the ground, for the duration of this pulse, so that the capacitor is charged much more rapidly than otherwise. The output signal of the combination of the operational amplifier and the capacitor is supplied to an actuator for the movable member so that the extent of excursion of the movable member out of its desired position after the electric power supply has been restored is minimized.

Arrangement for controlling the position of an aircraft control surface

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

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.

Incorporation of pitch bias actuator function into an existing AFCS

An improved flight director go-around mode adjusts the collective stick position in a closed loop to increase lift to a desired value. If the engagement airspeed is below the maximum rate of climb airspeed, the airspeed is increased towards the maximum rate of climb airspeed at a moderate rate. When the maximum rate of climb airspeed is attained, or if the engagement airspeed is at least the maximum rate of climb airspeed, and the rate of climb is satisfactory, the airspeed is maintained. If the engagement airspeed is greater than the maximum rate of climb airspeed and the desired rate of climb is not achieved within ten seconds, the airspeed is decreased towards the maximum rate of climb airspeed. Heading or course is automatically maintained when go-around is engaged.

Flight director go-around mode

An improved PBA shutdown monitor circuit senses collective pitch control position rate of change magnitude and inhibits PBA shutdown in response to the continuous presence of sensed magnitude within a range of selected collective pitch rate magnitudes bounded by a slow rate limit signal value and a fast rate limit signal value.

Shutdown monitor for a helicopter pitch bias actuator

In an aircraft automatic flight control system, automatic shutdown (42) of the roll axis thereof (FIG. 1) as a consequence of faults detected therein also provides automatic shutdown of the yaw axis thereof (180) in the same fashion as detection of faults in the yaw axis (171, 167) provides shutdown of the yaw axis. Coupling of the roll axis shutdown into the yaw axis shutdown allows determining which of three axes cause lighting of a single trim shutdown indicator by resetting the yaw shutdown register (145) and simultaneously resetting both the pitch and roll shutdown registers by means of a single cyclic pitch control switch. Automatic yaw trim outer loop and yaw inner loop stability systems are disclosed (FIG. 2).

Aircraft roll-yaw fault protocols

In an aircraft automatic flight control system having a reference parameter synchronizing system (70) operable in response to a trim release switch (44), an initial trim release period (139), on the order of a large fraction of a second (137) causes (217, 218) a relatively slow effect trim reference integrator (208, 211) time constant, for smooth transitions of any error signal, followed by a relatively fast (216) effective reference integrator time constant for close, rapid tracking of the reference signal with the actual aircraft parameter.

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Fairfield, CT, United States of America

Don L Adams