United Technologies Corporation

Douglas H Clelford

Richard D Murphy

Donald W Fowler

Joseph P Skonieczny

Franklin A Tefft

Ricardo L Perez

Raymond J Brand

Collective stick inputs are provided (18) to a high pass filter (24) when collective stick reversals in excess of a threshold collective authority rate (15%/sec) and in excess of a threshold frequency (3 stick reversals per second) are sensed. The output of the high pass filter (24) is scaled (26) and subtracted (14) from the collective inputs so as to effectively cancel the threshold stick reversals. Otherwise, if no stick reversals in excess of the threshold collective authority rate are sensed for a threshold time interval (5 seconds), the collective stick inputs are not provided to the high pass filter.

Preventing augmenting vertical load induced oscillations in a helicopter

An oscillatory-failure monitor (101) compares a parameter of a plurality of signals (A, B) within a tolerance (TOL) to determine agreement or disparity among the signals. Each discrete miscompare occurrence (i.e., singular disparity following full agreement) is counted by incrementing (27; 54) a counter (OSCCT; CNTR). The counter is decremented (32; 62) whenever the signals compare for a predetermined time interval (29; 47). An oscillatory-failure is declared when the counter increments to a threshold (28, 21; 57, 58, 60). Both digital (FIG. 1) and dedicated hardware (FIG. 4) embodiments are disclosed.

Oscillatory failure monitor

A torque limiting altitude hold system for a helicopter engages torque limiting (56, 203) when excessive torque is anticipated (138, 202) as determined by the summation of present torque and torque rate times a reference value (126, 194) exceeds maximum torque with torque limiting engaged, altitude commands are faded out (42, 189) and torque commands are faded in (44, 190) and the collective command integrator is switched from altitude to torque (48, 54; 181, 185), torque limiting is ended in response to negative altitude commands or anticipated desired altitude signal (96, 150, 152; 205, 206); the anticipated desired altitude is determined by subtracting from the altitude error a time function of the altitude rate (84, 90; 193), torque limiting is not allowed to reengage for two seconds after disengaging (144, 204) nor within three seconds after reaching desired altitude during an automatic descent (146, 207), the system provides smooth transitions from altitude control to torque control, without oscillation or abrupt changes; control descents are arrested without torque limiting the engine.

Automatic limiting of helicopter engine torque

A maneuver force gradient system causes a helicopter, that otherwise tends to pitch up in a banked turn, to pitch nose-down. A command signal (25) is provided as a function of the roll angle to operate the longitudinal trim actuator (27) which automatically moves the cyclic control (50) resiliently (28) forward to push the nose down. The pilot must consequently pull back on the cyclic control (50) to achieve a desired pitch attitude, thereby establishing a longitudinal positive maneuver-force gradient. The system is operable only when the pilot initiates a roll (33, 38, 50, 51) and the roll angle equals or exceeds 30.degree. (30, 33, 36, 37). Both analog (FIG. 1) and digital (FIG. 2) embodiments are disclosed, and the invention may be practiced in association with an AFCS (101) having the longitudinal trim actuator (27) and resilient linkage (28).

Maneuver-force gradient system

An automatic flight control system, for an aircraft having a roll attitude retention outer loop actuator (29), a roll stability inner loop actuator (25) and a control stick (26) for positioning control surfaces of the aircraft to control its roll attitude, includes means (54, 55) to provide a roll error signal (56) indicative of the deviation in roll attitude from a desired roll attitude. The roll attitude retention outer loop actuator (29) is controlled by a proportional (61) and integral (62) function of the roll error when force is not applied to the stick, but only as a proportional function when force is applied to the stick (48, 49, 51, 62). The roll stability inner loop actuator is controlled by a washed out (72) proportional (64) function of the roll error signal to provide short-term roll retention at roll attitudes established by the control stick during turns against trim. Sensing force on the stick (51, 95, FIG. 1) and roll attitude of a predetermined threshold magnitude (107) provides adjustments in relative gain (113, 117, 70, 71, FIG. 1; 154, 157, 160-163, FIG. 4) during turns against trim.

Aircraft short-term roll attitude retention system

For each of two computer systems, logic flowcharts describe background program in which highly detailed memory checksum tests of fixed memory and complementary tests of variable memory are performed, the background program being interrupted for utility programs which are for the most part responsive to transducer or other sensor and discrete inputs to calculate control values for operation of control actuators or other responsive devices. The utility programs include specific self test routines. A direct memory access unit is included in each computer for moving data between inputs of either computer and memories of both, and between the memories of both computers. Periodic testing of fault codes registering the health of each computer is done during utility program routines, any variation from normal causing further health-analysis routines to be performed until dispositive action-causing conditions are determined. Neither computer checks the internal health of other, but inputs, results and data link transmissions must compare equally between the two computers, or routines determine whether one computer will recognize itself (or a component thereof) as being faulty, and disable itself. If not, then each computer disables itself after disabling the other. A variety of self tests and other checks and routines are included. Disablement is accomplished in a complex fashion of each computer's output, by itself, and additional disablement if instituted by the other computer.

Complementary commands in fail-operational, fail-safe multi-computer

A helicopter pitch axis autopilot system, including airspeed hold at cruise speeds and pitch attitude hold below cruise speeds, includes integrated longitudinal acceleration drift corrected by heavily filtered Pitot static airspeed as a filtered airspeed reference, use of a beeper to nudge the airspeed reference above cruise speeds or the pitch attitude reference below cruise speeds, resynchronizing of airspeed reference, pitch attitude reference, pitch autopilot integrator and stick trim position reference, momentarily in response to airspeed transitions between cruise and sub-cruise speeds or in response to initiation of beeping, and continuously during trim release. Pilot override without disengagement of the autopilot and ultimate return to a previous speed or attitude is achieved by sensing pilot activity and causing the pitch autopilot integrator value to be held during pilot override, and limiting the value of airspeed error that drives the pitch autopilot integrator once the pilot activity has subsided.

Helicopter attitude and airspeed acquisition and retention system

A bias voltage, which can be added to force commands applied to a helicopter control stick actuator, to compensate for static null offset errors, is generated as the average of the summation or integral of actuator pressure, or pressure differential, sensed while the force is moved from a first position which is halfway aft through null to a second position which is halfway forward and back through null to the first position, the pressure readings which are averaged being taken only through the central portion of the motion, the motion being controlled to be extremely slow compared to normal permissible stick motion.

Helicopter force feel actuator automatic static null compensation

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.

Adaptive aircraft actuator fault detection

Synchronized, fail-operational, fail-safe multi-computer control system

Faults in a sensor (10) are detected by excess rate fault determination by taking the differentiated (58, 107), rate limited (60, 100, 111) integral (64, 112) from the raw sensor output (72, 113) and indicating a fault (15, 117) in the event that the difference exceeds a predetermined magnitude (76, 114). A null fault (22, 91) is provided in the event that the sensor output does not show a significant change (52, 88) within a given time interval (20, 80) whenever a related sensor (31) indicates (29, 84) that the first sensor (10) should have measurable activity. Both analog (FIG. 1) and digital (FIG. 2, FIG. 3) embodiments are disclosed.

Sensor fault detection by activity monitoring

A system which provides feel-force to the control stick of an aircraft by means of hydraulic pressure is provided with non-nulling, proportional, direct feedback loop, and a limited integral feedback loop in the drive of the pressure control servo valve that commands the pressure-generating hydraulic force actuator. Static nulls are compensated in the integral path; integration is corrected when the static null and integral output exceed a limiting value. Embodiments include software and/or hardware portions.

Force feel actuator with limited proportional/integral error feedback

In a system which provides feel-force to the control stick of an aircraft by means of hydraulic pressure, a fault indicating system which responds to excessive differences between desired and actual force in the actuator, only of a polarity in the same direction as concurrent stick motion, to separate pilot induced pressure excesses from force-system faults, is provided with directional sensitivity in the fault-indicating output logic. Excessive pressure errors in either direction of stick travel institute the monitoring of changes in stick position. Changes in stick position of an excessive magnitude are recognized as fault only if they occur in the same direction as the direction of excessive pressure error. A simple analog embodiment utilizes threshold detectors and a track store device; a dedicated-digital embodiment utilizes comparators, registers and a subtractor; and a computer embodiment utilizes a software routine.

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Trumbull, CT, United States of America

Douglas H Clelford