DON LUIS ADAMS
Pilots at Benson Rd, Fairfield, CT

4279391, Jul 21, 1981

Jan 24, 1979

6/006314

Don L. Adams - Fairfield CT
Richard D. Murphy - Monroe CT
William C. Fischer - Monroe CT

United Technologies Corporation - Hartford CT

B64C 1144

244 1713

A helicopter having an automatic flight control system including an inner, stability loop is rendered less sensitive to short-term, inadvertent pilot inputs by applying a washed-out derivative of a stick position signal to the inner stability loop in a sense to countermand the pilot action. Using a washed-out signal countermands only short-term rapid stick motions, which may be induced by the pilot actively, but inadvertently, or inactively due to coupling between the pilot or the stick and motion of the fuselage, while permitting purposeful, long-term stick positions to have the full, intended effect.

4580223, Apr 1, 1986

Nov 7, 1983

6/549387

Stuart C. Wright - Woodbridge CT
Richard D. Murphy - Trumbull CT
Don L. Adams - Fairfield CT

United Technologies Corporation - Hartford CT

G06F 1550, G06F 778

364434

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.

4385356, May 24, 1983

Mar 30, 1981

6/249302

David J. Verzella - Guilford CT
William C. Fischer - Monroe CT
Don L. Adams - Fairfield CT
Stuart C. Wright - Milford CT

United Technologies Corporation - Hartford CT

G06F 1550, G06G 778

364434

In a helicopter automatic flight control system having both automatic pitch attitude retention (70, 71) and automatic airspeed hold (84) and responsive both to pitch attitude error (206) and integrated (241) airspeed error (230), saturation of either the attitude error or integrated airspeed error circuitry is avoided by sensing (171) integrated airspeed error buildup to a threshold magnitude (which may be a significant fraction of the instantaneous authority of the automatic flight control system), and causing (169, 167) automatic slewing of the pitch attitude reference signal toward the current pitch attitude and reduction of the integrated airspeed error. In a disclosed embodiment, the correction of attitude reference and reduction of airspeed error is effected by utilizing an equal effective time constant (220, 221; 258, 259) for both actions over the same period of time (170, 168, 175).

4599698, Jul 8, 1986

May 2, 1983

6/490698

William C. Fischer - Monroe CT
Don L. Adams - Fairfield CT
Stuart C. Wright - Milford CT
David J. Verzella - Guilford CT

United Technologies Corporation - Hartford CT

G05D 100, G05D 2302, G06F 1520

364551

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.

4477876, Oct 16, 1984

Mar 30, 1981

6/248768

Stuart C. Wright - Milford CT
Don L. Adams - Fairfield CT
William C. Fischer - Monroe CT
David J. Verzella - Guilford CT

United Technologies Corporation - Hartford CT

G06F 1550, B64C 1318, G05D 108

364434

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.

4387432, Jun 7, 1983

Mar 30, 1981

6/249300

William C. Fischer - Monroe CT
Don L. Adams - Fairfield CT
David J. Verzella - Guilford CT
Stuart C. Wright - Milford CT

United Technologies Corporation - Hartford CT

G06F 1550, G06G 778

364434

In an aircraft automatic flight control system, proportional commands (54, 55) provided to fast, limited authority inner loop actuators (12, 13) are integrated (41), and when the integrator output indicates that the inner loop actuators 12, 13 have been driven a certain percentage of their authority, a comparator (130, 132) activates a pulse generator (137, 138) to provide timed excitation of an actuator (150), thereby to position the aircraft control system outer loop by a commensurate increment. Driving the actuator for a longer time than the desired pulse width is detected (165-169) and causes automatic shutdown (190) of the actuator. Resetting the integrator at the start of each pulse (162, 104), and pulse-controlled gating of the pulse circuits (172, 135, 136) allow sensing of authority transitions which occur within a pulse, and permit a subsequent pulse in response thereto.

4330829, May 18, 1982

Jul 7, 1980

6/166010

William C. Fischer - Monroe CT
Don L. Adams - Fairfield CT

United Technologies Corporation - Hartford CT

G06G 770, B64C 1134

364434

Oscillations in helicopter attitude sustained by the aerodynamic response of the helicopter to an automatic flight control system which is responsive to an attitude sensor, are eliminated by band reject (notch) filtering of a control system stability command to the aircraft, derived from rate of changes of such attitude at a frequency related to the aircraft attitude oscillations induced by the rate-controlled stability compensation.

4696445, Sep 29, 1987

Sep 18, 1986

6/909046

Stuart C. Wright - Woodbridge CT
Lorren Stiles - Woodbury CT
Don L. Adams - Fairfield CT

United Technologies Corporation - Hartford CT

B64C 1312

244229

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.

4392203, Jul 5, 1983

Mar 30, 1981

6/249273

William C. Fischer - Monroe CT
Don L. Adams - Fairfield CT
David J. Verzella - Guilford CT
Stuart C. Wright - Milford CT

United Technologies Corporation - Hartford CT

G06F 1550, G06G 778

364434

An aircraft automatic flight control system having a roll channel (FIG. 1), a yaw outer loop trim system (bottom of FIG. 2) and a yaw stability inner loop (top of FIG. 2) provides proportional (129) and integral (114) lateral acceleration inputs to the yaw trim system (106, 93) during coordinated turns, and provides proportional (130) and lagged (131) roll rate inputs to the trim system to provide initial coordination to the turns.

4003532, Jan 18, 1977

Mar 15, 1976

5/666590

Don L. Adams - Fairfield CT
Raymond G. Johnson - Milford CT

United Technologies Corporation - Hartford CT

B64C 1134, G05D 108

244 1713

The automatic heading retention mode of operation of an automatic flight control system for a rotary wing aircraft is discontinued and reestablished through the use of a logic circuitry responsive to the yaw rate, bank angle and airspeed of the aircraft and to the force applied to the cyclic pitch control member by the pilot. The logic circuitry provides for the isolation of information commensurate with left and right turns and summation of the yaw rate and bank angle in each direction to determine if automatic disengagement of the heading hold is warranted.