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公开(公告)号：BR8402038A
公开(公告)日：1984-12-11
申请号：BR8402038
申请日：1984-05-02
Applicant: UNITED TECHNOLOGIES CORP
Inventor： FISCHER WILLIAM CHRISTIAN , WRIGHT STUART C , VERZELLA DAVID JOHN
Abstract: 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.
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公开(公告)号：AU537146B2
公开(公告)日：1984-06-07
申请号：AU6752581
申请日：1981-02-20
Applicant: UNITED TECHNOLOGIES CORP
Inventor： FISCHER WILLIAM CHRISTIAN , SIVAHOP ALBERT
IPC: B64C13/00 , B64C27/57 , G05B9/03 , G05D1/00 , G05D1/08 , G06F11/14 , G06F11/18 , B64C13/16 , B64C17/06 , G06F15/20
Abstract: The main roll and yaw axes of a helicopter having a main rotor operative in response to cyclic and collective pitch commands and a tail rotor operative in response to pitch commands, include stability inputs to actuators which are selectively provided by an analog channel, a digital channel, or one half by each, the digital channels being manifested in a digital computer system which includes programs of self-test, in-flight comparison between the inputs and outputs of corresponding analog and digital channels, and ground tests of testable rate gyro inputs to the various digital and analog channels and the outputs thereof.
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公开(公告)号：DE3210818A1
公开(公告)日：1982-10-21
申请号：DE3210818
申请日：1982-03-24
Applicant: UNITED TECHNOLOGIES CORP
Abstract: In an automatic flight control system (FIG. 1) an airspeed control engage function (84) is automatically engaged (136, FIG. 2; 244, FIG. 4) in response to airspeed above a threshold magnitude, such as 45 knots (88, FIG. 2) and will remain engaged (subject to a fault condition, 135) until the airspeed command (75) reaches a predetermined, insignificant magnitude (132). An airspeed error integrator 241 which accommodates the difference between a reference attitude and an attitude required for a reference airspeed, does not react to large airspeed errors as a consequence of pilot maneuvering due to pilot force on the control stick (35, 109) opening the input (252, FIG. 4) to the airspeed error integrator.
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公开(公告)号：IT8220488D0
公开(公告)日：1982-03-30
申请号：IT2048882
申请日：1982-03-30
Applicant: UNITED TECHNOLOGIES CORP
Abstract: 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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公开(公告)号：IT1210866B
公开(公告)日：1989-09-29
申请号：IT2048982
申请日：1982-03-30
Applicant: UNITED TECHNOLOGIES CORP
Abstract: In an automatic flight control system (FIG. 1) an airspeed control engage function (84) is automatically engaged (136, FIG. 2; 244, FIG. 4) in response to airspeed above a threshold magnitude, such as 45 knots (88, FIG. 2) and will remain engaged (subject to a fault condition, 135) until the airspeed command (75) reaches a predetermined, insignificant magnitude (132). An airspeed error integrator 241 which accommodates the difference between a reference attitude and an attitude required for a reference airspeed, does not react to large airspeed errors as a consequence of pilot maneuvering due to pilot force on the control stick (35, 109) opening the input (252, FIG. 4) to the airspeed error integrator.
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Patent Agency Ranking
公开(公告)号：IT1151537B
公开(公告)日：1986-12-24
申请号：IT2049182
申请日：1982-03-30
Applicant: UNITED TECHNOLOGIES CORP
Abstract: An aircraft automatic flight control system includes a pair of fast, limited authority inner loop actuators (12, 13) responsive to signals (52-55) indicative of aircraft attitude (68, 69) or other flight parameters such as airspeed (84), the inner loop being recentered by an outer loop actuator (37) responsive to attitude or other aircraft parameter-indicating signals (54, 55). Commands (40) applied to the outer loop are applied in a lagged fashion (58, 59) in opposite direction so as to drive the inner loop actuators (12, 13) back toward the center of their authority. The rate of response of the outer loop (FIG. 2, FIG. 5) is adaptive in dependence upon airspeed (93, 96, 212, 213) and in response to magnitude of inner loop input (101, FIG. 2). An integral gain (41), pulsed (39), open loop drive of the outer loop actuator (37) and outer loop automatic shutdown (38) are disclosed.
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公开(公告)号：BR8402037A
公开(公告)日：1984-12-11
申请号：BR8402037
申请日：1984-05-02
Applicant: UNITED TECHNOLOGIES CORP
Inventor： FISCHER WILLIAM CHRISTIAN , ADAMS DON LUIS , VERZELLA DAVID JOHN
Abstract: 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.
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公开(公告)号：DE3416243A1
公开(公告)日：1984-11-08
申请号：DE3416243
申请日：1984-05-02
Applicant: UNITED TECHNOLOGIES CORP
IPC: B64C13/00 , B64C13/08 , B64C13/18 , B64C13/44 , B64C27/56 , B64C27/57 , G05B9/03 , G05D1/08 , G05D1/00 , G05B23/00
Abstract: 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.
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公开(公告)号：AU529394B2
公开(公告)日：1983-06-02
申请号：AU5484880
申请日：1980-01-23
Applicant: UNITED TECHNOLOGIES CORP
Inventor： ADAMS DON LUIS , MURPHY RICHARD DENNIS , FISCHER WILLIAM CHRISTIAN
Abstract: 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.
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公开(公告)号：BR8201661A
公开(公告)日：1983-02-16
申请号：BR8201661
申请日：1982-03-24
Applicant: UNITED TECHNOLOGIES CORP
Abstract: In an automatic flight control system (FIG. 1) an airspeed control engage function (84) is automatically engaged (136, FIG. 2; 244, FIG. 4) in response to airspeed above a threshold magnitude, such as 45 knots (88, FIG. 2) and will remain engaged (subject to a fault condition, 135) until the airspeed command (75) reaches a predetermined, insignificant magnitude (132). An airspeed error integrator 241 which accommodates the difference between a reference attitude and an attitude required for a reference airspeed, does not react to large airspeed errors as a consequence of pilot maneuvering due to pilot force on the control stick (35, 109) opening the input (252, FIG. 4) to the airspeed error integrator.
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