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公开(公告)号：GB2140174A
公开(公告)日：1984-11-21
申请号：GB8410987
申请日：1984-04-30
Applicant: UNITED TECHNOLOGIES CORP
Inventor： FISCHER WILLIAM C , WRIGHT STUART C , VERZELLA DAVID JOHN , ADAMS DON LUIS
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：GB2140172A
公开(公告)日：1984-11-21
申请号：GB8410985
申请日：1984-04-30
Applicant: UNITED TECHNOLOGIES CORP
Inventor： FISCHER WILLIAM C , WRIGHT STUART C , VERZELLA DAVID JOHN
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公开(公告)号：BR8201663A
公开(公告)日：1983-02-16
申请号：BR8201663
申请日：1982-03-24
Applicant: UNITED TECHNOLOGIES CORP
Abstract: 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.
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公开(公告)号：BR8201662A
公开(公告)日：1983-02-16
申请号：BR8201662
申请日：1982-03-24
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：BR8201660A
公开(公告)日：1983-02-16
申请号：BR8201660
申请日：1982-03-24
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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Patent Agency Ranking
公开(公告)号：GB2140173B
公开(公告)日：1986-07-30
申请号：GB8410986
申请日：1984-04-30
Applicant: UNITED TECHNOLOGIES CORP
Inventor： FISCHER WILLIAM C , 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：DE3416241A1
公开(公告)日：1984-11-08
申请号：DE3416241
申请日：1984-05-02
Applicant: UNITED TECHNOLOGIES CORP
Inventor： FISCHER WILLIAM C , VERZELLA DAVID JOHN , WRIGHT STUART C
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：DE3210868A1
公开(公告)日：1982-10-21
申请号：DE3210868
申请日：1982-03-24
Applicant: UNITED TECHNOLOGIES CORP
Abstract: 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.
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公开(公告)号：DE3210867A1
公开(公告)日：1982-10-14
申请号：DE3210867
申请日：1982-03-24
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：NO145462C
公开(公告)日：1982-03-31
申请号：NO780718
申请日：1978-03-02
Applicant: UNITED TECHNOLOGIES CORP
Inventor： JOHNSON RAYMOND GORDON JR , COTTON LUOIS SAXON , VERZELLA DAVID JOHN
Abstract: A helicopter stabilator, combining the functions of tail stabilizer and aircraft elevator, has its angle of incidence with respect to the helicopter controlled by means of a pair of reversible actuators acting in series, controlled in a closed-loop fashion by electronic hardware, the controls being disconnected when the actuators fail to track within a threshold disparity of either position or rate of change of position of each other. A test switch introduces an imbalance to test the fault circuitry. The stabilator is biased to assume a maximum incidence position at low speeds, including hover; and inputs from airspeed and collective pitch position cause it to assume a substantially level position at higher, cruise speeds. A pitch rate gyro input controls the stabilator for stable flight against pitch-inducing flight commands and external effects, such as gusts. In one embodiment a canted tail rotor which provides tail lift; to overcome tail-up and tail-down effects of the downward component of the tail rotor due to more or less thrust and real or apparent lateral accelerations, a lateral accelerometer is used. To avoid main rotor downwash against the stabilator at hover, the lateral accelerometer and collective pitch inputs are washed-out at low speeds. The bias allows final adjustment for high speed flight.
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