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    31.
    发明专利
    未知

    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

    IPC: B64C13/00 ,  B64C13/08 ,  B64C13/18 ,  B64C13/44 ,  B64C27/56 ,  B64C27/57 ,  G05B9/03 ,  G05D1/08

    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.

    36.
    发明专利
    未知

    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

    Inventor： WRIGHT STUART CAMMETT ,  ADAMS DON LUIS ,  FISCHER WILLIAM CHRISTIAN ,  VERZELLA DAVID JOHN

    IPC: B64C13/18 ,  B64C27/57 ,  G05D1/08 ,  G05D13/00 ,  G05B11/40

    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.

    37.
    发明专利
    未知

    公开(公告)号：DE3210817A1

    公开(公告)日：1982-10-21

    申请号：DE3210817

    申请日：1982-03-24

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： WRIGHT STUART CAMMETT ,  ADAMS DON LUIS ,  FISCHER WILLIAM CHRISTIAN ,  VERZELLA DAVID JOHN

    IPC: B64C13/18 ,  B64C27/56 ,  G05D1/00 ,  G05D1/08 ,  B64C13/16

    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.

    38.
    发明专利
    未知

    公开(公告)号：IT8220490D0

    公开(公告)日：1982-03-30

    申请号：IT2049082

    申请日：1982-03-30

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： WRIGHT STUART CAMMETT ,  ADAMS DON LUISI ,  FISCHER WILLIAM CHRISTINA ,  VERZELLA DAVID JOHN

    IPC: B64C13/18 ,  B64C13/16 ,  B64C27/56 ,  B64C27/57 ,  G05D1/08 ,  B64C

    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.

    39.
    发明专利
    未知

    公开(公告)号：IT8220489D0

    公开(公告)日：1982-03-30

    申请号：IT2048982

    申请日：1982-03-30

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： WRIGHT STUART CAMMETT ,  ADAMS DON LUIS ,  FISCHER WILLIAM CHRISTIAN ,  VERZELLA DAVID JOHN

    IPC: B64C27/57 ,  B64C13/18 ,  G05D1/08 ,  B64D

    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.

    40.
    发明专利
    未知

    公开(公告)号：NO780718L

    公开(公告)日：1978-09-11

    申请号：NO780718

    申请日：1978-03-02

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： JOHNSON RAYMOND GORDON JR ,  COTTON LOUIS SAXON ,  VERZELLA DAVID JOHN

    IPC: B64C27/06 ,  G05D1/08 ,  B64C

    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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