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    12.
    发明专利
    未知

    公开(公告)号：IT1151380B

    公开(公告)日：1986-12-17

    申请号：IT2048882

    申请日：1982-03-30

    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 ,  G01C

    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.

    15.
    发明专利
    未知

    公开(公告)号：IT8220491D0

    公开(公告)日：1982-03-30

    申请号：IT2049182

    申请日：1982-03-30

    Applicant: UNITED TECHNOLOGIES CORP

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

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

    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.

    16.
    发明专利
    未知

    公开(公告)号：NO145462B

    公开(公告)日：1981-12-21

    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

    IPC: B64C27/06 ,  G05D1/08 ,  B64C27/54

    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.

    20.
    发明专利
    未知

    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

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

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

    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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