• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    This invention relates to aeronautical engineering. The purpose of the invention is to increase the comfort of piloting and the life of the aircraft by reducing the vibrations of the buffering type wing. On the wing 1, a rudder surface 3 is installed in the root part, as well as a sensor 5, which is connected to the servo valves of the two-chamber hydraulic drive installed between the steering surface drive 4 and the steering surface itself through the threshold filter 6 and the filtered signal processing unit 7. The signal from the sensor after processing is applied to a two-chamber drive, which causes alternating deflection of the steering surface for damping oscillations / buffering / wing. 1 hp f-ly, 2 ill.
    公开号:SU1512494A3
    申请号:SU833633575
    申请日:1983-08-10
    公开日:1989-09-30
    发明作者:Морис Дестуиндер Роже；Буттес Жак；Пуассон-Кентон Филипп
    申请人:Оффис Насьональ Д, Этюд Э Де Решерш Аэроспасьаль О.Н.Э.Р.А. (Фирма)；
    IPC主号:
    专利说明:

    1
    The invention relates to aeronautical engineering, in particular, to devices for reducing vibration (bafting) of an aircraft wing.
    The purpose of the invention is to improve the comfort of piloting and the life of the aircraft by reducing the vibrations of the krusha type of buffering.
    wing. The output electrical signal from the sensor 5 passes through a threshold circuit and a band-pass filter, the bandwidth of which corresponds to the shape of the bending wing vibration to be corrected. Behind a bandpass filter, the signal passes through a pseudo-integrating circuit.
    Fig, 1 presents the principle of 10 mu, designed to produce
    on the scheme of the proposed device; figure 2 - diagram of the drive with the use of the proposed device.
    In the root part of the wing 1 of the aircraft 2, a rudder surface 3 is installed, for example, an interceptor or flap is kinematically connected with the actuator 4. A sensor 5 is installed on the wing in its root part, which outputs a signal; characterizing in amplitude, frequency and phase the instantaneous value of the parameter associated with the buffering, for example, the magnitude of the dynamic voltage at the sensor installation site. The sensor is electrically connected to the threshold filter 6, which outputs the output signal when the amplitude of the signal outputted by the sensor exceeds the upper level of the normal wing vibration threshold. The output of the threshold filter 6 is connected to the input of the filtered signal processing unit 7, which includes a series-arranged bandpass filter whose bandwidth corresponds to the shape of the wing oscillations to be corrected, a pseudo-integrating circuit that generates a signal reflecting the speed of oscillatory motion, and a phase control circuit to reduce wing buffering. The output of unit 7 is electrically connected to servovalves 8 and 9 of two-chamber hydraulic actuator 10, which is installed in a kinematic chain between the actuator 4 of the steering surface 4 and the steering surface 3. The stroke of this actuator 10 is for safety purposes if it fails Conditions for additional deviation of the rudder at clearing the buffering. Amplitude of oscillations necessary to reduce the buffering, on the order of ± 1
    The sensor 5 measures the parameter reflecting the dynamic stress due to the deformation of the bearing surface caused by vibration.
    five
    0
    five
    five
    0
    five
    0
    five
    signal reflecting the speed of oscillatory motion. The signal then passes through a phase adjustment circuit, where a signal is generated that reflects the exact phase and amplitude, taking into account the intrinsic characteristics of the steering surface used to reduce the buffering. This signal enters the servo valves 8 and 9 of the two-chamber hydraulic actuator 10, causing alternating deflections of the steering surface, and therefore the appearance on the wing 1 of alternating aerodynamic loads with amplitude and phase to reduce the apparent shape of the wing vibration.
    Invention Formula
    1. A device for reducing aircraft wing vibrations, including a steering surface located at the root of a wing, mounted on the wing sensors for measuring the parameter associated with wing vibration, as well as a threshold filter that distinguishes the characteristics of one of the wing oscillations and the unit processing of the filtered signal, which is different from the fact that, in order to increase the comfort of piloting and the aircraft resource by reducing the vibrations of the buffalo wing, between the actuator and the steering surface Strongly an optional two-: hkamer ny and a hydraulic drive with servo-valve, which provides an oscillating movement of a steering surface
    „- | About
    the maximum amplitude is not more than 1, while the additional drive is electrically connected with the processing unit of the filtered signal, which characterizes the dynamic stresses caused by oscillations such as buffing,
    2. Device POPP.1 and 2, characterized in that the processing unit of the filtered signal is 15124946
    It is not in the form of a sequentially allocating circuit and a circuit of an adjustable band-pass filter, pseudo-inn phases.
    FIG. 2
    权利要求:
    Claims (2)
    [1]
    Claim
    1. A device to reduce the vibration of the wing of an aircraft, including a steering surface with a drive located in the root of the wing, sensors mounted on the wing for measuring a parameter associated with the vibration of the wing, and a threshold filter that distinguishes the characteristics of one of the waveforms of the wing, and a filtered processing unit a signal, which is distinguished by the fact that, in order to increase the comfort of piloting and the life of the aircraft by reducing the vibration of the wing such as buffing, between the drive and the steering surface An additional two-chamber and hydraulic actuator with servo valves has been installed, providing oscillatory movement of the steering surface with a maximum amplitude of not more than 1 °, while the additional actuator is electrically connected to the filter signal processing unit, which characterizes dynamic stresses caused by fluctuations like buffering.
    [2]
    2. The device according to PP.1 and 2, characterized in that the processing unit of the filtered signal is executed
    1512494 6 nn in the form of a sequentially located tagging circuit and a circuit of an adjusted band-pass filter, pseudo-phase inversion.
    11 3 FIG. 2
    类似技术:
    公开号 | 公开日 | 专利标题
    SU1512494A3|1989-09-30|Arrangement for fighting vibration of aircraft wing
    CN103983806B|2016-08-24|A kind of fiber grating high frequency acceleration transducer based on flexible hinge
    US6772979B2|2004-08-10|Method and device for reducing the vibratory motions of the fuselage of an aircraft
    US1853069A|1932-04-12|Stabilizing apparatus
    AU2004230232A1|2005-06-02|A method of tuning a vibration absorber
    US5895012A|1999-04-20|Method and device for reducing the effect of the vibration generated by the driveline of a helicopter
    JP6659492B2|2020-03-04|Engine test equipment
    US5816533A|1998-10-06|Method and device for reducing the vibration generated on the structure of a helicopter
    EP2653382B1|2014-06-25|Self tuning vibration absorber
    GB2149473A|1985-06-12|Actuator system
    US10532808B2|2020-01-14|Apparatus for using aircraft active vibration control system as pilot cueing aid
    US2411373A|1946-11-19|Torsional vibration absorber
    KR20170043961A|2017-04-24|The manufacturing method of a dynamic vibration absorber
    SU961007A1|1982-09-23|Vibration drive
    Carnegie et al.1974|Effect of forcing terms and general characteristics of torsional vibrations of marine engine systems with variable inertia
    SU1265414A1|1986-10-23|Vibration isolator
    RU2337309C1|2008-10-27|Relay-supply pneumatic actuator equipped with missile control system vibration linearisation
    SU1738378A1|1992-06-07|Device for excitation of vibrations
    RU906124C|1994-03-30|Device for dumping oscillations in helicopter mechanical transmission
    SU1509615A1|1989-09-23|Meter of vibration parameters
    JP2003137191A|2003-05-14|Lead lag damping device for helicopter and lead lag damping method for helicopter
    RU1839872C|2006-05-27|Method of enhancing vibration resistance of pendulum accelerometer
    SU1413696A1|1988-07-30|Electronic drive with damping of oscillation in elastic coupling of motor with mechanism
    SU1055924A2|1983-11-23|Automatically controlled shock-absorber
    SU1462135A1|1989-02-28|Method of checkuing technical condition of machine
    同族专利:
    公开号 | 公开日
    US4706902A|1987-11-17|
    FR2531676B1|1985-02-01|
    FR2531676A1|1984-02-17|
    JPS5984696A|1984-05-16|
    DE3382509D1|1992-03-26|
    EP0101384A1|1984-02-22|
    EP0101384B1|1992-02-12|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US2745613A|1951-12-10|1956-05-15|Douglas Aircraft Co Inc|Gust load alleviating device|
    US3020006A|1954-12-06|1962-02-06|James R Warren|Vibration damping mechanism|
    US3184188A|1961-06-19|1965-05-18|Bendix Corp|Stress sensor and control system for limiting stresses applied to aircraft|
    US3240447A|1962-01-10|1966-03-15|North American Aviation Inc|Vehicle control system|
    US3279725A|1963-12-11|1966-10-18|North American Aviation Inc|Flight controller for flexible vehicles|
    US3236478A|1964-04-23|1966-02-22|North American Aviation Inc|Flight control system|
    GB1086938A|1964-09-16|1967-10-11|Elliott Brothers London Ltd|Aircraft structural fatigue alleviators|
    US3734432A|1971-03-25|1973-05-22|G Low|Suppression of flutter|
    US3902686A|1973-10-17|1975-09-02|Us Air Force|Combined vertical and lateral identical location of accelerometer and force system|
    EP0268041B1|1980-09-02|1992-06-17|Deutsche Airbus GmbH|Servo unit for actuating control surfaces or the like of a flight control system|DE3310510C2|1983-03-23|1988-03-03|Messerschmitt-Boelkow-Blohm Gmbh, 8000 Muenchen, De|
    GB8314656D0|1983-05-26|1983-06-29|Secr Defence|Aircraft control|
    DE3803015A1|1988-02-02|1989-08-10|Pfister Gmbh|Method and system for operatingan aircraft|
    US5458304A|1992-08-26|1995-10-17|Gilbert; Raymond D.|Disk spoiler system|
    US5111766A|1991-01-18|1992-05-12|Ortemund Leon D|Sea wing|
    FR2672028B1|1991-01-29|1993-05-14|Aerospatiale|SYSTEM FOR IMPROVING THE FLOATING BEHAVIOR OF AN AIRCRAFT.|
    US5150864A|1991-09-20|1992-09-29|Georgia Tech Research Corporation|Variable camber control of airfoil|
    US5319296A|1991-11-04|1994-06-07|Boeing Commercial Airplane Group|Oscillatory servo-valve monitor|
    US5374011A|1991-11-13|1994-12-20|Massachusetts Institute Of Technology|Multivariable adaptive surface control|
    EP0696982B1|1993-05-06|2000-08-23|Northrop Grumman Corporation|Apparatus and method for controlling the shape of structures|
    US5588800B1|1994-05-31|2000-12-19|Mcdonell Douglas Helicopter Co|Blade vortex interaction noise reduction techniques for a rotorcraft|
    US5564656A|1994-08-29|1996-10-15|Gilbert; Raymond D.|Segmented spoilers|
    US5549260A|1995-01-27|1996-08-27|Dynamic Engineering, Inc.|Active control device for aircraft tail buffet alleviation|
    DE59510549D1|1995-03-14|2003-03-13|Sulzer Markets & Technology Ag|Method for actively damping global flow oscillations in separated unstable flows and device for applying the method|
    US5669582A|1995-05-12|1997-09-23|The Boeing Company|Method and apparatus for reducing unwanted sideways motion in the aft cabin and roll-yaw upsets of an airplane due to atmospheric turbulence and wind gusts|
    US5791875A|1996-09-10|1998-08-11|Mcdonnell Douglas Helicopter Co.|Tip vortex reduction system|
    US5813625A|1996-10-09|1998-09-29|Mcdonnell Douglas Helicopter Company|Active blowing system for rotorcraft vortex interaction noise reduction|
    US6543719B1|1997-06-05|2003-04-08|Mcdonnell Douglas Helicopter Co.|Oscillating air jets for implementing blade variable twist, enhancing engine and blade efficiency, and reducing drag, vibration, download and ir signature|
    US6234751B1|1997-06-05|2001-05-22|Mcdonnell Douglas Helicopter Co.|Oscillating air jets for reducing HSI noise|
    US5938404A|1997-06-05|1999-08-17|Mcdonnell Douglas Helicopter Company|Oscillating air jets on aerodynamic surfaces|
    US6092990A|1997-06-05|2000-07-25|Mcdonnell Douglas Helicopter Company|Oscillating air jets for helicopter rotor aerodynamic control and BVI noise reduction|
    DE19835191C1|1998-08-04|2000-04-20|Daimler Chrysler Ag|Flight control device for improving the longitudinal stability of a controlled aircraft|
    DE19841632C2|1998-09-11|2001-06-07|Daimler Chrysler Ag|Method for compensating structural vibrations of an aircraft due to external disturbances|
    FR2792285B1|1999-04-16|2001-06-08|Onera |AERODYNAMIC SURFACE OF AIRCRAFT WITH LEAK EDGE DEFLECTOR|
    DE19923087B4|1999-05-20|2004-02-26|Eads Deutschland Gmbh|Device for pressure, sound and vibration measurement, and method for flow analysis on component surfaces|
    US6375127B1|2000-07-07|2002-04-23|Kari Appa|Active control surface modal system for aircraft buffet and gust load alleviation and flutter suppression|
    US6478541B1|2001-08-16|2002-11-12|The Boeing Company|Tapered/segmented flaps for rotor blade-vortex interactionnoise and vibration reduction|
    EP1478569B1|2002-01-30|2009-09-23|Gulfstream Aerospace Corporation|Fuselage shaping and inclusion of spike on a supersonic aircraft for controlling and reducing sonic boom|
    US6698684B1|2002-01-30|2004-03-02|Gulfstream Aerospace Corporation|Supersonic aircraft with spike for controlling and reducing sonic boom|
    FR2838101B1|2002-04-08|2004-12-17|Airbus France|AIRCRAFT WITH ELECTRIC FLIGHT CONTROLS, PROVIDED WITH A FUSELAGE LIKELY TO DEFORM AND VIBRATE|
    US6915989B2|2002-05-01|2005-07-12|The Boeing Company|Aircraft multi-axis modal suppression system|
    FR2841211B1|2002-06-21|2004-12-17|Airbus France|METHOD AND DEVICE FOR REDUCING THE VIBRATORY MOVEMENTS OF THE FUSELAGE OF AN AIRCRAFT|
    US6659397B1|2002-10-18|2003-12-09|Richard Charron|Control system for ornithopter|
    US6766981B2|2002-10-25|2004-07-27|Northrop Grumman Corporation|Control system for alleviating a gust load on an aircraft wing|
    US6685143B1|2003-01-03|2004-02-03|Orbital Research Inc.|Aircraft and missile forebody flow control device and method of controlling flow|
    FR2852577B1|2003-03-19|2006-01-06|Centre Nat Rech Scient|DEVICE FOR CONTROLLING AIR FLOW AROUND AERODYNAMIC PROFILE FOR AN AIRCRAFT TYPE OF AIRCRAFT WING|
    FR2853094B1|2003-03-26|2005-05-27|Airbus France|METHOD FOR CONTROLLING VIBRATION INDUCED IN AN AIRCRAFT BY THE ROTATING OPERATION OF A BLOWER AND ELECTRIC FLIGHT CONTROL SYSTEM USING THE SAME|
    EP2146263B1|2003-11-03|2019-05-08|The Boeing Company|Aircraft multi-axis modal suppression system|
    US7510149B2|2004-08-02|2009-03-31|Lockheed Martin Corporation|System and method to control flowfield vortices with micro-jet arrays|
    US7367530B2|2005-06-21|2008-05-06|The Boeing Company|Aerospace vehicle yaw generating systems and associated methods|
    US20070114327A1|2005-11-18|2007-05-24|The Boeing Company|Wing load alleviation apparatus and method|
    CA2998361C|2005-12-15|2020-04-28|Gulfstream Aerospace Corporation|Isentropic compression inlet for supersonic aircraft|
    US8209071B2|2007-04-16|2012-06-26|Raytheon Company|Methods and apparatus for aircraft turbulence detection|
    US8393158B2|2007-10-24|2013-03-12|Gulfstream Aerospace Corporation|Low shock strength inlet|
    US8090485B2|2007-11-27|2012-01-03|Embraer S.A.|Low-frequency flight control system oscillatory faults prevention via horizontal and vertical tail load monitors|
    FR2929241B1|2008-03-31|2010-06-04|Airbus France|METHOD AND DEVICE FOR REDUCING THE TREMBLING OF AN AIRCRAFT|
    US9302766B2|2008-06-20|2016-04-05|Aviation Partners, Inc.|Split blended winglet|
    FR2946430B1|2009-06-03|2011-08-26|Airbus France|METHOD AND DEVICE FOR DETERMINING CRITICAL LOADS DUE TO A TREMBLING ON A STRUCTURE OF AN AIRCRAFT.|
    DE102010028311A1|2010-04-28|2011-11-03|Airbus Operations Gmbh|System and method for minimizing buffeting|
    ES2783984T3|2011-06-09|2020-09-21|Aviation Partners Inc|Split spiroid|
    EP3176074B1|2015-12-03|2019-11-27|Airbus Operations GmbH|System and method for detecting mechanical failure in the high lift system of an aircraft|
    CA3029128A1|2016-06-30|2018-01-04|Bombardier Inc.|Split-aileron control|
    CN108885466A|2017-11-22|2018-11-23|深圳市大疆创新科技有限公司|A kind of control parameter configuration method and unmanned plane|
    GB2569175A|2017-12-08|2019-06-12|Airbus Operations Ltd|Aircraft with active support|
    法律状态:
    优先权:
    申请号 | 申请日 | 专利标题
    FR8213998A|FR2531676B1|1982-08-11|1982-08-11|
    [返回顶部]