| 序号 | 专利名 | 申请号 | 申请日 | 公开（公告）号 | 公开（公告）日 | 发明人 |
|---|---|---|---|---|---|---|
| 1 | 结构连接器 | CN201610856773.9 | 2016-09-28 | CN106553750A | 2017-04-05 | J·G·伊尔斯 |
| 本发明所公开了包括第一部件和第二部件以及结构连接器的组件。该结构连接器包括第一连接元件和与第一连接元件互连的第二连接元件。第一和第二连接元件中的每一个均被附接到第一和第二部件中的每一个。 | ||||||
| 2 | 用于飞行器的飞行控制装置 | CN201580010945.4 | 2015-02-04 | CN106061837A | 2016-10-26 | B·蒂利; N·拉沃克斯 |
| 本发明涉及一种用于飞行器的飞行控制装置(1)，该装置包括底座(2)、枢转地安装在底座上的操作杆(3)以及用于产生操作杆的回复力的机械工具，所述工具包括弹簧(6)和第一电机构件(13)，弹簧(6)和第一电机构件(13)布置成使得弹簧的第一端被约束成随操作杆旋转运动，而弹簧的第二端被约束成随第一电机构件的输出轴旋转运动。根据本发明，飞行控制装置包括电辅助系统(18)，电辅助系统(18)用来辅助用于产生回复力的所述机械工具。 | ||||||
| 3 | 用于飞行器中的可活动尾翼配平的方法及设备 | CN200880014064.X | 2008-03-17 | CN101674981B | 2012-06-27 | 托尔斯滕·霍尔茨豪森 |
| 本发明涉及一种致动用于调节升降舵(12)的调节驱动装置和用于调节可活动尾翼(23)的调节驱动装置的方法，该方法包括如下步骤：生成升降舵命令用以致动升降舵调节驱动装置；以使可活动尾翼(23)跟踪升降舵输入信号(10)的方式计算用于致动可活动尾翼调节驱动装置的可活动尾翼命令(IHC1)；根据升降舵(12)的调节状态和/或可活动尾翼(23)的调节状态或飞行状态，保持可活动尾翼调节驱动装置的调节状态或利用用于改变可活动尾翼(23)调节状态的可活动尾翼命令(IHCMD)来致动可活动尾翼调节驱动装置，在利用用于改变升降舵(23)调节状态的升降舵命令来致动升降舵调节驱动装置期间以及在计算出的可活动尾翼命令(IHC1)偏离于命令的可活动尾翼命令(IHCMD)的情况下，按照升降舵命令来补偿可活动尾翼调节驱动装置的调节状态的保持。本发明还涉及一种执行该方法的控制设备。 | ||||||
| 4 | 飞行器的高升力系统 | CN201110299632.9 | 2011-09-28 | CN102442425A | 2012-05-09 | R.克兰汉斯; S.胡思; W.韦斯; B.希弗尔布什 |
| 本发明涉及一种飞行器的高升力系统，具有至少一个驱动装置、至少一个负载站以及一个或多个用于传递该驱动装置的驱动能量给该至少一个负载站的传动装置，其中的一个或多个所述传动装置以传动轴的形式构成，该传动轴由含钛材料构成或具有含钛材料。 | ||||||
| 5 | 用于控制飞行器机翼襟翼运动的系统和方法 | CN201710228710.3 | 2017-04-10 | CN107303946A | 2017-10-31 | M·R·芬; P·麦考密克; G·莫伊; R·T·源 |
| 本申请涉及用于控制飞行器机翼襟翼运动的系统和方法。其包括：从联接至相应的第一致动器(112)和第二致动器(114)的相应的第一传感器(120)和第二传感器(122)接收第一传感器信号和第二传感器信号。第一传感器信号和第二传感器信号与相应的第一致动器和第二致动器的位置或速度中的一个或两个有关。该系统和方法还包括：比较第一传感器信号与第二传感器信号，以确定第一传感器信号与第二传感器信号之间的差异，并且基于该差异来调节第一致动器或第二致动器中的一个或两个的速度。该系统和方法包括：确定第一襟翼(102)和第二襟翼(108)的速度或位置中的一个或两个之间的差异，并且基于该差异来调节第一襟翼和第二襟翼中的一个或两个的速度。 | ||||||
| 6 | 飞行器的高升力系统 | CN201110299632.9 | 2011-09-28 | CN102442425B | 2016-12-28 | R.克兰汉斯; S.胡思; W.韦斯; B.希弗尔布什 |
| 本发明涉及一种飞行器的高升力系统，具有至少一个驱动装置、至少一个负载站以及一个或多个用于传递该驱动装置的驱动能量给该至少一个负载站的传动装置，其中的一个或多个所述传动装置以传动轴的形式构成，该传动轴由含钛材料构成或具有含钛材料。 | ||||||
| 7 | 高完整性旋转式促动器和运行方法 | CN201210102492.6 | 2012-03-29 | CN102730186A | 2012-10-17 | T·R·戈丁 |
| 本发明涉及高完整性旋转式促动器和运行方法。一种用于航空器的促动器(1)包括第一驱动器件(2)和第二驱动器件(3)以及促动器输出(10)，它们由齿轮组件(4)互连，借助于齿轮组件(4)：促动器输出(10)能够由第一驱动器件(2)独立于第二驱动器件(3)而驱动；以及促动器输出(10)能够由第二驱动器件(3)独立于第一驱动器件(2)而驱动；以及促动器输出(10)能够由第一驱动器件和第二驱动器件(2，3)结合起来驱动。齿轮组件(4)包括一组行星齿轮。 | ||||||
| 8 | 用于飞行器中的可活动尾翼配平的方法及设备 | CN200880014064.X | 2008-03-17 | CN101674981A | 2010-03-17 | 托尔斯滕·霍尔茨豪森 |
| 本发明涉及一种致动用于调节升降舵(12)的调节驱动装置和用于调节可活动尾翼(23)的调节驱动装置的方法，该方法包括如下步骤：生成升降舵命令用以致动升降舵调节驱动装置；以使可活动尾翼(23)跟踪升降舵输入信号(10)的方式计算用于致动可活动尾翼调节驱动装置的可活动尾翼命令(IHC1)；根据升降舵(12)的调节状态和/或可活动尾翼(23)的调节状态或飞行状态，保持可活动尾翼调节驱动装置的调节状态或利用用于改变可活动尾翼(23)调节状态的可活动尾翼命令(IHCMD)来致动可活动尾翼调节驱动装置，在利用用于改变升降舵(23)调节状态的升降舵命令来致动升降舵调节驱动装置期间以及在计算出的可活动尾翼命令(IHC1)偏离于命令的可活动尾翼命令(IHCMD)的情况下，按照升降舵命令来补偿可活动尾翼调节驱动装置的调节状态的保持。本发明还涉及一种执行该方法的控制设备。 | ||||||
| 9 | Non-jamming screw actuator device | JP30031887 | 1987-11-30 | JPS63152761A | 1988-06-25 | BURENTO EI KUROTSUPENSUTEIN |
| 10 | Aircraft wing structure and control system | US14907067 | 2014-07-25 | US09846432B2 | 2017-12-19 | Michael Lam; Greg Cole |
| An aircraft includes a wing. The wing includes an aileron pivotally connected to a trailing edge of the wing, and a Lam aileron pivotally connected to the trailing edge of the wing. The aircraft includes a motor connected to the Lam aileron and configured to rotate the Lam aileron. The aircraft includes a controller configured to detect a deflection of the aileron from a neutral position, calculate a target deflection for the Lam aileron using the deflection of the aileron, and cause the motor to rotate the Lam aileron to the target deflection. | ||||||
| 11 | AIRCRAFT WING STRUCTURE AND CONTROL SYSTEM | US14907067 | 2014-07-25 | US20160161949A1 | 2016-06-09 | Michael LAM; Greg COLE |
| An aircraft includes a wing. The wing includes an aileron pivotally connected to a trailing edge of the wing, and a Lam aileron pivotally connected to the trailing edge of the wing. The aircraft includes a motor connected to the Lam aileron and configured to rotate the Lam aileron. The aircraft includes a controller configured to detect a deflection of the aileron from a neutral position, calculate a target deflection for the Lam aileron using the deflection of the aileron, and cause the motor to rotate the Lam aileron to the target deflection. | ||||||
| 12 | AIRCRAFT FLAP SYSTEM WITH AILERON FUNCTIONALITY | US14285712 | 2014-05-23 | US20150191240A1 | 2015-07-09 | Alexander BURCHARD |
| The present disclosure pertains to an actuation mechanism for a flap with aileron functionality, including a crank having a crank axle, a crank arm and a crank pivot; a displacement shaft articulated to the crank pivot at an actuating end portion; and a rotatable linear-motion bearing in which the displacement shaft is slidably supported, wherein the displacement shaft is fixedly connectable to a flap body of the aircraft flap at a flap end portion of the displacement shaft opposite to the actuating end portion, and wherein the linear-motion bearing is rotatably attachable at a pivot point to an aircraft wing comprising the aircraft flap, so that the displacement shaft is able to rotate around an axis running spanwise of the aircraft wing around the pivot point. The present disclosure further pertains to an aircraft flap with such an actuation mechanism and an aircraft having an aircraft flap. | ||||||
| 13 | Deployable Aerodynamic Devices with Reduced Actuator Loads | US13150334 | 2011-06-01 | US20110226345A1 | 2011-09-22 | Glenn S. Bushnell |
| Deployable aerodynamic devices with reduced actuator loads, and related systems and methods are disclosed. An external flow system in accordance with a particular embodiment includes an external flow body, a deployable device carried by and movable relative to the external flow body, and a coupling connected between the external flow body and the deployable device. The system can further include an actuator device operatively coupled between the external flow body and the deployable device, with the actuator device positioned to move the deployable device along a motion path between a stowed position and the deployed position. The motion path can have a first portion over which the load delivered by the actuator device increases as the deployed device moves toward the deployed position, and a second portion over which the load delivered by the actuator device decreases as the deployed device moves toward the deployed position. | ||||||
| 14 | Mechanisms and methods for providing rudder control assist during symmetrical and asymmetrical thrust conditions | US11869219 | 2007-10-09 | US07984880B2 | 2011-07-26 | Ierko de Magalhaes Gomes |
| Rudder assist mechanisms and methods are capable of being operably connected to an aircraft's rudder control system. The rudder assist mechanisms most preferably have over-the-center spring biasing functions so as to cause either substantially no spring force (i.e., when the linkage is over the spring-bias center) or substantially all spring force (i.e., when the linkage is left or right of the spring-bias center) to be exerted on the rudder control system. The rudder assist mechanism may include a control spring assembly, and a linkage assembly which operably connects the control spring assembly to the rudder control assembly. The linkage assembly is moveable operably between a null position wherein substantially no spring force of the control spring assembly is transferred to the rudder control system by the linkage assembly, and right and left spring-biased positions wherein right and left spring forces of the control spring assembly are transferred to the rudder control system, respectively. The null position may therefore establish a dead zone of rudder deflection within a selected range of right and left rudder control surface deflection angles. Thus, right and left spring forces may be transferred to the rudder control system when the rudder control surface is deflected at an angle which exceeds the selected range of deflection angles. Some embodiments include an actuator unit which is operably capable of moving the spring assembly and the linkage assembly connected thereto between a thrust symmetrical mode (TSM) condition and a thrust asymmetrical mode (TAM) condition so as to cause different spring forces to be exerted on the rudder control system in such different thrust conditions. | ||||||
| 15 | FLIGHT CONTROL SYSTEMS | US11669565 | 2007-01-31 | US20070267548A1 | 2007-11-22 | Philippe Ciholas; Mark Palmer |
| A system and method for a controlling an aircraft with flight control surfaces that are controlled both manually and by a computing device is disclosed. The present invention improves overall flight control operation by reducing the mechanical flight control surface components while providing sufficient back-up control capability in the event of either a mechanical or power-related failure. Through the present invention, natural feedback is provided to the operator from the mechanical flight control surface which operates independent of computer-aided flight control surfaces. | ||||||
| 16 | System pressure compensated variable displacement hydraulic motor | US808422 | 1991-12-16 | US5307630A | 1994-05-03 | John D. Tysver; Bruce A. Krandel; Wesley A. Burandt |
| Paper drive unit instability that includes a variable displacement hydraulic motor (34), a rotary output (48), and a displacement controlling wobbler (36, 40) in the case of an aiding load can be avoided through the use of a double acting hydraulic actuator (42, 44) connected to the wobbler (36, 40) along with a control valve (30) including a single spool (80) connected to the motor (34), and wobbler control pilot valve (62) and to the actuator (42, 44) for a) controlling the speed and direction of rotation of the motor output (48); and b) controlling the actuator (42, 44) to maintain minimum displacement of the motor 34 when an aiding load occurs. System optimization can be achieved by changing the differential pressure setting in response to the available system pressure. | ||||||
| 17 | Electropneumatic rotary actuator having proportional fluid valving | US216753 | 1988-07-08 | US4903578A | 1990-02-27 | Leslie S. Terp |
| An electropneumatic rotary actuator has proportional fluid valving providing introduction and exhaust of pressurized fluid to and from a working chamber of the actuator. By control of the admission angle and venting angle of the fluid pressure and flow with respect to top dead center and bottom dead center of a working piston of the actuator both the torque level and gas consumption rates of the actuator may be controlled. Further by modulating the average working fluid pressure within the actuator the stiffness or resistance to backdriving of the actuator may be controlled. A proportional linear electrical solenoid is disclosed and employed in the actuator of the present invention to drive a double-acting closed-centered valve device. Because of its lightweight, relatively simplicity, high stiffness to backdriving, and relatively low gas consumption rate, the present actuator is expected to find application in airborne vehicles. | ||||||
| 18 | Fluid control system | US39117373 | 1973-08-24 | US3915427A | 1975-10-28 | SWOGGER EMERY C |
| A fluid-powered control system of a type employing redundant control channels and multiple fluid signal amplifiers. Each fluid amplifier has two fluid outlets having communication with respective portions of a movable, force summing structure but having no communication with the other amplifiers. A respective, differential pressure monitoring means is connected between a source of fluid under pressure and each fluid amplifier and is operative to shut off fluid flow to the respective, associated, fluid amplifier upon the occurrence, across the outlets of the fluid amplifier, of a differential pressure exceeding a predetermined level.
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| 19 | Rotary actuator with high maintainability and method for working the same | JP2012072834 | 2012-03-28 | JP2012214218A | 2012-11-08 | GOLDING TERENCE ROSS |
| PROBLEM TO BE SOLVED: To provide a rotary actuator suitable for an aircraft which maximizes reliability, and reduces a dimension, weight, and complexity.SOLUTION: The actuator for an aircraft 1 includes: first and second drive means 2 and 3 connected with each other by a gear assembly 4; and an actuator output shaft 10. The actuator output shaft 10 can be driven by the first drive means 2 independently of the second drive means 3, the actuator output shaft 10 can be driven by the second drive means 3 independently of the first drive means 2, and the actuator output shaft 10 can be driven by the combination of the first drive means 2 and the second drive means 3. The gear assembly 4 includes a pair of planetary gears. | ||||||
| 20 | Kokukyotakinoakuchueetasochi | JP17567880 | 1980-12-12 | JPH0228517B2 | 1990-06-25 | KAARU DEIIN GURIFUISU; KENISU RII ORIBAA |
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