WO2018184396A1 - 飞机的电动脚蹬控制装置 - Google Patents

飞机的电动脚蹬控制装置 Download PDF

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Publication number
WO2018184396A1
WO2018184396A1 PCT/CN2017/113849 CN2017113849W WO2018184396A1 WO 2018184396 A1 WO2018184396 A1 WO 2018184396A1 CN 2017113849 W CN2017113849 W CN 2017113849W WO 2018184396 A1 WO2018184396 A1 WO 2018184396A1
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WIPO (PCT)
Prior art keywords
control device
ankle
motor
transmission mechanism
pedal
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PCT/CN2017/113849
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English (en)
French (fr)
Inventor
谢殿煌
李正强
戴烨飞
李剑
廖军辉
徐德胜
余圣晖
Original Assignee
中国商用飞机有限责任公司
中国商用飞机有限责任公司上海飞机设计研究院
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Application filed by 中国商用飞机有限责任公司, 中国商用飞机有限责任公司上海飞机设计研究院 filed Critical 中国商用飞机有限责任公司
Priority to US16/310,058 priority Critical patent/US11319057B2/en
Priority to EP17904754.3A priority patent/EP3456626B1/en
Publication of WO2018184396A1 publication Critical patent/WO2018184396A1/zh

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C13/00—Control systems or transmitting systems for actuating flying-control surfaces, lift-increasing flaps, air brakes, or spoilers
    • B64C13/02—Initiating means
    • B64C13/04—Initiating means actuated personally
    • B64C13/044—Initiating means actuated personally operated by feet, e.g. pedals
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C13/00—Control systems or transmitting systems for actuating flying-control surfaces, lift-increasing flaps, air brakes, or spoilers
    • B64C13/02—Initiating means
    • B64C13/04—Initiating means actuated personally
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C13/00—Control systems or transmitting systems for actuating flying-control surfaces, lift-increasing flaps, air brakes, or spoilers
    • B64C13/02—Initiating means
    • B64C13/04—Initiating means actuated personally
    • B64C13/12—Dual control apparatus
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C13/00—Control systems or transmitting systems for actuating flying-control surfaces, lift-increasing flaps, air brakes, or spoilers
    • B64C13/24—Transmitting means
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C13/00—Control systems or transmitting systems for actuating flying-control surfaces, lift-increasing flaps, air brakes, or spoilers
    • B64C13/24—Transmitting means
    • B64C13/26—Transmitting means without power amplification or where power amplification is irrelevant
    • B64C13/28—Transmitting means without power amplification or where power amplification is irrelevant mechanical
    • B64C13/30—Transmitting means without power amplification or where power amplification is irrelevant mechanical using cable, chain, or rod mechanisms
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/40—Weight reduction

Definitions

  • the invention relates to the field of aviation, and in particular to an electric pedal control device for an aircraft.
  • the pedal control device is one of the important interfaces of the human-machine interface of the aircraft cockpit. It is the core component of the fly-by-wire flight control system. It mainly provides the steering signal and realizes the manual trimming function and the follow-up function (with the autopilot). ) and the override pedal (disconnect the autopilot).
  • the distributed pedal control device mainly adopts discrete components and distributed layout forms, and the Boeing series aircraft adopts such a layout form.
  • the centralized pedal control unit is mainly integrated with discrete components and a centralized layout.
  • the Airbus series uses this type of layout.
  • the centralized pedal control device is a further structural layout optimization of the distributed pedal drive device, but the actual driving methods of the two are the same, and the design is complicated, the transmission mechanism is complicated, the assembly is difficult, the manufacturing cost is high, and the maintenance cost is low. High, single point of failure is likely to cause a series of defects such as the transmission mechanism card set.
  • the existing conventional ankle control device is difficult to meet the increasing development requirements of the flight control system, namely, reducing weight, reducing maintenance costs, and increasing dispatch rate.
  • the technical problem to be solved by the present invention is to overcome the defects that the transmission mechanism of the conventional ankle control device in the prior art is complicated, the assembly is difficult, the cost is high, and the single point failure easily causes the failure of the transmission mechanism card set, and an aircraft is proposed. Electric pedal control device.
  • the invention provides an electric pedal control device for an aircraft, which is characterized in that it comprises a plurality of sets of ankle drive assembly, each set of an ankle drive assembly comprising a motor, an elastic connecting member, a transmission mechanism, an angular displacement sensor and an ankle.
  • the elastic connecting member connects the electric An output shaft of the machine and one end of the transmission mechanism, the other end of the transmission mechanism is connected to the ankle via an ankle rotation pair, the angular displacement sensor is configured to collect the ankle of the ankle around the ankle Rotation position information, the transmission mechanism has a transmission mechanism rotation pair, and the transmission mechanism rotation pair of the transmission mechanism in each set of ankle transmission assembly is connected via a mechanical linkage mechanism to achieve linkage, the electric pedal control
  • the apparatus further includes a controller for receiving the rotational position information acquired by the angular displacement sensor, and controlling the motor to generate a damping effect on the elastic connecting member according to the rotational position information.
  • the angular displacement sensor employs an RVDT sensor, that is, a rotary variable differential transformer.
  • the controller is further configured to receive the input control command and directly control the operation of the motor according to the control command.
  • control command includes a manual trim command
  • controller is configured to directly control all or part of the motor according to the manual trim command to balance the elastic force of the elastic link.
  • the controller rejects The operation of the motor is directly controlled according to the input control command.
  • the output shaft of the motor is coupled to the resilient connector via a torque limiter.
  • the elastic connecting member is a spring.
  • the electric pedal control device further includes an output interface for transmitting the rotational position information to a flight control computer of the aircraft.
  • the plurality of sets of ankle drive components are two sets of left ankle drive assemblies and two sets of right ankle drive assemblies, and the two sets of left ankle drive assemblies and the two sets of right ankle drive assemblies share one
  • the transmission mechanism rotation pair, the transmission mechanism rotation pair shared by the two sets of left ankle transmission components and the transmission mechanism rotation pair shared by the two sets of right ankle transmission components are connected via a mechanical linkage mechanism to achieve linkage.
  • one end of the elastic connecting member is connected to the output shaft of the motor, and the other end is connected to the corresponding rotating portion of the transmission mechanism via a connecting rod, and the corresponding rotating shaft of the transmission mechanism is transmitted through the transmission member.
  • the ankle rotation pair is coupled to the ankle.
  • the electric pedal control device of the aircraft of the invention can simplify the transmission mechanism, reduce the volume, and is easy to maintain and reconfigurable, thereby avoiding the failure of the transmission mechanism card group caused by a single point of failure, and helping To improve the dispatch rate.
  • FIG. 1 is a schematic view of an electric pedal control device according to a preferred embodiment of the present invention.
  • an electric pedal control device for an aircraft can be configured with four sets of ankle drive assembly to form a four-degree electric pedal control device.
  • other embodiments of the electric pedal control device according to other preferred embodiments of the present invention may employ other numbers of ankle drive assemblies.
  • each set of ankle drive assembly includes a motor 1, an elastic link 3, a transmission mechanism, an angular displacement sensor 5, and an ankle (not shown), the elastic connection
  • the piece 3 is connected via the torque limiter 2 to the output shaft of the motor 1 and one end of the transmission.
  • the other end of the transmission mechanism is connected to the ankle via an ankle rotation pair 4 for collecting rotational position information of the ankle around the ankle rotation pair 4, the electric foot
  • the ⁇ control device further includes a controller for receiving the rotational position information collected by the angular displacement sensor 5, and controlling the motor 1 to generate the elastic connecting member 3 according to the rotational position information. Damping effect.
  • the transmission mechanism has a transmission rotating pair 6 , and the transmission rotating pair 6 of each set of the pedal transmission assembly is connected via the mechanical linkage mechanism 7 to achieve linkage.
  • an aircraft that uses hot backup for an ankle control device only needs to cut off the corresponding motor output when encountering a typical single point of failure situation, such as when a resilient connection member 3 fails to open,
  • a typical single point of failure situation such as when a resilient connection member 3 fails to open
  • FIG. 1 is a schematic diagram showing the connection and transmission relationship between the pedal and the motor in the electric pedal control device of the aircraft according to a preferred embodiment of the present invention, and between the plurality of sets of ankle drive assemblies.
  • the linkage relationship but does not limit the invention, must be implemented in the structure shown in FIG.
  • the two sets of left ankle drive assemblies and the two sets of right ankle drive assemblies may each share a transmission rotation pair 6, the transmissions shared by the two sets of left ankle drive assemblies
  • the mechanism rotating pair 6 is coupled to the two sets of right ankle drive assemblies to be coupled via a mechanical linkage 7 to effect linkage.
  • the angular displacement sensor 5 in the four-degree electric pedal control device may employ an RVDT sensor, each RVDT sensor corresponding to one foot, and the motor assembly may include a micro motor and an optical encoder disk.
  • the elastic connecting member 3 can adopt a spring.
  • the controller collects a signal obtained by the RVDT sensor to characterize the movement of the ankle position, thereby controlling the motor to generate a damping function, and facilitating the damping and force sensing functions of the ankle control device by using the motor and the spring.
  • the electric pedal control device of the present invention may further include an output interface for transmitting the rotational position information to The flight control computer of the aircraft.
  • the controller is further configured to receive the input control command and directly control the operation of the motor according to the control command.
  • control instruction may include a manual trimming instruction
  • controller is configured to directly control all or part of the motor according to the manual trimming instruction to balance the elastic force of the elastic connecting member.
  • the driver only needs to control some or all of the motors in the electric pedal control device through the trim switch operation, and output in different preset modes, thereby performing force balance, for example, for the above four-degree electric pedal control.
  • the unit can directly control all four motors through the trim switch.
  • the automatic flight computer can also accept the control command provided by one end of the automatic flight computer to directly control the forward or backward movement of any motor.
  • the controller refuses to directly control the operation of the motor according to the input control command.
  • the override function of the electric pedal control device can be realized. That is, the driver can force the pedal to force the RVDT sensor to generate a sufficiently strong current signal, and the controller turns off the passage of the motor according to the control command according to the current signal, and only realizes the function of the damping portion, and the position control cannot be realized.
  • the above-described preferred embodiment of the present invention reduces the size and weight of the entire electric pedal control device as compared to the conventional distributed and centralized layout, and contributes to an improved dispatch rate.

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  • Engineering & Computer Science (AREA)
  • Automation & Control Theory (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Mechanical Engineering (AREA)
  • Mechanical Control Devices (AREA)

Abstract

一种飞机的电动脚蹬控制装置，包括若干组脚蹬传动组件，每组脚蹬传动组件包括电机（1）、弹性连接件（3）、传动机构、角位移传感器（5）及脚蹬，角位移传感器（5）用于采集脚蹬的转动位置信息，各组脚蹬传动组件中的传动机构的传动机构转动副（6）经由机械连杆机构（7）连接以实现联动，电动脚蹬控制装置的控制器用于接收转动位置信息，并根据转动位置信息控制电机（1）对弹性连接件（3）产生阻尼作用。该电动脚蹬控制装置，能够使得传动机构得到简化，同时还减小了体积，并易于维护、遇到单点故障时的重构性强，有助于提高派遣率。

Description

飞机的电动脚蹬控制装置 技术领域
本发明涉及航空领域，尤其涉及一种飞机的电动脚蹬控制装置。
背景技术
脚蹬控制装置是飞机驾驶舱的人机界面的重要接口之一，是电传飞控系统的核心部件，主要提供方向舵的操纵信号和实现人工配平功能、随动功能(与自动驾驶仪随动)和超控脚蹬(断开自动驾驶仪)。
传统的大型民用飞机脚蹬控制装置主要分为两类：分布式脚蹬控制装置和集中式脚蹬控制装置。分布式脚蹬控制装置主要采用了分立部件和分布式布局形式，波音系列飞机采用此类布局形式。集中式脚蹬控制装置主要集成了分立部件和集中式布局形式，空客系列飞机采用此类布局形式。集中式脚蹬控制装置是分布式脚蹬驱动装置的进一步结构布局优化，但是两者实际上的驱动方式是相同的，都存在着设计复杂，传动机构复杂、装配困难，制造成本高、维护成本高，单点故障容易造成传动机构卡组等一系列缺陷。
因而，现有的传统脚蹬控制装置很难满足日益增进的飞控系统的研制需求，即减轻重量、降低维护成本和提高派遣率。
发明内容
本发明要解决的技术问题是为了克服现有技术中的传统脚蹬控制装置的传动机构复杂、装配困难，成本高，以及单点故障容易造成传动机构卡组故障的缺陷，提出一种飞机的电动脚蹬控制装置。
本发明是通过下述技术方案来解决上述技术问题的：
本发明提供了一种飞机的电动脚蹬控制装置，其特点在于，其包括若干组脚蹬传动组件，每组脚蹬传动组件包括电机、弹性连接件、传动机构、角位移传感器及脚蹬，所述弹性连接件连接所述电 机的输出轴与所述传动机构的一端，所述传动机构的另一端经由脚蹬转动副连接至所述脚蹬，所述角位移传感器用于采集所述脚蹬绕所述脚蹬转动副的转动位置信息，所述传动机构具有传动机构转动副，各组脚蹬传动组件中的所述传动机构的所述传动机构转动副经由机械连杆机构连接以实现联动，所述电动脚蹬控制装置还包括控制器，所述控制器用于接收由所述角位移传感器采集到的所述转动位置信息，并根据所述转动位置信息控制所述电机对所述弹性连接件产生阻尼作用。
较佳地，角位移传感器采用RVDT传感器，即旋转可变差动变压器。
较佳地，所述控制器还用于接收输入的控制指令，并根据所述控制指令直接控制所述电机的运转。
较佳地，所述控制指令包括人工配平指令，所述控制器用于根据所述人工配平指令直接控制全部或部分的所述电机以配平所述弹性连接件的弹性力。
较佳地，当所述控制器接收到的所述转动位置信息表征所述脚蹬中的至少一个绕所述脚蹬转动副的转动加速度超出预设的转动加速度阈值时，所述控制器拒绝根据输入的所述控制指令直接控制所述电机的运转。
较佳地，所述电机的输出轴经由扭矩限制器连接至所述弹性连接件。
较佳地，所述弹性连接件为弹簧。
较佳地，所述电动脚蹬控制装置还包括输出接口，所述输出接口用于将所述转动位置信息传输至飞机的飞控计算机。
较佳地，所述若干组脚蹬传动组件为两组左脚蹬传动组件和两组右脚蹬传动组件，所述两组左脚蹬传动组件和所述两组右脚蹬传动组件共用一个传动机构转动副，所述两组左脚蹬传动组件共用的传动机构转动副与所述两组右脚蹬传动组件共用的传动机构转动副经由机械连杆机构连接以实现联动。
较佳地，每组脚蹬传动组件中，所述弹性连接件的一端连接所述电机的输出轴，另一端经由连杆连接至相应的传动机构转动副，相应的传动机构转动副经由传动部件及所述脚蹬转动副连接至所述脚蹬。
在符合本领域常识的基础上，上述各优选条件，可任意组合，即得本发明各较佳实例。
本发明的积极进步效果在于：
本发明的飞机的电动脚蹬控制装置，能够使得传动机构得到简化，同时还减小了体积，并易于维护、重构性强，避免了单点故障造成传动机构卡组的故障发生，有助于提高派遣率。
附图说明
图1为本发明一较佳实施例的电动脚蹬控制装置的示意图。
具体实施方式
下面结合说明书附图，进一步对本发明的优选实施例进行详细描述，以下的描述为示例性的，并非对本发明的限制，任何的其他类似情形也都落入本发明的保护范围之中。
在以下的具体描述中，方向性的术语，例如“左”、“右”、“上”、“下”、“前”、“后”、等，参考附图中描述的方向使用。本发明的实施例的部件可被置于多种不同的方向，方向性的术语是用于示例的目的而非限制性的。
参考图1所示，根据本发明一较佳实施例的飞机的电动脚蹬控制装置可以四组脚蹬传动组件，从而构成一种四余度电动脚蹬控制装置。当然，根据本发明其他较优实施方式的电动脚蹬控制装置也可以采用其他数量的脚蹬传动组件。
参考图1所示，下方的框中为两组左脚蹬传动组件，上方的框中为两组右脚蹬传动组件。每组脚蹬传动组件包括电机1、弹性连接件3、传动机构、角位移传感器5及脚蹬(未示出)，所述弹性连接 件3经由扭矩限制器2连接所述电机1的输出轴与所述传动机构的一端。所述传动机构的另一端经由脚蹬转动副4连接至所述脚蹬，所述角位移传感器5用于采集所述脚蹬绕所述脚蹬转动副4的转动位置信息，所述电动脚蹬控制装置还包括控制器，所述控制器用于接收由所述角位移传感器5采集到的所述转动位置信息，并根据所述转动位置信息控制所述电机1对所述弹性连接件3产生阻尼作用。所述传动机构具有传动机构转动副6，各组脚蹬传动组件中的传动机构转动副6经由机械连杆机构7连接以实现联动。
根据本发明的一个方面，针对脚蹬控制装置采用热备份的飞机，在遇到典型的单点故障情形时，例如当某一个弹性连接件3失效断开时，只需切断相应的电机输出，并对剩余三个电机或者两个电机或者其他数量的电机通过适当的重构算法重新计算其输出，就能利用剩余电机实现力平衡的控制。
应当理解的是，图1示意性地示出了本发明一较佳实施例的飞机的电动脚蹬控制装置中，自脚蹬至电机的连接及传动关系，以及多组脚蹬传动组件之间的联动关系，但并不因此限制本发明必须以图1所示的结构实施。
根据本发明的一个方面，可选地，两组左脚蹬传动组件和所述两组右脚蹬传动组件可以各自共用一个传动机构转动副6，所述两组左脚蹬传动组件共用的传动机构转动副6与所述两组右脚蹬传动组件共用的传动机构转动副6经由机械连杆机构7连接以实现联动。
根据本发明的另一个方面，四余度电动脚蹬控制装置中的角位移传感器5可采用RVDT传感器，每个RVDT传感器对应于一个脚蹬，电机组件可包含微电机及光电编码盘。弹性连接件3可采用弹簧。通过由弹簧产生力感，控制器采集RVDT传感器得到的表征脚蹬位置运动的信号，从而控制电机产生阻尼功能，便利用电机和弹簧实现了脚蹬控制装置的阻尼和力感功能。
根据本发明的一些优选实施方式，本发明的电动脚蹬控制装置还可包括输出接口，所述输出接口用于将所述转动位置信息传输至 飞机的飞控计算机。
根据本发明的一些优选实施方式，所述控制器还用于接收输入的控制指令，并根据所述控制指令直接控制所述电机的运转。
其中，所述控制指令可包括人工配平指令，所述控制器用于根据所述人工配平指令直接控制全部或部分的所述电机以配平所述弹性连接件的弹性力。
由此，驾驶者只需通过配平开关操作来控制电动脚蹬控制装置中的部分或全部电机，以不同的预设模式输出，从而进行力感配平，例如，对于上述四余度电动脚蹬控制装置，可通过配平开关直接控制全部四个电机。
此外，还可接受自动飞行计算机一端提供的控制指令，直接控制任何一个电机前进或者后退运动，力感配平驾驶者通过脚蹬控制任何一个电机的位置及力感，实现随动功能。
根据本发明的另一优选实施方式，当所述控制器接收到的所述转动位置信息表征所述脚蹬中的至少一个绕所述脚蹬转动副的转动加速度超出预设的转动加速度阈值时，所述控制器拒绝根据输入的所述控制指令直接控制所述电机的运转。由此，即可实现电动脚蹬控制装置的超控功能。即，驾驶者可以通过用力蹬脚蹬，使得RVDT传感器产生足够强的电流信号，控制器根据这一电流信号断开根据控制指令控制电机的通路，仅实现阻尼部分功能，而无法实现位置控制。
本发明的上述优选实施方式，相较于传统的分布式和集中式布局，减小了整个电动脚蹬控制装置的体积和重量，有助于提高派遣率。
虽然以上描述了本发明的具体实施方式，但是本领域的技术人员应当理解，这些仅是举例说明，本发明的保护范围是由所附权利要求书限定的。本领域的技术人员在不背离本发明的原理和实质的前提下，可以对这些实施方式做出多种变更或修改，但这些变更和修改均落入本发明的保护范围。

Claims (9)

  1. 一种飞机的电动脚蹬控制装置，其特征在于，其包括若干组脚蹬传动组件，每组脚蹬传动组件包括电机、弹性连接件、传动机构、角位移传感器及脚蹬，所述弹性连接件连接所述电机的输出轴与所述传动机构的一端，所述传动机构的另一端经由脚蹬转动副连接至所述脚蹬，所述角位移传感器用于采集所述脚蹬绕所述脚蹬转动副的转动位置信息，所述传动机构具有传动机构转动副，各组脚蹬传动组件中的所述传动机构的所述传动机构转动副经由机械连杆机构连接以实现联动，所述电动脚蹬控制装置还包括控制器，所述控制器用于接收由所述角位移传感器采集到的所述转动位置信息，并根据所述转动位置信息控制所述电机对所述弹性连接件产生阻尼作用。
  2. 如权利要求1所述的电动脚蹬控制装置，其特征在于，所述控制器还用于接收输入的控制指令，并根据所述控制指令直接控制所述电机的运转。
  3. 如权利要求2所述的电动脚蹬控制装置，其特征在于，所述控制指令包括人工配平指令，所述控制器用于根据所述人工配平指令直接控制全部或部分的所述电机以配平所述弹性连接件的弹性力。
  4. 如权利要求2所述的电动脚蹬控制装置，其特征在于，当所述控制器接收到的所述转动位置信息表征所述脚蹬中的至少一个绕所述脚蹬转动副的转动加速度超出预设的转动加速度阈值时，所述控制器拒绝根据输入的所述控制指令直接控制所述电机的运转。
  5. 如权利要求1所述的电动脚蹬控制装置，其特征在于，所述电机的输出轴经由扭矩限制器连接至所述弹性连接件。
  6. 如权利要求1所述的电动脚蹬控制装置，其特征在于，所述弹性连接件为弹簧。
  7. 如权利要求1所述的电动脚蹬控制装置，其特征在于，所述 电动脚蹬控制装置还包括输出接口，所述输出接口用于将所述转动位置信息传输至飞机的飞控计算机。
  8. 如权利要求1-7中任意一项所述的电动脚蹬控制装置，其特征在于，所述若干组脚蹬传动组件为两组左脚蹬传动组件和两组右脚蹬传动组件，所述两组左脚蹬传动组件和所述两组右脚蹬传动组件共用一个传动机构转动副，所述两组左脚蹬传动组件共用的传动机构转动副与所述两组右脚蹬传动组件共用的传动机构转动副经由机械连杆机构连接以实现联动。
  9. 如权利要求8所述的电动脚蹬控制装置，其特征在于，每组脚蹬传动组件中，所述弹性连接件的一端连接所述电机的输出轴，另一端经由连杆连接至相应的传动机构转动副，相应的传动机构转动副经由传动部件及所述脚蹬转动副连接至所述脚蹬。
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Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN110654531B (zh) * 2019-09-19 2022-05-03 中国商用飞机有限责任公司 一种用于飞机方向舵控制的脚蹬系统及其控制方法
US11396362B2 (en) 2019-11-01 2022-07-26 Woodward, Inc. Rudder and brake pedal assembly
CN110920873A (zh) * 2019-12-13 2020-03-27 北京青云航空仪表有限公司 一种基于扭力杆的脚蹬刹车机构
CN112173081B (zh) * 2020-09-25 2022-05-24 兰州万里航空机电有限责任公司 一种方向舵调整片操纵装置的数字化控制器
CN113460288B (zh) * 2021-09-03 2021-12-10 中国商用飞机有限责任公司 用于控制飞行器的主动式方向舵脚蹬组件以及飞行器

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3999726A (en) * 1975-05-29 1976-12-28 Textron, Inc. Helicopter pylon-fuselage coupling for expanded CG range
DE3825809A1 (de) * 1988-07-29 1990-02-08 Messerschmitt Boelkow Blohm Steuereinrichtung mit einem steuerknueppel
US20110108674A1 (en) * 2009-11-10 2011-05-12 Cessna Aircraft Company Cockpit Rudder Control Mechanism For An Aircraft
CN103057697A (zh) * 2013-01-07 2013-04-24 中国商用飞机有限责任公司 一种用于飞机的方向舵脚蹬控制装置和控制方法
CN103587679A (zh) * 2013-11-28 2014-02-19 江西洪都航空工业集团有限责任公司 机械式飞机脚蹬调节机构

Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3576302A (en) * 1968-07-10 1971-04-27 Bendix Corp Solid-state position sensor for sensing an adjusted position of a control element
US4470570A (en) * 1982-09-29 1984-09-11 The Boeing Company Control assembly for aircraft
US5797564A (en) * 1995-05-15 1998-08-25 The Boeing Company System for backdrive of flight deck controls during autopilot operation
US5806806A (en) * 1996-03-04 1998-09-15 Mcdonnell Douglas Corporation Flight control mechanical backup system
DE19931865A1 (de) * 1999-07-09 2001-01-11 Schopf Maschb Gmbh Schleppfahrzeug für Flugzeuge
US7644893B2 (en) * 2006-02-15 2010-01-12 Sikorsky Aircraft Corporation Full authority fly-by-wire pedal system
FR2952447B1 (fr) * 2009-11-06 2012-08-17 Ratier Figeac Soc Dispositif de controle electronique de fonctionnement d'un organe de pilotage a surveillance croisee, dispositif de pilotage et aeronef
US9764830B2 (en) * 2012-02-10 2017-09-19 Bell Helicopter Textron Inc. Pilot control system with adjustable pedals
FR2995873B1 (fr) * 2012-09-27 2014-10-03 Airbus Operations Sas Procede et dispositif d'alerte contre un actionnement inapproprie d'un palonnier d'un aeronef.
CN103473967B (zh) * 2013-08-29 2015-12-02 南京航空航天大学 具有操纵力感的飞机模拟操纵装置
FR3029502B1 (fr) * 2014-12-08 2018-03-23 Airbus Operations Palonnier pour un aeronef
US10106245B2 (en) * 2015-10-19 2018-10-23 Honeywell International Inc. Automatic flight control actuator systems

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3999726A (en) * 1975-05-29 1976-12-28 Textron, Inc. Helicopter pylon-fuselage coupling for expanded CG range
DE3825809A1 (de) * 1988-07-29 1990-02-08 Messerschmitt Boelkow Blohm Steuereinrichtung mit einem steuerknueppel
US20110108674A1 (en) * 2009-11-10 2011-05-12 Cessna Aircraft Company Cockpit Rudder Control Mechanism For An Aircraft
CN103057697A (zh) * 2013-01-07 2013-04-24 中国商用飞机有限责任公司 一种用于飞机的方向舵脚蹬控制装置和控制方法
CN103587679A (zh) * 2013-11-28 2014-02-19 江西洪都航空工业集团有限责任公司 机械式飞机脚蹬调节机构

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See also references of EP3456626A4 *

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