All-electric flight control actuation system of unmanned aerial vehicle

Abstract

The invention belongs to the technical field of aviation unmanned aerial vehicle flight control, and discloses an unmanned aerial vehicle all-electric flight control actuation system which comprises a plurality of dual-redundancy actuator controllers and a plurality of electrical dual-redundancy electromechanical actuators. And the actuator controller receives a flight management control instruction, solves an electromechanical actuator control instruction through a main and standby channel dissimilar control algorithm, and drives the electromechanical actuator to drive a control surface to deflect, so that aircraft attitude control is realized. The safety problem of an existing unmanned aerial vehicle flight control actuation system is solved, it can be ensured that when control of any control surface fails, the remaining control surfaces can be driven to have the aircraft attitude compensation capacity, and therefore high reliability and high safety are achieved.

Description

Technical Field

[0001] The present invention belongs to the technical field of unmanned aircraft design, and in particular relates to a fully electric flight control actuation system for an unmanned aircraft. Background Art

[0002] With the rapid development of medium and large UAVs, high reliability and high safety have gradually become their important design goals while pursuing economic costs. The flight control actuation system, as a key system for controlling and driving the rudders of UAVs, must have high reliability and high safety design features to ensure the safe and reliable flight of UAVs.

[0003] Most existing UAV flight control actuation systems adopt a single-redundancy system architecture, resulting in a low level of system safety. A few UAV flight control actuation systems adopt a dual-redundancy system architecture, but the similar control strategies for the main and backup channels also lead to safety risks in the system.

[0004] In summary, the present application proposes a highly reliable dual-redundant heterogeneous all-electric flight control actuation system, which will effectively solve the above-mentioned problems currently existing. Summary of the invention

[0005] In view of the problems of the background technology, the present invention realizes a highly reliable and safe UAV flight control actuation system through system redundancy architecture configuration, system internal bus network architecture, system main and standby channel sensor heterogeneity, system main and standby channel non-similar control algorithm, and equipment dual redundancy design.

[0006] The technical solution of the present invention is achieved in this way:

[0007] A fully electric flight control actuation system for an unmanned aerial vehicle comprises: a plurality of dual-redundancy actuator controllers and a plurality of electrical dual-redundancy electromechanical actuators, wherein the dual-redundancy actuator controllers receive flight control instructions, solve the electrical dual-redundancy electromechanical actuator control instructions through a primary-standby channel non-similar control algorithm, drive the electrical dual-redundancy electromechanical actuators to drive the rudder deflection, and realize aircraft attitude control; the flight control actuation system is interconnected with an external flight control computer through a dual-redundancy RS422 bus; a redundant CAN bus network is built between the plurality of dual-redundancy actuator controllers and between the primary-standby channels of a single dual-redundancy actuator controller to realize high-reliability interaction of the system; the flight control actuation system adopts a primary-standby channel heterogeneous sensor configuration to realize the primary-standby channel adopting a non-similar control strategy, thereby ensuring high efficiency and high-precision control during normal operation, and simple and reliable control during failure; the flight control actuation system adopts a redundancy switching management logic of full switching, partial switching, and full removal of the primary-standby channel to ensure reliable reconstruction after a system failure.

[0008] As a further solution of the present invention: it includes two dual-redundancy actuator controllers and eight electrical dual-redundancy electromechanical actuators; one dual-redundancy actuator controller controls the left aileron, right rudder, right elevator, and left flap electromechanical actuators, and the other dual-redundancy actuator controller controls the right aileron, left rudder, left elevator, and right flap electromechanical actuators, so as to realize cross-control of various rudder surfaces and ensure that when a single-axis single rudder surface fails, the remaining rudder surfaces have compensation capability.

[0009] As a further solution of the present invention: the dual-redundancy actuator controller includes a primary and standby dual-redundancy control drive module; the electrical dual-redundancy electromechanical actuator includes a dual-winding motor, a primary chord variable standby Hall motor position sensor, and a primary and standby LVDT sensor.

[0010] As a further solution of the present invention: a dual redundant RS422 bus is set between the two dual redundant actuator controllers and the flight tube computer to ensure that when the communication between the flight control actuation system and the flight tube computer fails to leave only one RS422, the flight control actuation system can still work normally and feedback its own status information.

[0011] The flight control actuation system adopts a redundancy switching management logic of full switching, partial switching, and full removal of the main and standby channels to ensure reliable reconstruction after a system failure.

[0012] As a further solution of the present invention: the flight control actuation system adopts a main-backup working mode, the main channel works in normal state, and the backup channel is hot-backed up; when the main fault occurs, the backup channel control is automatically switched to; full switching, partial switching or full removal is performed according to different fault modes; when the main channel control chip and peripheral circuits, power supply module, etc. fail, the four electrical dual-redundancy electromechanical actuators controlled by the main channel of the dual-redundancy actuator controller are all switched to the backup channel control; when the driving circuit of a single electrical dual-redundancy electromechanical actuator fails or the electrical channel of the electrical dual-redundancy electromechanical actuator fails, the dual-redundancy actuator controller only controls the control channel of the electrical dual-redundancy electromechanical actuator to switch to the backup channel, and the remaining electrical dual-redundancy electromechanical actuators are still controlled by the main channel; when the electrical dual-redundancy electromechanical actuator has a mechanical jam or breakage fault, the dual-redundancy actuator controller completely cuts off the control of the electrical dual-redundancy electromechanical actuator.

[0013] As a further solution of the present invention: the main and backup channels of the flight control actuation system adopt a non-similar servo closed-loop control strategy, wherein the main channel adopts a space vector control algorithm and the backup channel adopts a simple square wave control to reduce the common mode risk of the main and backup channels and improve system safety.

[0014] As a further solution of the present invention: the electrical dual-redundancy electromechanical actuator includes a control chip and peripheral circuits, a power conversion module and a communication module.

[0015] The beneficial effects of this application are:

[0016] 1. A dual-redundant multi-control surface cross-control architecture of "two dual-redundant actuator controllers + eight electrical dual-redundant electromechanical actuators" is proposed, which can ensure that when the system fails once, it still has the ability to drive and control each control surface; it can also ensure that when any control surface fails, the remaining control surfaces can be driven to have the ability to compensate for the aircraft attitude.

[0017] 2. A redundant bus architecture network is proposed to interconnect the system with external flight control and management computers through dual redundant RS422, and a dual redundant CAN bus network is built inside the system to achieve highly reliable interaction of signal data inside and outside the system, ensuring that when the system fails to communicate with the outside and only one RS422 is left, the flight control actuation still has the ability to fully drive each control surface.

[0018] 3. Based on heterogeneous configurations such as main and backup channel sensors, non-similar control of the system's main and backup channels is achieved. The main channel adopts the space vector control law algorithm, and the backup channel adopts simple square wave control, which can effectively avoid the common mode problem of the main and backup channel control.

[0019] 4. Propose a redundancy management strategy based on system architecture and failure mode. Through fault handling measures such as full switching, partial switching and complete removal, ensure efficient reconstruction under system failure conditions and improve system availability under failure conditions.

[0020] 5. In summary, the dual-redundant heterogeneous all-electric flight control actuation system proposed in this application can greatly improve the reliability and fault reconstruction capability of the flight control actuation system, and can provide a high-safety flight control actuation implementation path for large and medium-sized UAVs.

[0021] The present application is further described in detail below in conjunction with the accompanying drawings of the embodiments. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Figure 1 A system architecture diagram of a preferred embodiment of the flight control actuation system of the present application;

[0023] Figure 2 This is a CAN bus network diagram of a preferred embodiment of the present application;

[0024] Figure 3 A schematic diagram of a control algorithm block diagram of the main and backup channels of a preferred embodiment of the present application;

[0025] Figure 4 A schematic diagram of the switching of the primary and backup channels according to a preferred embodiment of the present application. DETAILED DESCRIPTION

[0026] The technical solution of the present application will be described clearly and completely below in conjunction with the accompanying drawings of the embodiments.

[0027] This application provides a fully electric flight control actuation system for a UAV, such as Figure 1 As shown, it mainly includes two dual-redundant actuator controllers and eight electrical dual-redundant electromechanical actuators; one actuator controller controls the left aileron, right rudder, right elevator, and left flap electromechanical actuators, and the other actuator controller controls the right aileron, left rudder, left elevator, and right flap electromechanical actuators to achieve cross-control of various rudder surfaces and ensure that when a single-axis single rudder surface fails, the remaining rudder surfaces have compensation capability.

[0028] like Figure 1 As shown, the flight control actuation system adopts a main-standby dual-redundancy configuration. Under normal conditions, the main channel works and the standby channel is in hot standby. When the main channel fails, it automatically switches to the standby channel control to achieve controllability of the rudder surface under one system failure (except mechanical failure).

[0029] Such as 1. Figure 2 As shown, a redundant CAN bus network is set up between the two dual-redundancy actuator controllers of the flight control actuation system and between the main and standby channels of a single dual-redundancy actuator controller; a dual-redundancy RS422 bus is set up between the two dual-redundancy actuator controllers and the flight tube computer to ensure that when the communication between the flight control actuation system and the flight tube computer fails to leave only one RS422, the flight control actuation system can still work normally and feedback its own status information.

[0030] like Figure 1 , Figure 3 As shown, the flight control actuation system is based on the heterogeneity of the main and backup channel motor position sensors to support the main channel using a space vector control algorithm and the backup channel using a simple square wave control to implement different servo closed-loop control strategies, so as to reduce the common mode risk of the main and backup channels and improve system safety.

[0031] like Figure 4 As shown, in the normal state, the main channel of the flight control actuation system works and the standby channel is hot-backed up; when the main channel fails, the main and standby channels are switched redundantly, and full switching, partial switching or full removal is performed according to different failure modes; when the main channel control chip and peripheral circuits, power supply module, etc. fail, the four electrical dual-redundant electromechanical actuators controlled by the main channel of the dual-redundant actuator controller are all switched to the standby channel control; when the drive circuit of a single electrical dual-redundant electromechanical actuator fails or the electrical channel 1 of the electrical dual-redundant electromechanical actuator fails, the dual-redundant actuator controller only controls the control channel of the electrical dual-redundant electromechanical actuator to switch to the standby channel, and the remaining electrical dual-redundant electromechanical actuators are still controlled by the main channel; when the electrical dual-redundant electromechanical actuator has mechanical jamming or fracture, etc., the dual-redundant actuator controller completely removes the control of the electrical dual-redundant electromechanical actuator.

[0032] So far, the purpose of the present invention has been achieved.

[0033] The above are only preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the protection scope of the present invention.

Claims

1. A fully electric flight control actuation system for a UAV, characterized in that: include: A plurality of dual-redundancy actuator controllers and a plurality of electrical dual-redundancy electromechanical actuators, wherein the dual-redundancy actuator controllers receive flight control instructions, solve the electrical dual-redundancy electromechanical actuator control instructions through the primary and standby channel non-similar control algorithm, drive the electrical dual-redundancy electromechanical actuators to drive the rudder deflection, and realize aircraft attitude control; the flight control actuation system is interconnected with the external flight control computer through a dual-redundancy RS422 bus; a redundant CAN bus network is built between the plurality of dual-redundancy actuator controllers and between the primary and standby channels of a single dual-redundancy actuator controller to realize high-reliability interaction of the system; the flight control actuation system adopts a primary and standby channel heterogeneous sensor configuration to realize the primary and standby channels adopting a non-similar control strategy, ensuring high efficiency and high-precision control during normal operation, and simple and reliable control during failure.

2. The UAV all-electric flight control actuation system according to claim 1, characterized in that: It includes two dual-redundant actuator controllers and eight electrical dual-redundant electromechanical actuators; one of the dual-redundant actuator controllers controls the left aileron, right rudder, right elevator, and left flap electromechanical actuators, and the other dual-redundant actuator controller controls the right aileron, left rudder, left elevator, and right flap electromechanical actuators, to achieve cross-control of each rudder surface and ensure that when a single-axis single rudder surface fails, the remaining rudder surfaces have compensation capability.

3. The UAV all-electric flight control actuation system according to claim 2, characterized in that: The dual-redundancy actuator controller includes a primary and standby dual-redundancy control drive module; the electrical dual-redundancy electromechanical actuator includes a dual-winding motor, a primary chord variable standby Hall motor position sensor, and a primary and standby LVDT sensor.

4. The UAV all-electric flight control actuation system according to claim 3, characterized in that: A dual redundant RS422 bus is set between the two dual redundant actuator controllers and the flight tube computer to ensure that when the communication between the flight control actuation system and the flight tube computer fails to only one RS422 line, the flight control actuation system can still work normally and feedback its own status information.

5. The UAV all-electric flight control actuation system according to claim 4, characterized in that: The flight control actuation system adopts a redundancy switching management logic of full switching, partial switching, and full removal of the main and standby channels to ensure reliable reconstruction after a system failure.

6. The UAV all-electric flight control actuation system according to claim 5, characterized in that: The flight control actuation system adopts a master-slave working mode, in which the main channel works in normal state and the backup channel is hot-backed up; when the main fault occurs, the backup channel control is automatically switched to; full switching, partial switching or full removal is performed according to different fault modes; when the main channel control chip and peripheral circuits, power supply module, etc. fail, the four electrical dual-redundancy electromechanical actuators controlled by the main channel of the dual-redundancy actuator controller are all switched to the backup channel control; when the driving circuit of a single electrical dual-redundancy electromechanical actuator fails or the electrical channel of the electrical dual-redundancy electromechanical actuator fails, the dual-redundancy actuator controller only controls the control channel of the electrical dual-redundancy electromechanical actuator to switch to the backup channel, and the remaining electrical dual-redundancy electromechanical actuators are still controlled by the main channel; when the electrical dual-redundancy electromechanical actuator has mechanical jamming or fracture, etc., the dual-redundancy actuator controller completely cuts off the control of the electrical dual-redundancy electromechanical actuator.

7. The UAV all-electric flight control actuation system according to claim 6, characterized in that: The main and standby channels of the flight control actuation system adopt a non-similar servo closed-loop control strategy, wherein the main channel adopts a space vector control algorithm and the standby channel adopts a simple square wave control to reduce the common mode risk of the main and standby channels and improve system safety.

8. The UAV all-electric flight control actuation system according to claim 7, characterized in that: The electrical dual-redundancy electromechanical actuator comprises a control chip and peripheral circuits, a power conversion module and a communication module.

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