Universal control system for a flight trimmer with an active primary channel and a passive backup channel

Abstract

This universal control system (1) for a flight trim system of an aircraft (2) comprises a control element (3) pivoting about at least one axis of rotation (5), and a primary force feedback means (9) comprising an electric motor (11) and connected to the control element (3), a backup force feedback means (25) connected to the control element (3) via a clutch (27), and a control unit (37) for the primary force feedback means (9). Figure for the abstract: Fig. 1

Description

Description Title of the invention: Universal control system for a flight compensator with an active primary channel and a passive backup channel technical field

[0001] — The present invention relates, in general, to control systems of flight for aircraft and relates more specifically to a flight trimmer for aircraft, especially for helicopters. Previous techniques

[0002] = Flight trimmers, or trims, have the particular function of compensating for the drifts linked to disturbances likely to influence the parameters of aircraft flight, without the pilot having to act on the flight controls.

[0003] — In addition, flight trimmers conventionally restore to the pilot the aircraft a resisting force on the flight controls in the form of a return of effort. This effort is generally passive, either linear (spring trims) or constant (friction trims), and can also be active. The resisting force is also called reminder, reminder force, or return of effort.

[0004] — Conventional flight trimmer architectures generally incorporate a primary passive force return path, for example spring-loaded or friction-based. These architectures are reliable but not configurable, especially once the spring stiffness or defined friction. These architectures therefore do not allow the command to fully adapt to flight conditions and the pilot's wishes the aircraft.

[0005] = Optionally, active secondary paths can be used to assist the return effort exerted by the main path, for example using a motor, in such a way to allow for slight adaptability of the flight compensators.

[0006] — Moreover, each command, of the cyclic axis, of the collective axis, or of the yaw, is seen generally assign a different flight trimmer architecture in order to solve problems specific to each command, for example, a trim to Friction is preferred for controlling the collective axis, while a spring-loaded trim is preferred for controlling the cyclic axis.

[0007] — Consequently, there is no architecture that is entirely configurable and meeting the safety and ergonomic requirements of the controls. Furthermore, the architects Existing trim structures are not universal; they only correspond to one cyclic, collective, or lace control, and therefore require development complex. Description of the invention The present invention therefore aims to overcome the aforementioned drawbacks and to provide a trim control system capable of generating a freely parameterizable and universally usable force law. The present invention relates to a universal control system for an aircraft flight compensator, comprising a piloting element pivoting about at least one axis of rotation, the system further comprising a main force feedback means comprising an electric motor and linked to the piloting element, a backup force feedback means linked to the piloting element via a clutch, and a control unit for the main force feedback means. Thus, this architecture can be used for any trim control, such as cyclic, collective, or yaw control. Furthermore, the primary control is active, allowing for free and real-time parameterization of the force feedback law. Finally, system safety is ensured by the backup control. Advantageously, the main means of force feedback includes a torque sensor between the electric motor and the control unit and / or includes a control loop for the electric motor supply current. In one embodiment, the main means of force feedback includes a position and / or speed sensor of the electric motor. In a particular embodiment, the control unit and the electric motor are configured to simulate a force law of a spring and / or a force law of a friction module. Advantageously, the backup means of force return includes a dry friction module and / or a controllable brake with current failure. In a particular embodiment, the backup force return means includes a damping motor. Advantageously, the control unit includes a control module for the main force feedback means, and a monitoring module for the main force feedback means. Advantageously, the main force return means is linked to the pilot member via a clutch and / or a safety pin, and in which the monitoring module is configured to, in the event of failure of the main force return means, decouple the main force return means from the pilot member and to couple the backup force return means to the pilot member. Advantageously, the control system includes a main shaft connecting the pilot member to the main force feedback means and / or the backup force feedback means, and at least one angular position sensor for the main shaft. The invention also relates to a helicopter comprising three systems as defined above, respectively for the control of the collective axis, the axis cyclic, and yaw, and an onboard computer controlling each control unit. Brief description of the drawings Other objects, features and advantages of the invention will become apparent from the following description, given solely by way of non-limiting example, and made with reference to the accompanying drawings in which: [Fig.1] is a schematic representation of a control system according to the invention. Detailed description of at least one embodiment A universal control system 1 of a flight trimmer of an aircraft 2 according to the invention has been schematically represented in [Fig.1]. The aircraft is for example a helicopter comprising said system 1 controlling the collective axis, the cyclic axis or the yaw of said helicopter. Universal means that the control system is suitable for controlling the collective axis, the cyclic axis, or the yaw. The control system 1 includes a frame (not shown) and a control element 3 mounted pivotally relative to the frame around at least one axis of rotation 5. The control element 3 is for example a control stick 3 and allows for example to control the collective axis, the cyclic axis or the yaw of the aircraft 2. The frame includes mechanical stops 7 against which the piloting element 3 butts to limit its angular stroke. The control system includes at least one main force return means 9 linked to the control member 3, the restoring force of the main force return means 9 being directed in the opposite direction to an external force applied by the aircraft pilot on the control member 3. The main means 9 comprises an electric motor 11 and is connected to the control unit 3 via a main shaft 13 of the control system 1. The electric motor 11 is for example a three-phase synchronous permanent magnet motor, and / or a brushed motor, and / or a piezoelectric motor. Optionally, the main force feedback means 9 includes a position and / or speed sensor 15 configured to measure the position and / or speed of the electric motor 11. Optionally, the main means 9 of force feedback includes a torque sensor 17 positioned between the electric motor 11 and the main shaft 13 and enabling monitoring of the pilot's force on the piloting element. The main means 9 is a purely active and adaptive means of force feedback. In particular, it does not include a spring or friction module. The presence of the electric motor 11 allows the main means 9 to be fully and freely parameterizable in real time in order to generate a force feedback law adapted to each use, and to each phase of flight of the aircraft 2. This force feedback law is comparable to a complex mechanical force feedback law, more advanced than a force law obtained using a purely passive or active system including an analog controller. The main means 9 further includes a gear system 19, for example positioned between the electric motor 11 and the torque sensor 17. In order to be able to be decoupled from the control member 3, the main means 9 includes a clutch 21 and / or a safety pin 23. Thus, in the event of a failure in the main means 9 of force return, it is possible to disengage the clutch 21 and / or break the safety pin 23 in order to separate the main means 9 from the safety member 3. The control system also includes a backup force feedback means 25 connected to the control element 3, for example, via the main shaft 13. The backup means 25 is a passive force feedback means and therefore ensures the safety of the system 1 only in the event of a failure of the active main means 9. In normal operation, the backup means 25 does not influence the force feedback on the control element 3. This backup means 25 thus contributes to the free parameterization of the force feedback by the main force feedback means 9. The backup means 25 optionally includes a clutch 27 allowing the backup means 25 to be decoupled from the control unit 3 in nominal mode and to be coupled in the event of a failure, i.e. when the main means is unusable and / or disconnected. The backup means 25 further includes a dry friction module 29, the friction generating a feedback force at the control unit. Alternatively, the backup means 25 includes a fail-safe controllable brake (not shown). This brake is unique in that it only activates when it is no longer powered, for example during a power failure, thus allowing it to generate and transmit a force between the friction module 29 and the main shaft 13 even without the main means being powered. The controllable brake includes a friction module 29 associated with a clutch 27 controlled in an "on or off" manner. The fact that it is under power failure depends on the type of clutch used, here normally closed. Which allows, in case of lack of power to the clutch, the transmission of forces between the friction module (29) and the output shaft of the actuator (13). Optionally, the backup means 25 includes a damping motor 31, particularly useful when a friction module 29 is present in order to smooth out torque fluctuations due to friction. In addition, the backup means 25 includes a gear 33 and, optionally, a safety pin 35. The gear 33 may include a reduction gearbox to adapt the forces supplied. The system also includes a control unit 37 of the main means 9. This control is digital and is carried out in closed loop and / or open loop. Optionally, the control unit 37 also controls the backup means 25, in particular the clutch 27. The control unit 37 includes a control module 39 for the main means 9 of force feedback, as well as a monitoring module 41 for the main means 9. The control module 39 performs closed-loop force law calculations using certain information provided by the torque sensor 17. In particular, the monitoring module 41 is configured to decouple the main force feedback means 9 and, if necessary, to couple the backup force feedback means 25 to the control unit 3. The monitoring means 41 sends a command to clutches 21 and 27 for this purpose, for example, in the event of a failure of the main force feedback means 9. Clutches 21 and 27 are complementary; one is engaged while the other is disengaged. The control unit 37 and the electric motor 11 are configured to simulate the force law of a spring and / or the force law of a friction module. Any force law of a different form can also be generated by the electric motor 11 under the command of the pilot or an onboard computer. System 1 optionally includes one or more angular position sensors 43 of the main shaft 13. These sensors communicate with the control unit 37. A sensor 43 is for example of the type rotary variable differential transformer also called RVDT sensor, or of the resolver type, hall effect sensor, potentiometer or any other type of sensor. Optionally, each angular position sensor 43 is connected to the main shaft 13 by a safety pin 45, allowing the sensor 43 to be disconnected from the main shaft 13, for example in the event of a failure of the sensor 43. The control unit 37 receives information from the position sensor 15, the torque sensor 17 and the angular position sensor 43, as well as from an on-board computer 46. Using this information, the control module 39 controls the electric motor 11 by current and the monitoring module 41 controls the clutch 21 and the clutch 27 by current. In one particular embodiment, system 1 does not include a torque sensor 17, as this could be considered bulky and expensive, and the control unit can put An open-loop control system for the electric motor 11's supply current is implemented, with static and dynamic compensation to overcome static and viscous friction, respectively. The supply current control loop provides a representation of the applied force. During its use, the control unit 37 implements a first position control loop of the piloting element 3. The control unit 37 can integrate, if necessary, into this first loop the current control loop, used for the control of the motor 11. The first control loop will be implemented by providing access to a parameter table included in the control unit 37 and which contains all the singular points of the effort law profile that can be realized. Therefore, it will be possible to freely modify the effort law of system 1 within the limits of its range of validity. The control unit 37 also implements a protection to provide a smooth torque transition between a current setpoint and a new setpoint when a setpoint is changed by a pilot or the on-board computer. An alert will be raised by control unit 37 if a risk of inappropriate configuration is detected, for example near the mechanical stops, in order to guarantee the mechanical integrity of system 1. The invention also relates to a helicopter 2 comprising three systems 1, respectively for controlling the collective axis, the cyclic axis, and the yaw of said helicopter 2. The helicopter also includes an onboard computer 46 controlling each control unit 37 by communicating position, and / or speed, and / or thrust law commands. Optionally, a single control unit 37 can be used for all three systems 1. Optionally, some sensors such as the angular position sensor 43 communicate with the on-board computer 46.

Claims

Claims

1. System (1) for universal control of a flight compensator of a aircraft (2), comprising a piloting member (3) pivoting around at least one at least one axis of rotation (5), characterized in that it comprises means main (9) force feedback comprising an electric motor (11) and linked to the control organ (3), a backup means (25) of return of effort linked to the control organ (3) via a clutch (27), and a control unit (37) of the main means (9) of return of effort.

2. System (1) according to claim |, wherein the main means (9) force feedback includes a torque sensor (17) between the motor electric (11) and the control member (3) and / or comprises a loop of control of the supply current of the electric motor (11).

3. System (1) according to one of claims 1 and 2, in which the means main force feedback (9) includes a position sensor (15) and / or speed of the electric motor (11).

4. System (1) according to any one of claims 1 to 3, in which the control unit (37) and the electric motor (11) are configured to simulate a spring force law and / or a law of effort of a friction module.

5. System (1) according to any one of claims 1 to 4, in which the emergency force feedback means (25) comprises a module dry friction (29) and / or a brake that can be controlled by a lack of current.

6. System (1) according to any one of claims 1 to 5, in in which the emergency force feedback means (25) comprises a motor depreciation (31).

7. System (1) according to any one of claims 1 to 6, in which the control unit (37) comprises a control module (39) of the main means (9) of force feedback, and a monitoring module (41) of the main means (9) of force feedback.

8. System (1) according to any one of claims 1 to 7, in which the main means (9) of force feedback is linked to the organ of control (3) by means of a clutch (21) and / or a safety pin (23), and in which the monitoring module (41) is configured for, in the event of failure of the main means (9) of return of effort, decouple the main means (9) of force feedback from the organ steering (3) and to couple the emergency means (25) back of effort to the steering organ (3).

9. System (1) according to any one of claims 1 to 8, comprising a main shaft (13) connecting the control member (3) to the main means (9) of force feedback and / or emergency means (25) of force feedback, and at least one angular position sensor (43) of the main shaft (13).

10. Helicopter (2) characterized in that it comprises three systems (1) according to any one of claims 1 to 9, respectively for the collective axis, cyclic axis, and yaw control, and an or- on-board diner (46) controlling each control unit (37).

Images