Together for an aircraft

Abstract

The present description relates to an assembly for an aircraft comprising: a compensator (5) comprising: a chassis (50) intended to be fixedly mounted on an aircraft cell; and an actuator (52) movably mounted on the chassis (50); a mechanical transmission system (40, 42, 44, 46) configured to transmit a force on the one hand between a flight control (30, 32, 34) of the aircraft and a control surface (20, 22) of the aircraft, and on the other hand between the actuator (52) and the control surface (20, 22), the mechanical transmission system (40, 42, 44, 46) being without a hydraulic servomotor; and a control system (6) configured to drive the actuator (52) according to a force exerted on the flight control (30, 32, 34) so ​​as to control a movement of the rudder (20, 22) relative to the airframe (10, 14). Figure for the abbreviation: Fig. 2

Description

Title of the invention: Assembly for an aircraft Technical field

[0001] This disclosure relates to the aeronautical field. More specifically, this disclosure relates to the control of the control surfaces of an aircraft such as a helicopter. STATE OF THE ART

[0002] A pilot can control the movement of a helicopter in space using control surfaces mechanically connected to the cockpit controls. Some helicopters are, in addition, equipped with hydraulic servomotors providing the interface between the controls and the control surfaces in order to reduce the effort exerted by the pilot on the controls to control the movement of the helicopter in space. However, the presence of hydraulic servomotors makes the helicopter heavier, which reduces its flight autonomy. To overcome this disadvantage, in some helicopters, the hydraulic servomotors are replaced by electric actuators, which are lighter. However, the presence of electric actuators to control the control surfaces requires the integration of power electronics into the system connecting the controls to the control surfaces, which can be complex and expensive. GENERAL STATEMENT

[0003] An aim of the present disclosure is to increase the flight autonomy of an aircraft in a reliable, simple and inexpensive manner.

[0004] For this purpose, according to one aspect of the present disclosure, there is proposed an assembly for an aircraft comprising: a compensator comprising: a frame intended to be fixedly mounted on an airframe of the aircraft; and an actuator movably mounted on the frame; a mechanical transmission system configured to transmit a force on the one hand between a flight control of the aircraft and a control surface of the aircraft, and on the other hand between the actuator and the control surface, the mechanical transmission system being devoid of a hydraulic servomotor; and a control system configured to drive the actuator as a function of a force exerted on the flight control so as to control a movement of the control surface relative to the cell.

[0005] Advantageously, but optionally, the assembly may comprise at least one of the following characteristics, taken alone or in any combination: - it further comprises a force sensor configured to measure an intensity of the force, the control system being further configured to control the actuator in function of a measurement of the intensity of the effort carried out by the effort sensor; - the control system is configured to drive the actuator so that: if the measurement of the intensity of the effort is less than an intensity threshold, the control system is designed to maintain the actuator in position; and if the measurement of the intensity of the force is greater than the intensity threshold, the control system is designed to control a speed of movement of the actuator relative to the chassis as a function of the intensity of the force; - it further comprises a switch, in which the aircraft is configured so that the force causes the switch to switch from a first state to a second state, the control system being further configured to control the actuator according to a switching duration during which the switch is switched to the second state; and - the control system is configured to drive the actuator so that: if the switching duration is less than a duration threshold, the control system is designed to maintain the actuator in position; and if the switching time is greater than the duration threshold, the control system is designed to control a speed of movement of the actuator relative to the chassis depending on the switching time.

[0006] According to another aspect of the present disclosure, there is provided an aircraft comprising: an airframe; a control surface mounted movably on the cell; a flight control mounted movably on the airframe; and an assembly as previously described, in which the chassis is fixedly mounted on the cell.

[0007] Advantageously, but optionally, the aircraft may comprise at least one of the following characteristics, taken alone or in any combination: - the rudder is one of: a yaw rudder, a pitch rudder and a roll rudder; - the flight control is one of: a stick, a rudder pedal and a collective; and - the aircraft is a helicopter.

[0008] According to another aspect of the present disclosure, there is provided a method of controlling an aircraft comprising: a cell; a control surface mounted movably on the cell; a flight control mounted movably on the cell; a compensator comprising: a chassis mounted fixedly on the cell; and an actuator mounted movably on the chassis; a mechanical transmission system configured to transmit a force on the one hand between the flight control and the control surface, and on the other hand between the actuator and the control surface, the mechanical transmission system being devoid of a hydraulic servomotor; the control method comprising piloting the actuator as a function of a force exerted on the flight control so as to control a movement of the control surface relative to the cell. DESCRIPTION OF FIGURES

[0009] Other characteristics, aims and advantages will emerge from the following description, which is purely illustrative and non-limiting, and which must be read in conjunction with the appended drawings in which:

[0010] [Fig.l] schematically illustrates an aircraft.

[0011] [Fig.2] schematically illustrates a part of the aircraft.

[0012] [Fig.3] is a flowchart illustrating an implementation of aircraft control.

[0013] Throughout the figures, similar elements bear identical references. DETAILED DESCRIPTION

[0014] [Fig.l] illustrates an aircraft 1 which is a device capable of rising or circulating in the air.

[0015] The aircraft 1 extends along a longitudinal axis XX and comprises a cell 10, 14, which constitutes the structural part of the aircraft 1, and a wing 12.

[0016] The cell 10, 14 comprises a fuselage 10 and a tailplane 14. The fuselage 10 is the spindle-shaped frame which constitutes the body of the aircraft 1 and whose shape is intended to facilitate the penetration of the aircraft 1 into the air. Typically, the cockpit, intended to accommodate the pilot, is arranged within the fuselage 10. The tailplane 14 is notably configured to ensure the stability of the aircraft 1.

[0017] The wing 12 is in particular configured to provide lift for the aircraft 1.

[0018] The aircraft 1 further comprises at least one control surface 20, 22, preferably a plurality of control surfaces 20, 22, each mounted movably relative to the cell 10, 14 and configured to produce and / or control a movement of the aircraft 1 around its center of gravity along at least one of three axes: the pitch axis YY, the roll axis XX and the yaw axis ZZ. These three axes XX, YY, ZZ form a fixed reference point connected to the center of gravity of aircraft 1. When aircraft 1 is stationary on the ground, the yaw axis ZZ coincides with the vertical direction, orthogonal to the ground, while the pitch axis YY and the roll axis XX belong to a plane parallel to the ground, the roll axis XX being parallel to the longitudinal axis XX and the pitch axis YY being orthogonal to the longitudinal axis XX.

[0019] Furthermore, the aircraft 1 comprises at least one flight control 30, 32, 34, preferably a plurality of flight controls 30, 32, 34, each positioned within the cockpit so as to be accessible to the pilot. The movements of the aircraft 1 are controlled by the pilot by means of the flight controls 30, 32, 34. In this regard, the aircraft 1 comprises a transmission system 40, 42, 44, 46, 48 in particular configured to transmit a force between the flight controls 30, 32, 34 and the control surfaces 20, 22, in order to control a movement of the control surface 20, 22 relative to the cell 10, 14 in order to produce and / or control a movement of the aircraft 1 around its center of gravity. The transmission system 40, 42, 44, 46, 48 may be mechanical, hydraulic and / or electrical, that is to say comprise a plurality of links connecting the flight controls 30, 32, 34 to the control surfaces 20, 22, the links being mechanical, hydraulic and / or electrical.

[0020] Furthermore, the aircraft 1 comprises at least one compensator 5, preferably a plurality of compensators 5. Each compensator 5 comprises a chassis 50 fixedly mounted on the airframe 10, 14 and an actuator 52 mounted movably on the chassis 50, for example in rotation or in translation relative to the chassis 50. The chassis 50 may take the form of a housing fixed to the floor of the cockpit and the actuator 52 may be an arm mounted for rotation about an axis extending from the chassis 50. The compensator 5 may further comprise at least one jack arranged within the chassis 50 and the movable part of which is connected to the actuator 52, so as to multiply the forces likely to be generated by the actuator 52 when it moves relative to the chassis 50. The jack and / or the actuator 52 may be actuated by means of an electric motor.Alternatively, or in addition, the trim tab 5 may comprise at least one, or even several, mechanical or magnetic return members, connecting the chassis 50 to the actuator 52 so as to control the movement and / or position thereof relative to the chassis 50. In addition, the trim tab 5 may comprise a clutch device making it possible to engage and disengage the actuator 52 from the members of the trim tab 5 provided for controlling its movement and / or its position. The transmission system 40, 42, 44, 46, 48 is also configured to transmit a force between the actuator 52 and at least one of the control surfaces 20, 22 in order to control a movement of the control surface 20, 22 relative to the cell 10, 14 in order to produce and / or control a movement of the aircraft 1 around its center of gravity.In this way, the trim tabs 5 can compensate, that is to say dampen, in whole or in part, the forces perceived by the pilot on the flight controls 30, 32, 34 linked to the aerodynamic forces to which the control surfaces 20, 22 are subjected. In other words, the trim tabs 5 offer the possibility of controlling the force feedback perceived by the pilot in the flight controls 30, 32, 34, and this in a precise manner. Finally, the trim tabs 5 can each maintain and / or place at least one of the control surfaces. 20, 22 in a position allowing the aircraft to be balanced so as to ensure automatic piloting of the aircraft 1 by maintaining the position imposed by the pilot around the center of gravity of the aircraft 1, whatever the speed and mass of the aircraft 1. In this regard, the frames 50 provide an anchoring point for the control surfaces 20, 22 on the cell 10, 14, via the transmission system 40, 42, 44, 46, 48.

[0021] The transmission system 40, 42, 44, 46, 48 may also comprise at least one linear actuator member 46, preferably a plurality of linear actuator members 46s, each mounted movably relative to the cell 10, 14. Each linear actuator member 46 is configured to cause low amplitude movements of the control surfaces 20, 22 relative to the cell 10, 14, and this at a high frequency, in order to improve the stability of the aircraft 1 in flight. In this regard, each linear actuator member 46 may comprise at least one electric motor. The linear actuator members 46s are designed to operate without their action being felt by the pilot in his manipulation of the flight controls 30, 32, 34.

[0022] Furthermore, the aircraft 1 comprises a control system 6, illustrated in [Fig. 2], configured to control at least one of the actuators 52 of the compensators 5, typically a position and / or a speed of a movement of the actuator 52 relative to the chassis 50. The control system 6 may typically comprise a computer (or processor) configured to receive information and to process it so as to transmit a command to the compensator 5, typically to the electric motor of the compensator 5, to control the actuator 52. If necessary, the control system 6 may also be configured to control at least one of the linear actuation members 46.

[0023] The aircraft 1 illustrated in [Fig.l] is a helicopter 1.

[0024] The wing 12 of the helicopter 1 is rotating, that is to say that it comprises at least one main rotor 20 mounted to rotate relative to the fuselage 10. In this respect, the main rotor 20 comprises a plurality of blades 200, each blade 200 having an aerodynamic shape and being capable of being controlled, in particular in incidence, to generate the lift necessary for the support of the helicopter 1 when the rotor is rotated. In addition, the position of the plane of rotation of the blades 200 can also be controlled to control a movement of the helicopter 1 around the pitch axis YY and the roll axis XX. Therefore, the main rotor 20 constitutes both the pitch control surface and the roll control surface of the helicopter 1.

[0025] The empennage 14 of the helicopter 1 comprises an anti-torque rotor 22, or tail rotor 22, which also comprises a plurality of blades 220 each having an aerodynamic shape and also being able to be controlled, in particular in incidence, to control the movements of the helicopter 1 around the yaw axis Z- Z, the axis of rotation of the tail rotor 22 being orthogonal to the axis of rotation of the main rotor 20. The tail rotor 22 therefore constitutes the yaw control surface of the helicopter 1.

[0026] The flight controls 30, 32, 34 of the helicopter 1 comprise a stick 30, which is designed to extend between the pilot's legs and makes it possible to control the movements of the helicopter 1 around the pitch axis YY and the roll axis XX, in particular by controlling the inclination of the plane of rotation of the blades 200 of the main rotor 20, a rudder pedal 32, preferably two rudder pedals 32, which make it possible to control the movements of the helicopter 1 around the yaw axis ZZ, in particular by controlling the rotation speed of the tail rotor 22 and / or the incidence of the blades 220 of the tail rotor 22, and a collective 34 which makes it possible to control the movements of the helicopter 1 along the yaw axis ZZ, i.e. the vertical elevation movements of the helicopter 1, in particular by controlling the speed of rotation of the main rotor 20 and / or of the incidence of the blades 200 of the main rotor 20.

[0027] [Fig.l] illustrates a mechanical transmission system 40, 42, 44, 46, 48 which comprises a set of linkages 40, cables 42 and horns 44 connected to each other, to the flight controls 30, 32, 34, to the trim tabs 5 and to the control surfaces 20, 22. In such a transmission system 40, 42, 44, 46, 48, the linear actuating members 46 are each arranged on one of the cables 42 so as to connect two free cable ends 42 to each other. Furthermore, the free end of the arm formed by the actuator 52 is connected to at least one of the cables 42.

[0028] [Fig. 1] further illustrates that the mechanical transmission system 40, 42, 44, 46, 48 may comprise at least one hydraulic servomotor 48, preferably a plurality of hydraulic servomotors 48, for example a hydraulic servomotor 48 connected to the tail rotor 22 and a plurality of hydraulic servomotors 48 connected to the main rotor 20. The hydraulic servomotors 48 are configured to multiply the force transmitted by the pilot to the flight controls 30, 32, 34 and / or the force generated by the actuator 52, which forces are intended to be transmitted to the main rotor 20 and to the tail rotor 22. In this way, the control of the control surfaces 20, 22 is facilitated, in particular in conditions where the aerodynamic forces to which the blades of the main rotor 20 and of the tail rotor 22 are subjected are high, which makes the control of their movement relative to the main rotor 20 and of the tail rotor 22 easier. to cell 10,14 particularly difficult.

[0029] The weight of the hydraulic servomotors 48 and the hydraulic power supply systems necessary for their operation reduces the flight autonomy of the helicopter 1. This is why it can be provided that the mechanical transmission system 40, 42, 44, 46, 48 is devoid of them.

[0030] Where appropriate, it is however necessary to ensure the function of the hydraulic servomotors 48, typically by means of electric power actuators, which have the advantage of no longer requiring the presence of hydraulic supply systems, but require the integration of power electronics controlled by the compensators 5 and / or the linear actuating members 46.

[0031] Another way of ensuring the function of the hydraulic servomotors 48 is to use at least one of the compensators 5 by adapting the piloting carried out by the control system 6 so as to allow the compensator 5 to generate greater forces to move and / or maintain the control surfaces 20, 22 in position relative to the cell 10, 14 without the haptic feeling of the pilot manipulating the flight controls 30, 32, 34 being modified.

[0032] In this regard, as illustrated in [Fig. 2], the control system 6 is configured to control the actuator 52 of the trim tab 5 as a function of a force exerted on the flight control 30, 32, 34 so as to control a movement of the control surface 20, 22 relative to the cell 10, 14. The force is exerted on the flight control 30, 32, 34 by the pilot. But, instead of being transmitted to at least one of the control surfaces 20, 22 by the mechanical transmission system 40, 42, 44, 46, 48, it is measured to control the trim tab 5 so that it is the forces exerted by the actuator 52 which are transmitted to the control surfaces 20, 22 by the mechanical transmission system 40, 42, 44, 46 in order to control them.In this way, in addition to its functions of regulating the force feedback perceived by the pilot and of automatic piloting to maintain the balance of the helicopter 1, the compensator 5 fulfills the function of assisting the manual piloting of the helicopter 1 which was previously fulfilled by the hydraulic servomotors 48.

[0033] To do this, it may be provided that the helicopter 1 comprises a force sensor 7 configured to measure an intensity of the force exerted on the flight control 30, 32, 34. The control system 6 is then configured to control the actuator 52 as a function of a measurement of the intensity of the force carried out by the sensor.

[0034] Alternatively, or in addition, it may be provided that the helicopter 1 comprises a switch 7. The helicopter 1 is then configured so that the force exerted on the flight control 30, 32, 34 causes the switch 7 to switch from a first state to a second state, typically the force exerted on the stick 30 causes a movement of the stick 30 which comes into contact with the switch 7 and causes it to switch. The control system 6 is, in this case, configured to control the actuator 52 as a function of a switching duration during which the switch 7 is switched to the second state.

[0035] [Fig. 3] illustrates more precisely the logic of piloting the actuator 52 of the compensator 5 which is implemented by the control system 6 as a function of the force exerted on the flight control 30, 32, 34, whether the piloting is implemented in depending on the intensity of the effort measured by the effort sensor 7 or on the switching duration during which the switch 7 is switched to the second state.

[0036] If the measurement of the intensity of the force is less than a first intensity threshold, or if the switching duration is less than a first duration threshold, the control system 6 is provided to maintain the actuator 52 in position. In other words, as long as the pilot does not express, by requesting the flight controls 30, 32, 34, his desire to take back control of the control surfaces 20, 22, the trim tab 5 maintains at least one of the control surfaces 20, 22 in position relative to the cell 10, 14, whether following a prior command received from the flight controls 30, 32, 34 or according to an automatic pilot logic.

[0037] If the measurement of the intensity of the force is greater than the first intensity threshold, or if the switching duration is greater than the first duration threshold, the control system 6 is provided to control a speed of a movement of the actuator 52 relative to the chassis 50 and / or a position of the actuator 52 relative to the chassis 50, as a function of the intensity of the force measured or, where appropriate, the switching duration. The position and / or the speed of the actuator 52 determines a force exerted by the actuator 52 transmitted by the mechanical transmission system 40, 42, 44, 46 to at least one of the control surfaces 20, 22. In other words, depending on the way in which the pilot requests the flight controls 30, 32, 34, the control system 6 controls the actuator 52 so that the force that the actuator 52 exerts on the at least one of the control surfaces 20, 22 is a force that corresponds to what the pilot wants by requesting the flight control 30, 32, 34.The compensator 5 therefore replaces the hydraulic servomotors 48 during manual piloting.

[0038] [Fig. 3] illustrates that a second intensity threshold and / or a second duration threshold, each respectively higher than the first intensity threshold and the first duration threshold, may be provided. In this case, the first intensity threshold, respectively the first duration threshold, may only constitute a detection threshold for the control system 6 which is then informed that the pilot wishes to switch from automatic piloting by the trim tabs 5 to manual piloting. But the piloting of the speed and / or the position of the actuator 52 is only implemented once the second intensity threshold, respectively the second duration threshold, has been exceeded.

[0039] Although the control logic of the trim tabs 5 has been described with reference to a helicopter 1 in the present description, this is of course not limiting. Indeed, such control logic of the trim tabs applies to any aircraft whose transmission system is mechanical and devoid of hydraulic servomotors to assist the control of the control surfaces.

[0040] Similarly, the control logic of the compensators 5 described in the present description is preferably implemented for the control of the yaw control surface 22, which undergoes much less aerodynamic forces than the pitch and roll control surfaces 20, especially within a helicopter 1.

Claims

Claims

1. Assembly for an aircraft (1) comprising: a compensator (5) comprising: a chassis (50) intended to be fixedly mounted on a cell (10, 14) of the aircraft (1); and an actuator (52) movably mounted on the chassis (50); a mechanical transmission system (40, 42, 44, 46) configured to transmit a force on the one hand between a flight control (30, 32, 34) of the aircraft (1) and a control surface (20, 22) of the aircraft (1), and on the other hand between the actuator (52) and the control surface (20, 22), the mechanical transmission system (40, 42, 44, 46) being devoid of a hydraulic servomotor; a switch (7) configured such that a force exerted on the flight control (30, 32, 34) causes the switch (7) to switch from a first state to a second state;and a control system (6) configured to control the actuator (52) as a function of a switching duration during which the switch (7) is switched into the second state so as to control a movement of the control surface (20, 22) relative to the cell (10, 14).;

2. Assembly according to claim 1, further comprising a force sensor (7) configured to measure an intensity of the force exerted on the flight control (30, 32, 34), the control system (6) being further configured to control the actuator (52) as a function of a measurement of the intensity of the force carried out by the force sensor (7) so as to control a movement of the control surface (20, 22) relative to the cell (10, 14).

3. Assembly according to claim 2, in which the control system (6) is configured to control the actuator (52) so that: if the measurement of the intensity of the force exerted on the flight control (30, 32, 34) is less than an intensity threshold, the control system (6) is provided to maintain the actuator (52) in position; and if the measurement of the intensity of the force exerted on the flight control (30, 32, 34) is greater than the intensity threshold, the control system (6) is provided to control a speed of a movement of the actuator (52) relative to the chassis (50) as a function of the intensity of the force exerted on the flight control (30, 32, 34).

4. An assembly according to any one of claims 1 to 3, wherein the control system (6) is configured to control the actuator (52) so that: if the switching duration is less than a duration threshold, the control system (6) is provided to maintain the actuator (52) in position; and if the switching duration is greater than the duration threshold, the control system (6) is provided to control a speed of a movement of the actuator (52) relative to the chassis (50) as a function of the switching duration.

5. An aircraft (1) comprising: an airframe (10, 14); a control surface (20, 22) movably mounted on the airframe (10, 14); a flight control (30, 32, 34) movably mounted on the airframe (10, 14); and an assembly according to any one of claims 1 to 4, wherein the frame (50) is fixedly mounted on the airframe (10, 14).

6. An aircraft (1) according to claim 5, wherein the control surface (20, 22) is one of: a yaw control surface (22), a pitch control surface (20) and a roll control surface (20).

7. An aircraft (1) according to any one of claims 5 and 6, wherein the flight control (30, 32, 34) is one of: a stick (30), a rudder pedal (32) and a collective (34).

8. An aircraft (1) according to any one of claims 5 to 7, wherein the aircraft (1) is a helicopter.

9. A method for controlling an aircraft (1) comprising: a cell (10, 14); a control surface (20, 22) movably mounted on the cell (10, 14); a flight control (30, 32, 34) movably mounted on the cell (10, 14); a trim tab (5) comprising: a chassis (50) fixedly mounted on the cell (10, 14); and an actuator (52) movably mounted on the chassis (50); a mechanical transmission system (40, 42, 44, 46) configured to transmit a force on the one hand between the flight control (30, 32, 34) and the control surface (20, 22), and on the other hand between the actuator (52) and the control surface (20, 22), the mechanical transmission system (40, 42, 44, 46) being devoid of a hydraulic servomotor; and a switch (7) configured so that a force exerted on the flight control (30, 32, 34) causes a switching of the switch (7) from a first state to a second state; the control method comprising controlling the actuator (52) as a function of a switching duration during which the switch (7) is switched into the second state so as to control a movement of the control surface (20, 22) relative to the cell (10, 14).

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