Method and device for assisted piloting of an aircraft comprising a hybrid powerplant, by tactile feedback to indicate the electric energy of the electric energy source

Abstract

Method and device for assisting the piloting of an aircraft comprising a hybrid powerplant by tactile feedback of the electric energy of the electric energy source. The invention relates to a piloting assistance method for an aircraft comprising a hybrid powerplant and a drive train (35) controlling aerodynamic control surfaces (20) and provided with flight controls (31) and control actuators (30) connected to said controls (31). According to the method, at least one characteristic of the hybrid powerplant is measured, then a threshold value related to a functional parameter of the hybrid powerplant is determined as a function of said characteristic. The limit position of the controls (31) for which said functional parameter is equal to said threshold value is then calculated, the current position of the controls (31) is measured, and when the current position exceeds the limit position, a first force is applied to the controls as a function of a predetermined force law as a function of the electric energy of the electric energy source of the hybrid powerplant.

Description

[0001] Cross-reference of related applications

[0002] This application claims the benefit of French patent application FR25 03451, filed on April 2, 2025, the disclosure of which is incorporated herein by reference in its entirety. Technical Field

[0003] This invention relates to the technical field of aircraft flight control, and more particularly to aircraft equipped with a hybrid power system.

[0004] The present invention relates to a method for assisting the piloting of an aircraft equipped with a hybrid power system and a device for assisting the piloting of such an aircraft, so as to indicate the amount of electrical energy of the hybrid power system through tactile feedback.

[0005] An aircraft, particularly a rotorcraft, may include one or more rotors driven to rotate by a power unit. A rotorcraft may, for example, include a lift rotor that participates in lifting or even moving the rotorcraft. A rotorcraft may also include an auxiliary rotor (e.g., a tail rotor) for specifically controlling the yaw motion of the rotorcraft. A rotorcraft may also include one or more propellers.

[0006] In the following text, the term "rotor" may specifically refer to a lifting rotor, an auxiliary anti-torque rotor used to control the yaw motion of a rotorcraft, or a propeller or traction propeller.

[0007] The aircraft also includes at least one aerodynamic control surface controlled by flight actuators (e.g., servo control) to steer the aircraft. The aerodynamic control surface can be a rotor blade. It can also be, for example, an aileron or rudder flap arranged on a wing or tail unit, whose tilt angle is changed by the flight actuators.

[0008] In addition, the drivetrain includes flight control components, such as a joystick, stick, or rudder stick that is operated by the pilot via at least one flight actuator to control aerodynamic control surfaces.

[0009] A flight control system can be described as "electrical," in which flight control components send electrical or optical signals to control flight actuators. Alternatively, a flight control system can be described as "mechanical," in which flight control components are connected to flight actuators via multiple connecting rods to control them.

[0010] Document FR 2946620 describes a control actuator that is directly connected to flight controls and configured to anchor the flight controls in a predefined position, particularly during the operation of an aircraft's autopilot system. This control actuator can also be used as a soft or hard stop.

[0011] In the context of mechanical flight control, this control actuator is called a "parallel actuator" or "trimming actuator," and it directly connects the flight control components or transmission system to a fixed point in the aircraft's structure. This parallel actuator is one of the components of an aircraft's autopilot system.

[0012] The power plant of this aircraft can be a hybrid power system. The hybrid power system may include at least one heat engine and at least one electric motor, as well as a mechanical transmission system that uses the at least one heat engine and / or the at least one electric motor to drive one or more rotors to rotate.

[0013] This hybrid power system also includes electrical energy that powers one or more motors via an electrical system. In the following text, the term "source" is used more simply to refer to the source of electrical energy.

[0014] The source may include, for example, at least one rechargeable battery, a fuel cell connected to a fuel tank, and / or a supercapacitor.

[0015] The motor can be an electric motor that functions solely as a motor, converting electrical energy supplied by a source into mechanical energy that is transmitted to a mechanical transmission system and then to the rotor. Alternatively, the motor can be a generator that functions solely as a generator, converting mechanical energy transmitted by a mechanical transmission system into electrical energy for purposes such as recharging the aircraft's source or powering the aircraft's equipment. The motor can also be an electric generator that functions as both a motor and a generator.

[0016] It should be noted that, for convenience, the term "heat engine" throughout the text refers to any heat engine, such as a turboshaft engine or a piston engine. The term "heat engine" is used in contrast to the terms "electric motor" or "electric motor" used to describe motors driven by electrical energy.

[0017] Depending on the operation of the hybrid power unit, one or more heat engines and one or more electric motors may be installed separately and may be used independently or in combination, simultaneously or sequentially, to drive the rotation of the rotor.

[0018] However, the use of such an electric motor is subject to certain constraints, particularly those related to electrical energy. In fact, the low energy density of batteries compared to the energy density of fuel supplying a heat engine is a constraint considered in the size setting of the hybrid power unit and its operating modes. Therefore, the current amount of electrical energy that can be supplied to the motor in flight is limited, thus allowing for a limited duration of power supply. Furthermore, the electrical energy source can be rechargeable and can therefore be at least partially recharged during flight, thereby increasing its current electrical energy.

[0019] Therefore, the electrical energy of the source can vary significantly during flight. However, the pilot needs to constantly know the amount of electrical energy in the power source in order to determine the aircraft's achievable performance. An indicator can show the pilot this electrical energy from the source, as well as, optionally, the electrical energy consumed by the hybrid power unit. However, the pilot may have a significant workload, preventing him from constantly consulting the indicator. Background Technology

[0020] Document US 10193374 is outside the technical field of this invention. This document describes an electronic device equipped with a battery, such as a smartphone, laptop, tablet, or watch. If desired, the electronic device indicates, for example, that the battery's electrical energy has dropped below a predetermined threshold by generating one or more vibrations.

[0021] The technical background of this invention also includes documents US2005 / 0151672, FR3117447 and FR3021627. Summary of the Invention

[0022] Purpose and Overview of the Invention

[0023] Therefore, one object of the present invention is to provide an apparatus and method for assisting the piloting of an aircraft, which makes it possible to overcome the above-mentioned limitations and aims to reduce the workload of the pilot of the aircraft.

[0024] The primary objective of this invention is to provide a method for assisting the piloting of an aircraft, the aircraft including at least one transmission system for controlling aerodynamic control surfaces, the transmission system being provided with flight control components and control actuators connected to the flight control components, the aircraft including a hybrid power unit, the hybrid power unit being provided with at least one heat engine for driving rotor rotation, at least one electric motor, an electrical energy source, and a mechanical transmission system.

[0025] According to the present invention, this method for assisting in the piloting of an aircraft includes the following steps: • Measure at least one characteristic of the at least one heat engine, mechanical transmission system, or rotor; • Determine a threshold related to the functional parameters of the hybrid power unit based on at least one of the aforementioned characteristics; • Calculate the extreme positions of the flight control unit where the functional parameters of the hybrid power unit are equal to the threshold; • Measure the current position of the flight controls; and • When the current position of the flight control unit exceeds the limit position, a first force is applied to the flight control unit using the control actuator according to a predetermined force law, which is at least a function of the current electrical energy of the source.

[0026] Therefore, the present invention lies in returning force feedback to the pilot of the aircraft substantially in real time via flight controls, the force feedback being a function of the current electrical energy of the source. In this way, the pilot knows the amount of energy available at the source without needing to look at a display, and can remain advantageously focused on the outside of the aircraft and the maneuvers to be performed.

[0027] Optionally, a computer dedicated to performing the method according to the invention on an aircraft is configured to receive measurements of one or more characteristics related to the operation of the at least one heat engine, mechanical transmission system, or rotor, and then determine thresholds related to functional parameters of the hybrid power unit and limit positions of flight controls corresponding to those thresholds. The thresholds and limit positions of the flight controls can be determined using laws previously recorded in or connected to the computer's memory.

[0028] Then, when the limit position is exceeded, the computer controls the actuators to generate a first drag on the flight control components. The first force varies specifically according to the current electrical energy of the source, allowing the pilot of the aircraft to use the flight controls and to know the current electrical energy of the source through this first force.

[0029] When the flight control components return to below their limit positions, the generation of the first force ceases.

[0030] The method according to the invention can be applied to any flight control of an aircraft, controlling the position of aerodynamic control surfaces that influence the movement of the aircraft, and for example, influencing the blades of a rotor, aileron, or rudder flap. In particular, the method can be applied to flight control of the propulsion power of an aircraft. In the specific case of a conventional helicopter, the method can therefore be applied to a lever that controls the collective pitch variation of the lift rotor blades.

[0031] The control actuator can be dedicated to performing the method according to the invention. Alternatively, the control actuator can perform other functions and may optionally take the form of a trim actuator. The control actuator can be a hydraulic, pneumatic, or electric cylinder, or an electric motor.

[0032] The method according to the invention may also include one or more of the following features, used individually or in combination.

[0033] According to one possibility, the current position and extreme position of the flight control unit can be defined by the angle of the flight control unit relative to the reference position of the flight control unit.

[0034] The reference position of the flight control is, for example, the neutral position of the flight control when the pilot does not apply any force to it.

[0035] According to a possibility compatible with the foregoing possibilities, the at least one characteristic may include one or more of the following characteristics: • The internal temperature, engine torque, maximum engine torque, and aging rate of the at least one heat engine, and the duration of use of the heat engine at that engine torque; • The internal temperature of the mechanical transmission system, engine torque, maximum engine torque, and aging rate; and • The rotational speed of the rotor.

[0036] Therefore, one or more characteristics enable characterization of the operation of the at least one heat engine, the power transmission gearbox, and / or the rotor. Thus, one or more of these characteristics can be used to determine threshold values ​​for the functional parameters of the hybrid power unit.

[0037] The aging rate of the at least one heat engine or the power transmission gearbox is, for example, equal to the operating duration, or at least determined based on the operating duration.

[0038] According to one possibility compatible with the foregoing, the method may include measuring at least one environmental parameter from the temperature and atmospheric pressure outside the aircraft; determining a threshold and calculating the limit positions of flight controls that can be executed according to the at least one environmental parameter.

[0039] According to a possibility compatible with the aforementioned possibilities, the calculation of the extreme position can be performed based on a threshold, the functional parameters of the hybrid power unit, the forward speed of the aircraft, and the rotational speed of the rotor.

[0040] In this case, the method involves measuring the forward velocity of the aircraft and the rotational speed of the rotor. These measurements can be performed by sensors dedicated to the process or as part of the aircraft's avionics system, which transmits these measurements to a computer responsible for implementing the method.

[0041] According to a possibility compatible with the foregoing possibilities, the law of predetermined force can be selected from a list including, for example, the following: • Step law, where each step is associated with an interval of the amount of electrical energy of the source, the first force being constant for each step and increasing from one step to the next in parallel with a decrease in the electrical energy of the source. • The law of elasticity states that the first force varies with the current position of the flight control component and the electrical energy of the source. The first force is equal to the current angle of the flight control component multiplied by a first stiffness, which increases continuously or gradually as the electrical energy of the source decreases. The current angle of the flight control component is defined relative to a reference position of the flight control component. • Damping law, the first force varies with the moving speed of the flight control component and the electrical energy of the source, the first force being equal to the moving speed multiplied by a second stiffness, the second stiffness increasing continuously or gradually as the electrical energy of the source decreases; and • Vibration law: The first force generates vibration in the flight control unit each time a predetermined level of electrical energy from the source passes through, the intensity of which is the same each time or gradually increases from one pass to another as the electrical energy from the source decreases.

[0042] In this way, a pilot who knows the predetermined force law that applies to generating the first force can know the current electrical energy of the source through tactile feedback.

[0043] These laws may have been recorded in the computer's memory or connected to the computer's memory.

[0044] According to one possibility compatible with the foregoing possibilities, the method may include supplying non-zero mechanical power to the rotor when the current position of the flight control is greater than or equal to the limit position.

[0045] In this way, once the current position of the flight control unit reaches its limit position—that is, once the power unit's functional parameters are substantially equal to a threshold—the motor is controlled to supply non-zero mechanical power to the rotor via a mechanical transmission system. This mechanical power is then injected into the propulsion system and, individually or in combination with one or more heat engines of the power unit, contributes to driving the rotor's rotation. For this purpose, the computer can control the motor directly or via a controller associated with it.

[0046] Therefore, when the flight controls reach their current extreme position, the motors can be automatically activated to supply non-zero mechanical power to the rotor. This activation can also be signaled by displaying a message on the cockpit screen.

[0047] Alternatively, the motors can be pre-started while supplying zero mechanical power, provided that the current position of the flight control unit is below the limit position.

[0048] Therefore, when the functional parameters reach the threshold, the motor automatically supplies additional power to the rotor in order to, for example, prevent the functional parameters from exceeding the threshold or exceeding the threshold by too much, thereby advantageously maintaining one or more heat engines and mechanical transmission systems.

[0049] When the flight controls return to below their limit position, the motor stops, or alternatively remains running without supplying mechanical power to the rotor.

[0050] According to one possibility compatible with the foregoing possibilities, the functional parameter may be the engine torque of the at least one heat engine, and the threshold may be a predetermined limit torque of the at least one heat engine.

[0051] This limiting torque corresponds, for example, to torque saturation of a heat engine, i.e., the heat engine reaches its operating limit, which can cause damage if exceeded, especially if the exceedance of this limit is sustained for, for example, several minutes. Specifically, this limiting torque is a function of the external temperature and atmospheric pressure of the aircraft. This limiting torque can also be a function of the aging rate of the heat engine.

[0052] According to another example, the limiting torque can be the torque achieved when a heat engine operates at a specific speed. This specific speed allows the heat engine to supply high power and high engine torque, and particularly above the nominal power and torque for a finite duration, such as 30 seconds or five minutes. Beyond this finite duration, degradation of the heat engine or power transmission system is possible.

[0053] In this way, once the current position of the flight control unit reaches its limit position, that is, once the functional parameters related to the operation of the heat engine are substantially equal to the threshold, the motor is controlled to supply non-zero mechanical power to the rotor via the mechanical transmission system.

[0054] Alternatively, the functional parameter could be the engine torque measured in the mechanical transmission system, and the threshold could be the maximum torque of the mechanical transmission system.

[0055] For example, the measured engine torque can be the input or output torque of the power transmission gearbox of the mechanical transmission system, or even the internal torque, and the maximum engine torque is the maximum torque.

[0056] In this way, once the current position of the flight control unit reaches its limit position, that is, once the functional parameters associated with the mechanical transmission system are substantially equal to the threshold, the motor is controlled to supply non-zero mechanical power to the rotor via the mechanical transmission system, thereby limiting or even avoiding the risk that the functional parameters of the mechanical transmission system will exceed the threshold, and subsequently limiting the risk of degradation of the mechanical transmission system.

[0057] Alternatively, the functional parameter can be the rotor's rotational speed, which is equal to the rotational speed of the output shaft of the mechanical transmission system, and the threshold can be the rotor's minimum rotational speed.

[0058] In practice, rotors can have a nominal rotational speed to adhere to, especially in the case of rotorcraft, to ensure safe flight. Therefore, when the rotational speed decreases and falls below this nominal speed, the control motors supply non-zero mechanical power to the rotor in order to maintain an acceptable rotational speed.

[0059] In this case, the law of predetermined force can preferably be the law of damping or the law of vibration, but other laws can be used.

[0060] Alternatively, the functional parameter could be the distance the aircraft can reach, and the threshold could be, for example, the target distance to an identified landing point. This target distance is a function of the amount of fuel available to power one or more heat engines, or even the electrical energy of the source, and can be associated with the maximum mechanical power that the power unit must supply to the rotor to ensure the aircraft travels at least that target distance. In practice, exceeding this maximum mechanical power will specifically result in excessive consumption of the source's electrical energy and will prevent the aircraft from traveling that target distance.

[0061] In this scenario, the motor can supply non-zero power independently of the functional parameters, for example, before the functional parameters reach a threshold. Then, the first force notifies the pilot that by continuing to fly at the current power, the aircraft risks not reaching the landing point.

[0062] According to one possibility compatible with the foregoing possibilities, the source may include a battery, and the current electrical energy of the source is the charging level of the battery's electrical energy.

[0063] Alternatively or complementaryly, the source may include a fuel cell, and the current electrical energy of the source is a function of the amount of fuel contained in the fuel cell housing, and is, for example, proportional to, or even equal to, the amount of fuel (in terms of mass or volume).

[0064] According to one possibility compatible with the foregoing possibilities, the method may include applying a second force to the flight control component using a control actuator when the current position of the flight control component is equal to the limit position.

[0065] Therefore, the second force is applied to the flight control components when the functional parameters of the power unit are substantially equal to the threshold, or if applicable, even when the motor supplies non-zero mechanical power to the rotor.

[0066] Therefore, pilots can use flight controls and second force sensing to know that the power unit's functional parameters have reached a threshold, or even that the motor is supplying non-zero mechanical power to the rotor.

[0067] This perception can take the form of a soft stop, a hard stop, or friction on the flight control components generated by a second force. A soft or hard stop is characterized by requiring a larger or smaller force to bring the controller to a position beyond its limit. Once the limit is exceeded, no force is applied to the flight control components. Friction is characterized by a constant force during the movement of the flight control components, and is therefore independent of their position.

[0068] Therefore, the method according to the invention enables flight control with force feedback, which is controlled according to the use of the aircraft's power unit (in particular one or more motors), so as to first notify the one or more motors of the non-zero mechanical power provided by the one or more motors using a second force signal, and secondly notify the availability level of these one or more motors using a first force signal.

[0069] Therefore, this method can provide the pilot with information in real time about whether he is using one or more motors, as well as information about the amount of electrical energy available to power those one or more motors.

[0070] Furthermore, the second force can be integrated into a predetermined law of force, which thus defines both the first and second forces. Alternatively, the second force can be defined and generated independently of the first force and the law of force.

[0071] Another object of the present invention is a pilot assistance device for an aircraft, the pilot assistance device comprising at least one transmission system for controlling aerodynamic control surfaces, the transmission system being provided with flight control elements and control actuators connected to the flight control elements, the pilot assistance device including a computer connected to the control actuators, the aircraft including a hybrid power unit, the hybrid power unit being provided with a mechanical transmission system for driving rotor rotation, at least one heat engine, at least one electric motor, and a source of electrical energy.

[0072] The notable feature of this driving assistance device is that the computer is configured to implement the aforementioned driving assistance methods.

[0073] Therefore, the flight assistance device may specifically include a measurement sensor for measuring at least one characteristic of the aircraft, an angle sensor for measuring the current position of the flight control components, and a memory for storing predetermined force laws.

[0074] Finally, one object of the present invention is an aircraft comprising at least one transmission system for controlling aerodynamic control surfaces, the transmission system being provided with flight control elements and control actuators connected to the flight control elements, the aircraft comprising a hybrid power unit, the hybrid power unit being provided with a mechanical transmission system for driving rotor rotation, at least one heat engine, at least one electric motor, and an electrical energy source.

[0075] The aircraft includes the pilot assistance devices described above. Attached Figure Description

[0076] The invention and its advantages will become more apparent from the following description of examples given by way of illustration with reference to the accompanying drawings, wherein:

[0077] Figure 1 It is a view of the aircraft's mechanical flight control system;

[0078] Figure 2 This is a view of the aircraft's electrical flight control system;

[0079] Figure 3 It includes according to Figure 1 or Figure 2 A view of the aircraft's flight control system;

[0080] Figure 4 This is a block diagram of the method according to the present invention;

[0081] Figure 5 It is a graph showing the law of step force;

[0082] Figure 6 It is a graph illustrating the law of elasticity;

[0083] Figure 7 It is a graph showing the law of damping force;

[0084] Figure 8 It is a graph illustrating an example of stiffness variation associated with the law of force; and

[0085] Figure 9 This is a graph illustrating an example of stiffness variation associated with the law of force. Detailed Implementation

[0086] Elements present in more than one figure are given the same reference numerals in each figure.

[0087] Figure 1 The mechanical flight control system 3 for aircraft 1 is shown.

[0088] This mechanical flight control system 3 has the function of controlling the movement of one or more aerodynamic control surfaces 20, enabling the aircraft 1 to turn. These aerodynamic control surfaces 20 may include blades of rotor 25, particularly blades of a lifting rotor or rotor for controlling yaw motion, or even blades of a propeller, as well as blades of ailerons or rudder flaps.

[0089] Independent of the properties of the aerodynamic control surface 20, the mechanical flight control system 3 includes at least one transmission system 35 and at least one flight actuator 28 connected to the aerodynamic control surface 20. For example, the flight actuator 28 may be in the form of a servo controller capable of controlling the pitch of one or more blades.

[0090] This mechanical flight control system 3 may include a single flight actuator 28 or multiple flight actuators 28. Furthermore, the flight actuator 28 may be associated with a single mechanical flight control system 3 or shared by multiple mechanical flight control systems 3. For example, the flight actuator 28 controlling the pitch of the rotor blades 25 may be controlled by a collective pitch control stick and a cyclic pitch control stick.

[0091] The transmission system 35 is provided with a flight control element 31 and optionally includes at least one connecting rod, or even, according to this example, multiple connecting rods 32, 33, 34, 36, 37, 38 hinged to each other, or even includes at least one connecting element, such as a bell crank 39, or even a combiner. The transmission system 35 also includes one or more control actuators 30. The reference numeral "30" indicates any control actuator of the transmission system 35, and if necessary, reference numerals "301, 302" indicate specific control actuators.

[0092] according to Figure 1 For example, the drivetrain 35 may include a parallel actuator 301 and a series actuator 302.

[0093] According to the example shown, the parallel or trim actuator 301 connects the connecting rod 33 of the transmission system 35 to a fixed point of the structure 5 of the aircraft 1 via the connecting rod 303. The parallel actuator 301 may include, for example, a cylinder or an electric motor. Therefore, the parallel actuator 301 can act on the flight control unit 31 to anchor the flight control unit 31 in a defined position and / or apply force to the flight control unit 31.

[0094] As illustrated in the example, the series actuator 302 connects the two connecting rods 34 and 36 of the drivetrain 35 together. The series actuator 302 may include, for example, an electric, pneumatic, or hydraulic actuator. Thus, the series actuator 302 can control the flight actuator 28 in a manner transparent to the pilot via the connecting rods 36, 37, and 38, without changing the position of the flight control element 31, which is anchored via the parallel actuator 301.

[0095] Figure 2 An electrical flight control system 4 for an aircraft 1 is shown. Similar to the mechanical flight control system 3, the electrical flight control system 4 may include a single flight actuator 28 or multiple flight actuators 28, which may be associated with a single electrical flight control system 4 or shared by multiple electrical flight control systems 4.

[0096] The electrical flight control system 4 includes a drivetrain 45 and a flight actuator 28 connected to an aerodynamic control surface 20. The drivetrain 45 is equipped with a flight control element 41, such as a mini-joy or joystick, and has at least one electrical connection 42 connecting the flight control element 41 to the flight actuator 28 via at least one computer 27. According to the example shown, the drivetrain 45 may also include one or more control actuators 40 that connect the flight control element 41 to a fixed point on the structure 5 of the aircraft 1. The flight actuator 40 may include, for example, a cylinder or an electric motor. Therefore, the control actuator 40 can act on the flight control element 41 to anchor it in a defined position and / or apply force to it. Thus, the flight control element 41 can transmit digital or analog, electrical or optical signals to the computer 27 that controls the flight actuator 28 using the electrical link 42.

[0097] according to Figure 3 For example, aircraft 1 may be a helicopter including a main rotor 7 and a yaw motion control rotor 8. In this case, aircraft 1 may include multiple mechanical flight control systems 3 or electrical flight control systems 4, each of which is provided with multiple flight control elements 31, 41 and, for example, a collective pitch control stick 75 for controlling the collective pitch change of the first blades 71 of the main rotor 7, a periodic pitch stick 76 for controlling the periodic pitch change of these first blades 71, and a rudder stick 77 for controlling the collective pitch change of the second blades 81 of the yaw motion control rotor 8.

[0098] The aircraft 1 also includes a hybrid power unit 10, which is equipped with one or more heat engines 11, 12, one or more electric motors 13, an electrical energy source 15, and a mechanical transmission system 16 to drive one or more rotors 25. Figure 3 The exemplary aircraft 1 shown has a mechanical transmission system 16 that can rotate the main rotor 7 and the yaw motion control rotor 8.

[0099] Source 15 is electrically connected to one or more motors 13, which may be electric motors, generators, or motor-generators. Source 15 may include one or more (optionally rechargeable) batteries. Alternatively or complementaryly, source 15 may include one or more fuel cells, such as hydrogen or methanol fuel cells, and a tank containing the fuel and connected to one or more fuel cells.

[0100] According to the example shown, the mechanical transmission system 16 includes a power transmission gearbox 17, and the hybrid power unit 10 includes two heat engines 11 and 12 and a motor 13. The two heat engines 11 and 12 and the motor 13 are respectively connected to the input shaft of the power transmission gearbox 17, and the output shaft of the power transmission gearbox 17 is mechanically connected to rotors 7 and 8, respectively, so that rotors 7 and 8 can be driven to rotate jointly or independently by the two heat engines 11 and 12 and the motor 13. The heat engines 11 and 12 are respectively connected to engine computers 111 and 121, such as a Full Authority Digital Engine Control (FADEC) system computer, to control the operation of these heat engines 11 and 12. The motor 13 is connected to a motor controller 131 that controls the operation of the motor 13.

[0101] Alternatively, the motor 13 can be inserted between the heat engine 11 and the mechanical transmission system 16, for example, between the input shaft of the power transmission gearbox 17 of the heat engine 11 and the mechanical transmission system 16. In this case, the motor 13 and the heat engine 11 can jointly drive the rotation of the input shaft of the power transmission gearbox 17.

[0102] Alternatively, the hybrid power unit 10 may include one or more heat engines 11, 12 that drive the rotation of one or more generators 13 to convert mechanical energy into electrical energy to supply source 15 or even other motors 13. The hybrid power unit 10 also includes one or more electric motors 13 that drive one or more rotors 25 to rotate via a mechanical transmission system 16.

[0103] The aircraft 1 may also include a flight assistance device 2 for the aircraft 1. Regardless of whether the aircraft 1 includes one or more mechanical flight control systems 3 or electrical flight control systems 4, the flight assistance device 2 is provided with a computer 9 connected to one or more control actuators 30, 40 of the transmission systems 31, 41. The computer 9 may be dedicated to the flight assistance device 2, or optionally shared with other equipment of the aircraft 1.

[0104] The pilot assistance device 2 may include multiple sensors that measure various characteristics or parameters of the aircraft 1 and are connected to the computer 9 via wired or wireless links. For example, the pilot assistance device 2 may include an angle sensor 21 that measures the current position of the flight controls 31, 41. This current position may be measured relative to a reference position of the flight controls 31, 41, which may be the neutral position of the flight controls 31, 41 when the pilot is not applying any force to them.

[0105] The pilot assistance device 2 may include a thermometer 22 for measuring the external temperature and a barometer 23 for measuring the atmospheric pressure outside the aircraft 1.

[0106] The driving assistance device 2 may include other sensors for measuring characteristics of the power unit 10. For example, the driving assistance device 2 may include a torque meter for measuring engine torque within the mechanical transmission system 16, or even a thermometer and timer for measuring the internal temperature and aging rate of the mechanical transmission system 16, respectively. The driving assistance device 2 may also include one or more tachometers for measuring the rotational speed of the input or output shaft of the mechanical transmission system 16. The driving assistance device 2 may also include torque meters, tachometers, and timers for measuring engine torque, rotational speed on the input or output shaft of the power transmission gearbox 17, and the aging rate of the power transmission gearbox 17.

[0107] Another characteristic of the power unit 10 is the rotational speed of the rotor 25, which is equal to the rotational speed of the output shaft of the mechanical transmission system 16, measured using a tachometer, and in particular the rotational speed of the output shaft of the power transmission gearbox 17.

[0108] Computer 9 can also be connected to engine computers 111 and 121 via a wired or wireless link to receive information related to the characteristics of heat engines 11 and 12. These characteristics of heat engines 11 and 12 may include, for example, the internal temperature of heat engines 11 and 12, engine torque, the duration of use of the engine torque, total service life, and / or aging rate.

[0109] Then, engine computers 111 and 121 and sensors transmit digital or analog, electrical or optical signals to computer 9 via wired or wireless links, which carry the characteristic or the value of the characteristic.

[0110] Furthermore, the computer 9 may include or be connected to a memory. This memory may store one or more code segments or computer programs implemented by the computer 9 to perform steps of the pilot assistance method for the aircraft 1. Figure 4 A block diagram of the driving assistance method is shown. The method may include the following steps.

[0111] First, the method includes measuring 110 one or more characteristics of one or more heat engines 11, 12, mechanical transmission system 16, or rotor 25. This measurement 110 of one or more characteristics can be performed using an engine computer 111, 121 associated with one of the heat engines 11, 12 and / or sensors located on one or more heat engines 11, 12, mechanical transmission system 16, or rotor 25.

[0112] The method may also optionally include measuring one or more environmental parameters. These environmental parameters include, for example, the temperature and atmospheric pressure outside the aircraft 1 measured using thermometer 22 and barometer 23.

[0113] Next, the method includes using computer 9 to determine a threshold 130 related to the functional parameters of hybrid power unit 10 based on one or more previously measured characteristics. This threshold may also optionally be determined based on one or more environmental parameters. The threshold may be determined using one or more laws or tables previously recorded in or connected to the memory of computer 9, and relates to one or more characteristics of one or more heat engines 11, 12, mechanical transmission system 16, and / or rotor 25, or even, where appropriate, one or more environmental parameters.

[0114] The functional parameters can be the engine torque of one of the heat engines 11 or 12, transmitted by the engine computer 111 or 121 associated with the heat engine 11 or 12, or measured by sensors located on the heat engine 11 or 12, and then the threshold is the predetermined limit torque of the heat engine 11 or 12.

[0115] Alternatively, the functional parameter could be engine torque measured by a suitable sensor located in the mechanical transmission system 16, with a threshold value being the maximum torque of the mechanical transmission system 16. In particular, the functional parameter could be engine torque measured by a torque meter located on the input or output shaft of the power transmission gearbox 17.

[0116] Alternatively, the functional parameter could be the rotational speed of rotor 25 as measured by a tachometer, and the threshold could be the minimum rotational speed of rotor 25.

[0117] Alternatively, the functional parameters could be the distance that aircraft 1 can reach, determined by a dedicated computer or avionics system of aircraft 1, and then the threshold and target distance.

[0118] The method then includes calculating the extreme positions of flight controls 31, 41 where the functional parameters of the hybrid power unit 10 are equal to threshold values. This calculation 140 is performed by computer 9 based on the threshold values, the functional parameters of the hybrid power unit 10, or even the forward velocity of aircraft 1 and the rotational speed of rotor 25. This calculation 140 can be performed using one or more laws or tables previously recorded in or connected to the memory of computer 9.

[0119] The method then includes, for example, measuring the current position of flight controls 31, 41 using angle sensor 21. The current position and the extreme positions of flight controls 31, 41 can be defined by the angles of flight controls 31, 41 relative to reference positions of flight controls 31, 41, respectively.

[0120] Finally, the method includes applying a first force of 190° to the flight controls 31 and 41 using control actuators 30 and 40 when the current position of the flight controls 31 and 41 exceeds the limit position, according to a predetermined force law, which is at least a function of the current electrical energy of the source 15. Figure 1 In the context of the example of the mechanical flight control system 3 shown, the first force can be preferentially applied by the parallel actuator 301, although the series actuator 302 can also apply the first force alone, or even in conjunction with the parallel actuator 301.

[0121] To this end, computer 9 transmits digital or analog, electrical or optical signals to control actuators 30 and 40 via wired or wireless links. These signals carry information related to the first force. Then, once the current position of flight control components 31 and 41 exceeds the limit position, control actuators 30 and 40 can generate a first drag on flight control components 31 and 41 by applying a predetermined force law.

[0122] To generate this resistance, when the control actuators 30 and 40 are pneumatic or hydraulic cylinders, the control actuators 30 and 40 adjust the pressure of the fluid in the cylinder chamber according to a predetermined force law and the movement of the flight control components 31 and 41. When the control actuators 30 and 40 are electric motors, the torque applied directly or indirectly by the control actuators 30 and 40 to the flight control components 31 and 41 is adjusted according to a predetermined force law and the movement of the flight control components 31 and 41, for example, as described in document FR 2946620.

[0123] The current electrical energy of source 15 is, for example, the battery's charge level when source 15 includes a battery. When source 15 includes a fuel cell, the current electrical energy can be a function of, or even equal to, the amount of fuel contained in the fuel cell's casing.

[0124] Then, the pilot of aircraft 1 feels the first force when he acts on the flight control components 31, 41, and can become aware of the current electrical energy of source 15. He can then pilot aircraft 1 by sensing a tactile feedback image of the amount of energy available from one or more motors 13 of the hybrid power unit 10.

[0125] One or more predetermined force laws may be recorded in the memory of computer 9 or connected to the memory of computer 9.

[0126] Such a law of predetermined force can be a step law. An example of this step law (see...) Figure 5 (As shown in the figure) includes the amount of electrical energy Q from source 15. nrj The intervals are associated with the steps. Therefore, the first force F is constant for each step, and the amount of electrical energy Q from source 15 is... nrjWhen decreasing, it increases from one step to the next.

[0127] Such a law of predetermined force can also be a law of elasticity. Examples of such a law of elasticity are found in... Figure 6 As shown, the first force F changes with the current position θ of the flight control components 31 and 41. 31、41 And the change. The first force F, for example, is equal to the current angle θ of flight control components 31 and 41. 31、41 The product of the first stiffness K1 and the first stiffness K1 is the amount of electrical energy Q in source 15. nrj It increases when the amount of electrical energy decreases. The first stiffness can increase with the amount of electrical energy Q. nrj Decrease and continuously increase, such as Figure 8 As shown. Alternatively, when this amount of electrical energy Q... nrj As the stiffness decreases, the initial stiffness can be gradually increased, such as... Figure 9 As shown.

[0128] Such a law of predetermined force can be replaced by a law of damping. An example of such a law of damping is found in… Figure 7 As shown, the first force F increases with the moving speed V of the flight control components 31 and 41. 31、41 And it changes. For example, the first force F is equal to the moving speed V. 31、41 The product of the second stiffness K2 and the second stiffness K2 is the amount of electrical energy Q in source 15. nrj It increases when it decreases. For example... Figure 8 As shown, when the amount of electrical energy Q of source 15 nrj When the stiffness is decreased, the second stiffness K2 can be continuously increased. Or, as... Figure 9 As shown, when the amount of electrical energy Q of source 15 nrj When the stiffness is reduced, the second stiffness K2 can be gradually increased.

[0129] Such a predetermined force law can also be a vibration law. According to such a vibration law, a first force generates vibrations in the flight control components 31 and 41 whenever a predetermined level of electrical energy passes through source 15. These predetermined levels correspond, for example, percentages of the maximum capacity of electrical energy 15, such as 25%, 50%, and 75%. The intensity of these vibrations can be the same for each pass. Alternatively, as the amount of electrical energy in source 15 decreases, the intensity of these vibrations can increase from one pass to the next.

[0130] For example, when several predetermined force laws are recorded in memory, computer 9 can automatically select the law to be used according to the flight phase. In particular, the avionics system of aircraft 1 can typically provide computer 9 with signals carrying information related to that flight phase. Alternatively, before takeoff or during flight, the pilot can use a human-machine interface (e.g., a screen with a touch panel) connected to computer 9 to select the law to be used.

[0131] Furthermore, the method of the present invention may include supplying 160 non-zero mechanical power to the rotor 25 by the motor 13 when the current position of the flight controls 31, 41 is greater than or equal to the limit position. Then, the computer 9 transmits digital or analog, electrical or optical signals to the motor 13 or the motor controller 131 associated with the motor 13 via a wired or wireless link, the signals carrying information related to the mechanical power that the motor 13 must supply.

[0132] In fact, when the current position of the flight control components 31 and 41 reaches or even exceeds the limit position, the functional parameters of the power unit 10 reach or even exceed the associated threshold.

[0133] Therefore, for example, to protect the power unit 10 and to prevent the deterioration of the heat engines 11, 12 or the mechanical transmission system 16, the motor 13 can supply non-zero mechanical power to help drive the rotation of the rotor 25. This is especially true when the functional parameters involve the heat engines 11, 12 or the mechanical transmission system 16.

[0134] Furthermore, when the functional parameter is the rotational speed of rotor 25 and the parameter decreases until it reaches an associated threshold, motor 13 can then provide non-zero mechanical power to help drive the rotation of rotor 25, so as to limit the decrease in the rotational speed, or even increase it.

[0135] Furthermore, the method according to the invention may include applying a second force of 180° to the flight controls 31, 41 using control actuators 30, 40 when the current position of the flight controls 31, 41 is equal to the limit position. Thus, the pilot knows that the functional parameters have reached the associated threshold, or that the motor 13 begins to supply non-zero mechanical power to facilitate the rotation of the rotor 25. This second force may take the form of a soft stop, a hard stop, friction, or even vibration.

[0136] To this end, computer 9 transmits digital or analog, electrical or optical signals to control actuators 30 and 40 via wired or wireless links. These signals carry information related to the second force. Once the current position of flight control components 31 and 41 reaches its limit position, control actuators 30 and 40 can generate a second drag on the flight control components 31 and 41.

[0137] Naturally, the present invention can be subject to many variations in its implementation. Although several embodiments have been described above, it should be readily understood that it is impossible to exhaustively list all possible embodiments. Without departing from the scope of the invention and the claims, any of the described devices can naturally be replaced by equivalent means.

Claims

1. A method for assisting the piloting of an aircraft (1), the aircraft (1) comprising at least one transmission system (35, 45) for controlling aerodynamic control surfaces (20), the transmission system (35, 45) being provided with flight control elements (31, 41) and control actuators (30, 40) connected to the flight control elements (31, 41), the aircraft (1) comprising a hybrid power unit (10), the hybrid power unit (10) being provided with a mechanical transmission system (16) for driving rotors (25) to rotate, at least one heat engine (11, 12), at least one electric motor (13), and a source of electrical energy (15). The method includes the following steps: Measure at least one characteristic of the at least one heat engine (11, 12), the mechanical transmission system (16), or the rotor (25); Based on the at least one characteristic, a threshold related to the functional parameters of the hybrid power unit (10) is determined (130); Calculate the extreme positions of the flight control unit (31, 41) where the functional parameters of the hybrid power unit (10) are equal to the threshold; Measure the current position of the flight control components (31, 41) mentioned in (150); as well as When the current position of the flight control unit (31, 41) exceeds the limit position, the control actuator (30, 40) applies a first force (190) to the flight control unit (31, 41) according to a predetermined force law, which is at least a function of the current electrical energy of the source (15).

2. The method according to claim 1, This includes supplying (160) non-zero mechanical power from the motor (13) to the rotor (25) when the current position of the flight control unit (31, 41) is greater than or equal to the extreme position.

3. The method according to claim 1, in, The current position and the extreme position of the flight control unit (31, 41) are respectively defined by the angle of the flight control unit (31, 41) relative to the reference position of the flight control unit (31, 41).

4. The method according to claim 1, include: When the current position of the flight control unit (31, 41) is equal to the limit position, the control actuator (30, 40) applies a second force (180) to the flight control unit (31, 41).

5. The method according to claim 1, in, The at least one characteristic includes one or more of the following characteristics: The internal temperature, engine torque, maximum engine torque, and aging rate of the at least one heat engine (11, 12); The internal temperature, engine torque, maximum engine torque, and aging rate of the mechanical transmission system (16); as well as Rotational speed of rotor (25).

6. The method according to claim 1, include: (120) At least one environmental parameter is measured from the temperature and atmospheric pressure outside the aircraft (1), (130) the threshold is determined, and (140) the extreme position of the flight control unit (31, 41) is calculated, which is a function of at least one environmental parameter.

7. The method according to claim 1, in, The functional parameter is the engine torque of the at least one heat engine (11, 12), and the threshold is a predetermined limit torque of the at least one heat engine (11, 12).

8. The method according to claim 1, in, The functional parameter is the torque measured in the mechanical transmission system (16), and the threshold is the maximum torque of the mechanical transmission system (16).

9. The method according to claim 1, in, The functional parameter is the rotational speed of the rotor (25), and the threshold is the minimum rotational speed of the rotor (25).

10. The method according to claim 1, The functional parameter is the distance that the aircraft (1) can reach, and the threshold is the target distance.

11. The method according to claim 1, in, The calculation of the extreme positions of the flight control components (31, 41) in (140) is performed based on the threshold of the functional parameters of the hybrid power unit (10), the forward speed of the aircraft (1), and the rotational speed of the rotor (25).

12. The method according to claim 1, The predetermined force law mentioned above is selected from: The step law states that each step is associated with an interval of electrical energy from the source (15), the first force is constant for each step, and the first force increases from one step to the next in parallel with the decrease of electrical energy from the source (15). The elastic law states that the first force varies with the current position of the flight control unit (31, 41) and is equal to the current angle of the flight control unit (31, 41) multiplied by a first stiffness, which increases continuously or gradually as the electrical energy of the source (15) decreases, and the current angle of the flight control unit is defined relative to the reference position of the flight control unit (31, 41). According to the damping law, the first force varies with the moving speed of the flight control components (31, 41), and the first force is equal to the moving speed multiplied by the second stiffness, which increases continuously or gradually as the electrical energy of the source (15) decreases. as well as According to the law of vibration, the first force generates vibration in the flight control components (31, 41) each time the electrical energy of the source (15) reaches a predetermined level, and the intensity of the vibration is the same each time it passes, or gradually increases from one pass to another as the electrical energy of the source (15) decreases.

13. The method according to claim 1, in, The source (15) includes a battery, and the current electrical energy is the electrical charge level of the battery, or the source (15) includes a fuel cell, and the current electrical energy is a function of the amount of fuel contained in the tank of the fuel cell.

14. A pilot assistance device (2) for an aircraft (1), the pilot assistance device (2) comprising at least one transmission system (35, 45) for controlling an aerodynamic control surface (20), the transmission system (35, 45) being provided with flight control elements (31, 41) and control actuators (30, 40) connected to the flight control elements (31, 41), the pilot assistance device (2) comprising a computer (9) connected to the control actuators (30, 40), the aircraft (1) comprising a hybrid power unit (10), the hybrid power unit (10) being provided with a mechanical transmission system (16) for driving a rotor (25) to rotate, at least one heat engine (11, 12), at least one electric motor (13), and a source of electrical energy (15). in, The computer (9) is configured to implement the method according to claim 1, wherein the driving assistance device (2) includes a measurement sensor for measuring at least one characteristic of the aircraft (1), an angle sensor (21) for measuring the current position of the flight control unit (31, 41), and a memory for storing the predetermined force law.

15. An aircraft (1) comprising at least one transmission system (35, 45) for controlling aerodynamic control surfaces (20), the transmission system (35, 45) being provided with a flight control system (31, 41) and control actuators (30, 40) connected to the flight control system (31, 41), the aircraft (1) comprising a hybrid power unit (10), the hybrid power unit (10) being provided with a mechanical transmission system (16) for driving a rotor (25) to rotate, at least one heat engine (11, 12), at least one electric motor (13), and a source of electrical energy (15). in, The aircraft (1) includes a pilot assistance device (2) according to claim 14.

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