Device and method for generating a fluttering phenomenon on at least a part of an aircraft.

Abstract

- Method and device for generating a fluttering phenomenon on at least part of an aircraft.- The device (1) comprises a plurality of sensors (4) arranged on the aircraft (AC) and configured to measure input data, an avionics computer (5) configured to determine, from said input data, at least one control command for at least one control surface (3) of the aircraft (AC) using a control law comprising at least one gain value, said control law being configured to obtain a control command enabling the generation of a flutter mode on at least one part (2) of the aircraft (AC), and a control system (6) configured to control the control surface (3) of the aircraft (AC) so as to generate said flutter mode using the control command determined by the avionics computer (5), said device (1) thus enabling the generation of a flutter phenomenon under controlled conditions and the easy and safe study of the behavior of the aircraft (AC) subjected to this flutter. Figure for the abbreviation: Figure 1.

Description

Title of the invention: Device and method for generating a fluttering phenomenon on at least one part of an aircraft. Technical field

[0001] The present invention relates to a device and a method for generating a flutter phenomenon on at least one part of an aircraft. State of the art

[0002] During a flight, an aircraft is subjected to aerodynamic forces which deform its structure. These deformations coupled with the airflow on the external surfaces of the aircraft can generate undesirable vibratory phenomena, in particular at the level of the wing. Indeed, for an aircraft speed greater than a critical speed, aeroelastic instabilities appear leading to an oscillation phenomenon called flutter. This flutter phenomenon can be harmful to the aircraft which leads to limiting its maximum speed. It is therefore interesting to have solutions to counter the flutter phenomenon.

[0003] One method of flutter suppression involves modifying the damping properties of the aircraft's soft modes (or at least certain parts) in flight by acting on control surfaces, for example ailerons, flaps or spoilers of the wings or the rudder. This method is developed using simulations and requires that its effectiveness be evaluated in real conditions. To do this, a test aircraft is brought into conditions at the limit of flutter, i.e. at a speed close to its critical speed, so as to verify whether the damping increase capabilities make it possible to obtain effective flutter suppression or not.

[0004] However, this method is not completely satisfactory. Indeed, bringing an aircraft close to its critical speed presents risks. First of all, it assumes a high degree of confidence in knowing the precise critical speed of a given aircraft. In addition, this test method does not allow the robustness of flutter suppression to be assessed, nor does it allow the effectiveness of the self-adaptation algorithms associated with said flutter suppression to be assessed, if applicable. Furthermore, in the event of unexpected behavior during the test, the only way to return to stable flight conditions is to reduce the aircraft speed to move away from the critical speed, which is not instantaneous.

[0005] There is therefore a need to find a safer and more extensive solution for testing such a flutter suppression method on an aircraft. Statement of the invention

[0006] The object of the present invention is to propose a solution to overcome the aforementioned drawbacks. It relates to a method for generating a fluttering phenomenon on at least one part of an aircraft.

[0007] According to the invention, the method comprises at least the following series of successive steps, implemented repetitively: - a measurement step, implemented by a plurality of sensors arranged on the aircraft, to measure input data; - a data processing step, implemented by an avionics computer, to, from the input data measured in the measurement step, determine at least one control order for at least one control surface of the aircraft, the control order being determined using a control law comprising at least one adjustable parameter corresponding to a gain value, said control law being configured to obtain a control order making it possible to generate a flutter mode on at least one part of the aircraft; and - a control step for controlling the control surface so as to generate said floating mode using the control order determined in the data processing step.

[0008] Thus, thanks to the invention, it is possible to generate a flutter mode on a part of the aircraft without needing to approach its critical speed. It is therefore possible to study the behavior of the aircraft when it is subjected to a flutter phenomenon, and this, under controlled and safe conditions. Indeed, in the event of an undesirable event during the generation of the flutter, it is sufficient to stop the implementation of the method to almost instantly return to a stable situation for the aircraft.

[0009] Advantageously, the method comprises an adjustment step, implemented by an adjustment unit, to adjust the adjustable parameter(s) of the control law in real time, so as to obtain a control order for the control surface of the aircraft making it possible to generate a predetermined floating mode.

[0010] Furthermore, advantageously, the method comprises a data preprocessing step, implemented by the avionics computer before the data processing step, to calculate an input signal for the control law from the input data acquired in the acquisition step, using the following equation: U= E" / û; )-.V.aA nslaque ^ U is the input signal of the control law; N is the number of sensors; ai is an acceleration measured by one of said sensors; and ar is an acceleration measured by one of said sensors configured to measure a acceleration related to a rigid movement of the aircraft.

[0011] In a particular embodiment, the method comprises a data post-processing step, implemented by the avionics computer after the data processing step, to apply at least one of the following limitations to the control order determined by the control law: a deployment amplitude limitation, a deployment speed limitation, a delay.

[0012] Furthermore, advantageously, the control law comprises at least one filtering chain.

[0013] In a particular embodiment, the method comprises a monitoring step, implemented by a monitoring unit, for recording the input data measured by the plurality of sensors over time, for comparing said input data with predetermined threshold values, and, if at least one of said input data is greater than said corresponding predetermined threshold value, for inhibiting the deployment of the control surface of the aircraft controlled by the control order.

[0014] The present invention also relates to a device for generating a flutter phenomenon on at least one part of an aircraft. According to the invention, said device comprises at least: - a plurality of sensors arranged on the aircraft and configured to measure input data; - an avionics computer arranged on the aircraft and configured to determine, from said input data, at least one control order for at least one control surface of the aircraft, the control order being determined using a control law comprising at least one adjustable parameter corresponding to a gain value, said control law being configured to obtain a control order making it possible to generate a flutter mode on at least part of the aircraft; and - a control system configured to control said control surface so as to generate said floating mode using the control order determined by the avionics computer.

[0015] Advantageously, the device comprises an adjustment unit comprising an interactive interface making it possible to adjust the adjustable parameter(s) of the control law in real time, so as to obtain a control command for the aircraft's control surface making it possible to generate a predetermined floating mode.

[0016] Furthermore, in a particular embodiment, the device comprises a monitoring unit configured to record the input data measured by the plurality of sensors over time, to compare said input data to predetermined threshold values, and, if at least one of said input data is greater than said corresponding predetermined threshold value, to inhibit the deployment of the aircraft control surface controlled by the command order.

[0017] The present invention also relates to an aircraft. According to the invention, the aircraft comprises at least one device for generating a flutter phenomenon as described above. Brief description of the figures

[0018] The attached figures will make it clear how the invention can be implemented. In these figures, identical references designate similar elements.

[0019] [Fig.l] is a perspective view of an aircraft comprising a device for generating a flutter phenomenon according to a particular embodiment.

[0020] [Fig.2] is a block diagram of a method for generating a flutter phenomenon according to a particular embodiment.

[0021] [Fig.3] is a functional diagram of a control law used by the device for generating a flutter phenomenon of [Fig.l]. Detailed description

[0022] The device 1 for illustrating the invention is shown schematically in a particular embodiment in [Fig. 1]. It is a device for generating a fluttering phenomenon capable of generating flutter on all or part of an aircraft AC, in particular a transport aircraft.

[0023] By "flutter phenomenon" or "flutter" is meant a vibrational resonance phenomenon caused by aeroelastic coupling that appears naturally on an aircraft in flight when it reaches or exceeds a critical speed specific to it and which depends on its design. The objective of the device 1 is to artificially generate (or cause) the appearance of flutter on the aircraft AC in flight, without needing to exceed or approach too close to its critical speed.

[0024] The device 1 is particularly suitable for studying the flutter behavior of an aircraft under controlled conditions. It can, for example, allow solutions to be tested for eliminating or reducing flutter. This is the subject of a preferred application which will be detailed later, in which the aircraft AC corresponds to a test aircraft used to evaluate the performance of a flutter suppression function. In this preferred application, the device 1 is configured to generate flutter on the aircraft AC in a first step. Then, in a second step, a system (not described in detail in the present description), intended to counter this flutter, is tested so as to evaluate its ability to effectively counter the flutter.

[0025] In a particular embodiment, the device 1 is configured to carry out measurements on a part 2 of the aircraft AC, on which it is desired to generate flutter. From these measurements, the device 1 is able to determine an order of control for control surfaces 3 of the aircraft AC. The deployment (or movement) of these control surfaces 3 makes it possible to generate a desired floating mode on part 2, as explained below.

[0026] In the particular embodiment shown in [Fig. 1], the part 2 on which it is desired to generate flutter corresponds to wings 11 of the aircraft AC. However, in other embodiments, the flutter may be generated on one or more other parts of the aircraft AC, for example on a part of the fuselage or the rudder.

[0027] The expression "flutter mode" refers to the unstable vibration modes obtained naturally on certain parts of the aircraft AC when the latter exceeds its critical speed. Furthermore, "obtained naturally" means obtained under normal flight conditions, that is to say without acting on any element of the aircraft AC modifying its structural and aerodynamic properties in order to modify its vibration response.

[0028] The device 1 comprises a plurality of sensors, for example accelerometers 4, arranged on the part 2 of the aircraft AC. In the remainder of the description, the sensors will be illustrated by accelerometers, but could of course be other types of sensors. The accelerometers 4 are configured to enable the vibration modes of the part 2 to be measured when the aircraft AC is in flight. More precisely, they are configured to measure the excitation linked to the vibration modes of the part 2, namely the deformations undergone by said part 2 in flight.

[0029] The arrangement of the accelerometers 4 is chosen judiciously so that they can best measure the vibration modes on the part 2, in particular the flutter mode that it is desired to generate. The flutter mode to be generated being previously targeted via simulations using aeroelastic models of the aircraft AC, its characteristics are known. Consequently, it is possible to identify preferred locations where to position the accelerometers 4.

[0030] Furthermore, preferably, the accelerometers 4 are arranged in proximity to the control surfaces 3 used to generate the desired floating mode. This makes it possible to effectively measure the effect produced by the deployment of said control surfaces 3 and to improve the efficiency of the device 1.

[0031] The plurality of accelerometers 4 comprises an accelerometer 4R configured to measure an acceleration linked to a particularly rigid movement of the aircraft AC. This acceleration, called rigid acceleration, corresponds to a rigid dynamic component of the overall movement of the aircraft AC in flight. It is capable of serving as a reference as explained below. Preferably, the accelerometer 4R is arranged close to the center of gravity of the aircraft AC or in an area whose dynamics are consistent with said center of gravity.

[0032] For reasons of simplification of the drawings, only a few accelerometers 4 are shown on each wing 11 of the aircraft AC. However, the device 1 may comprise a higher number of accelerometers 4 to precisely measure the vibration modes, for example several tens of accelerometers.

[0033] Furthermore, in the embodiment of [Fig.l], the control surfaces 3 correspond to internal ailerons 10 of the wings 11 of the aircraft AC. Indeed, in this particular embodiment, these control surfaces 3 are adapted to generate the targeted flutter mode on the part 2, namely the wings 11. However, in other embodiments, the control surfaces 3 may correspond to other usual control surfaces of the aircraft AC. The control surfaces 3 are chosen according to the application considered. Indeed, they must be capable, by their deployment during a flight, of modifying the properties of the part 2 considered, in particular its aerodynamic damping characteristics, so as to obtain the appearance of the targeted flutter mode.

[0034] Furthermore, the device 1 comprises an avionics computer 5 configured to acquire and process the data measured by the accelerometers 4 in order to determine the control order for the control surfaces 3 of the aircraft AC. The avionics computer 5 is linked to the accelerometers 4 via a data transmission link, as shown schematically by dashed lines in [Fig.l]. Preferably, the avionics computer 5 corresponds to a standard flight management system of the FMS type (for “Flight Management System” in English).

[0035] As shown schematically in [Fig. 3], in a particular embodiment, the avionics computer 5 is configured to carry out a series of data processing operations in order to obtain the control order (denoted C) for the control surfaces 3.

[0036] The avionics computer 5 firstly performs operations on the data measured by the accelerometers 4, called input data A, to obtain an input signal U. The avionics computer 5 is configured to calculate the input signal U using the following equation: u= E^)-W anslaquelle: N is the number of accelerometers 4 (without the 4R accelerometer); a> is an acceleration measured by one of the 4 accelerometers; and ar is the rigid acceleration measured by the 4R accelerometer arranged on a rigid motion of the aircraft AC.

[0037] The avionics computer 5 therefore calculates the sum of the accelerations ai measured by the accelerometers 4 by subtracting the rigid acceleration ar, which makes it possible to have a reference acceleration so as not to take into account the overall movement of the aircraft AC in the measurement of the vibration modes on the part 2.

[0038] The avionics computer 5 then applies a control law to the input signal U to obtain an output signal Y. This control law is in the form of a transfer function F comprising a gain value G and a filter chain H. The gain value G and parameters of the filter chain H correspond to adjustable parameters allowing the control law to be adapted as described below.

[0039] The product of the input signal U and the gain value G makes it possible to obtain a control order to move the control surfaces 3 in a way that modifies the vibration mode of the part 2 of the aircraft AC. Indeed, by moving the control surfaces 3, the structural and aerodynamic properties of the aircraft AC (at least those of the part 2) and therefore its vibrational behavior are modified. This leads to modifying the speed for which particular vibration modes appear, in particular flutter modes.

[0040] Indeed, thanks to simulations based on previously established aeroelastic models of the aircraft AC, it is possible to identify the vibration modes of the part 2 as a function of the flight conditions of the aircraft AC, in particular its speed. It is also possible to determine how to modify certain properties of the part 2, such as the damping, to modify the properties of the vibration mode obtained for a given speed. Knowing this, it is possible to determine how to move the control surfaces 3 in order to modify said properties of the part 2, as desired. The deployment of the control surfaces 3 which is controlled by the control order is directly linked to the gain value G. Consequently, the choice of the gain value G makes it possible to determine the vibration mode obtained in the flight conditions considered.

[0041] Thus, by correctly choosing the gain value G, it is therefore possible to generate a desired flutter mode at a given speed. The higher the gain value G (in absolute value), the greater the instability linked to flutter and the more the speed of the aircraft AC for which said flutter appears decreases.

[0042] Furthermore, the filtering chain H makes it possible to improve the efficiency of the control law. For example, it can be configured to obtain suitable synchronization between the deployment of the control surfaces 3 and the measurement frequency of the accelerometers 4. It can also be configured to obtain an amplification of the input signal U for frequencies close to that corresponding to the flutter mode that it is desired to generate. Conversely, it can also be configured to reduce the output signal Y for frequency ranges far from that corresponding to the targeted flutter mode.

[0043] As a non-limiting example, a function representing the filtering chain H can be presented in the form of the following equation: .s A ( / 2>tA„2.0,7 ).5+( / 2^)2 in which: .5'2+ ( / 2^.2.0,7) ^+ ( f2rr )2 5 is a variable; f is the frequency of the flutter mode to be generated, in Hertz (Hz); and Am is a predetermined amplification factor.

[0044] The output signal Y obtained at the output of the transfer function F corresponds to a raw control order which can be used to control the control surfaces 3 of the aircraft AC. However, in the embodiment of [Fig. 3], the output signal Y is not used directly. Indeed, the avionics computer 5 is configured to carry out data post-processing operations on the signal Y so as to refine it to obtain the final control order C, actually used to control the control surfaces 3. The data post-processing operations correspond to limitations denoted D, L1 and L2 in [Fig. 3]. In a non-limiting manner, this may be a deployment amplitude limitation (L1), a deployment speed limitation (L2) and / or a delay (D).

[0045] Thus, the control order C used to control the control surfaces 3 can be adapted so as to avoid an excessively large or excessively sudden deployment of said control surfaces 3, which could be undesirable depending on the application considered, for example in terms of maneuverability of the aircraft AC. A delay, for its part, can make it possible to improve the synchronization between the deployment of the control surfaces 3 and the measurements taken by the accelerometers 4.

[0046] As a non-limiting example, it is considered that the floating mode that one wishes to generate has a frequency / (given value in Hz) and that there exists, without the application of a delay (delay) D, a phase shift of -40° between the input signal U obtained from the input data A and the output signal Y of the control law. In this case, a function representing a delay D making it possible to obtain a control order C in phase with the input signal U can be presented in the form of the following equation: 360. /

[0047] Furthermore, as shown schematically in [Fig.l], the device 1 comprises a standard control system 6 configured to control the deployment of the control surfaces 3. The avionics computer 5 transmits the control order C to the control system 6. The latter is linked to the control surfaces 3 via a data transmission link, as shown schematically by dashed lines in [Fig.l], so as to be able to control them and generate the desired floating mode.

[0048] Thus, thanks to the device 1, it is possible to generate a floating mode on the part 2 of the AC aircraft without having to approach its critical speed. It is therefore possible to study the behavior of the AC aircraft when it is subjected to a flutter phenomenon, and this, under controlled and safe conditions. Indeed, in the event of an undesirable event during the generation of flutter, it is sufficient to deactivate device 1 to return to a stable situation almost instantly.

[0049] Furthermore, as shown schematically in [Fig.l], the device 1 comprises an adjustment unit 7 comprising an interface for adjusting the adjustable parameters of the control law. This interface is configured to allow a user to act on the adjustable parameters of the control law by modifying them in real time, namely during the use of the device 1. For example, this interface may comprise a screen and input means such as a keyboard or a touch screen.

[0050] The adjustment unit 7 makes it possible to modify in real time the floating mode that one wishes to generate on the part 2 of the aircraft AC, in particular by acting on the gain value G. It can also make it possible to adjust the performance of the control law, in particular by acting on the parameters of the filtering chain H.

[0051] The adjustment unit 7 can also allow modification of other adjustable parameters such as the parameters of the limitations D, L1 and L2 applied during data post-processing.

[0052] Furthermore, as shown schematically in [Fig.l], the device 1 comprises a monitoring unit 8 for monitoring the behavior of the part 2 of the aircraft AC when using the device 1. The monitoring unit 8 is configured to record and monitor in real time the input data A measured by the accelerometers 4. For example, the avionics computer 5 can be configured to transmit said input data A to the monitoring unit 8 as and when said avionics computer 5 acquires them. The monitoring unit 8 stores the input data A thus received in a memory so as to be able to establish and use a history of the measurements.

[0053] In addition, the monitoring unit 8 is configured to compare the input data A transmitted by the avionics computer 5 with a table of predetermined threshold values ​​not to be exceeded. If one or more input data A is greater than said corresponding predetermined threshold value in said table of predetermined threshold values, the monitoring unit 8 is configured to inhibit the deployment of the control surfaces 3, controlled by the control order C. In this case, the avionics computer 5 is configured to ignore the control order C and to reposition the control surfaces 3 in their original position.

[0054] Depending on the application considered, the monitoring unit 8 can be integrated directly into the aircraft AC or be remote so as to carry out remote monitoring.

[0055] Thus, in the event that an abnormal or undesired evolution of the floating mode obtained on part 2 of the aircraft AC is detected, the monitoring unit 8 makes it possible to avoid any risk by guaranteeing a rapid return to a stable situation.

[0056] The device 1 described above is configured to implement a method P for generating a flutter phenomenon. As shown schematically in [Fig.2] in a particular embodiment, the method P comprises the series of successive steps E1, E2, E3, E5 and E6, implemented by the device 1.

[0057] Step E1 is a measurement step, implemented by the accelerometers 4, to measure the input data A.

[0058] Step E2 is a data processing step, implemented by the avionics computer 5, to determine the control order for the control surfaces 3 making it possible to generate the desired floating mode on the part 2 of the aircraft AC. The data processing carried out in step E2 corresponds to the application of the control law to the input data A, as described above.

[0059] Steps E5 and E6 are steps, respectively, of preprocessing and postprocessing of data implemented by the avionics computer 5. Step E5 is put before step E2 to calculate the input signal U of the control law from the input data A, as described above. Step E6, for its part, is implemented after step E2 to apply the limitations D, L1 and L2 to the output signal Y.

[0060] Step E3 is a control step, implemented by the control system, to control the control surfaces 3 so as to generate the desired floating mode using the control order C transmitted by the avionics computer 5.

[0061] The steps of the method P are implemented in a repetitive, continuous manner, thus forming an active control loop. Indeed, the vibrational behavior of the part 2 makes it possible to determine a control order C which moves the control surfaces 3. This deployment modifies the properties of the aircraft AC, which modifies the structural response of the part 2 which ultimately leads to a new vibrational behavior. The new vibrational behavior is measured by the accelerometers 4 which makes it possible to determine a new control order C for the control surfaces 3, and so on.

[0062] In this way, a coupling is obtained between the structural response of part 2 and the control law which makes it possible to change the vibratory behavior of said part 2 until the desired floating mode is generated.

[0063] Furthermore, the method P comprises an adjustment step E4, implemented by the adjustment unit 7, to adjust in real time the adjustable parameters of the control law. As detailed above, step E4 makes it possible, in particular, to target the floating mode that one wishes to generate and to adjust the performance of the control law. order.

[0064] Furthermore, the method P comprises a monitoring step E7 implemented by the monitoring unit 8. Step E7 is implemented in parallel with steps E1 to E6 to record the input data A measured by the accelerometers 4 and compare them with the threshold values ​​not to be exceeded. If one of the input data A is greater than the corresponding threshold value (to which said input data A is compared), step E7 inhibits the deployment of the control surfaces 3 controlled by the control order determined in step E2.

[0065] A preferred application of the device 1 concerns the study of the flutter phenomenon on a test aircraft and the evaluation of the performance of a flutter suppression system on this same aircraft. The objective of this application is to verify the effectiveness of the flutter suppression system in real flight conditions.

[0066] In this application example, it is considered that the design of the aircraft AC predisposes it to a flutter mode at the wings 11 for a certain critical speed. It is desired to avoid the occurrence of this flutter mode by implementing the aforementioned flutter suppression system. This would make it possible to increase the speed range usable by the aircraft AC. Indeed, if the occurrence of the undesirable flutter mode is prevented, the critical speed would be usable without risk for the aircraft AC.

[0067] The flutter suppression system is therefore implemented by a method whose object is not part of the present invention and which is therefore not explained in detail. In this example, it is considered that the wings 11 of the aircraft AC comprise external ailerons 9 making it possible to influence the flutter mode that it is desired to suppress. The flutter suppression system is therefore configured to act on these external ailerons 9 so as to counter the flutter mode to be suppressed.

[0068] In the context of this preferred application, it is desired to verify that the flutter suppression system is correctly configured by testing it under real conditions, without however bringing the aircraft AC so that its speed is close to the critical speed. Also, the device 1 is used to generate the aforementioned flutter mode on the wings 11 of the aircraft AC. To do this, it is necessary to identify the control surfaces 3 of the aircraft AC which make it possible to generate the desired flutter mode on the wings 11, other than the external ailerons 9 (because they are already used by the flutter suppression system).

[0069] In this example, it is considered that the aircraft AC has other ailerons, for example the internal ailerons 10, which can be used by the device 1 to generate the desired flutter mode. Consequently, the accelerometers 4 are arranged on the wings 11 in a suitable manner, close to the internal ailerons 10. In addition, configures the control law of the device 1 so as to determine a control order C for the internal ailerons 10. In this way, it is possible to generate the desired flutter mode on the wings 11.

[0070] One approach to obtaining a particular floating mode may be to proceed in stages. We begin by determining the theoretical gain value G allowing to obtain a control order C which would generate the targeted floating mode. We then configure the control law with a gain value G lower than that required, then we gradually increase this value until actually generating the targeted floating mode.

[0071] Once the desired flutter mode is generated on the wings 11, the flutter suppression system is implemented using the external ailerons 9. In this way, the effectiveness of said flutter suppression system can be tested under controlled conditions. In addition, its robustness and, if present, its self-adaptation algorithms can also be tested. Indeed, thanks to the adjustment unit 7, the control law can be adjusted during the test flight to vary the characteristics of the flutter mode generated on the wings 11. This can make it possible to evaluate the capabilities of the flutter suppression system to adapt to variations in the flutter mode to be suppressed.

[0072] The device 1 for implementing the method P as described above has numerous advantages. In particular: - it allows a flutter phenomenon to be generated on the AC aircraft under controlled conditions without needing to approach the critical speed; - it allows to generate a particular floating mode which can be targeted by adjusting the adjustable parameters of the control law; - it allows the study of the efficiency and robustness of a flutter suppression device for flutter modes with varied characteristics, by simply modifying the adjustable parameters of the control law; and - it is particularly safe since it allows for an easy and rapid return to a stable situation in the event of an adverse event.

Claims

Claims

1. Method for generating a flutter phenomenon on at least one part of an aircraft, characterized in that it comprises at least the following series of successive steps, implemented repetitively: - a measurement step (El), implemented by a plurality of sensors (4) arranged on the aircraft (AC), to measure input data (A); - a data processing step (E2), implemented by an avionics computer (5), to, from the input data (A) measured in the measurement step (El), determine at least one control order (C) for at least one control surface (3) of the aircraft (AC), the control order (C) being determined using a control law comprising at least one adjustable parameter corresponding to a gain value (G), said control law being configured to obtain a control order (C) making it possible to generate a flutter mode on at least one part (2) of the aircraft (AC);and - a control step (E3), implemented by a control system (6), for controlling the control surface (3) so as to generate said floating mode using the control order (C) determined in the data processing step (E2).;

2. Method according to claim 1, characterized in that it comprises an adjustment step (E4), implemented by an adjustment unit (7), to adjust the adjustable parameter(s) of the control law in real time, so as to obtain a control order (C) for the control surface (3) of the aircraft (AC) making it possible to generate a predetermined floating mode.

3. Method according to any one of the preceding claims, characterized in that it comprises a data preprocessing step (E5), implemented by the avionics computer (5) before the data processing step (E2), to calculate an input signal (U) for the control law from the input data (A) acquired in the acquisition step (El), using the following equation: in which: U is the input signal of the control law; N is the number of sensors (4); ai is an acceleration measured by one of said sensors (4); and ar is an acceleration measured by one of said sensors (4) configured to measure an acceleration linked to a rigid movement of the aircraft (AC).

4. Method according to any one of the preceding claims, characterized in that it comprises a step (E6) of post-processing data, implemented by the avionics computer (5) after the step (E2) of processing data, to apply at least one of the following limitations to the control order determined by the control law: a limitation of deployment amplitude (L1), a limitation of deployment speed (L2), a delay (D).

5. Method according to any one of the preceding claims, characterized in that the control law comprises at least one filtering chain (H).

6. Method according to any one of the preceding claims, characterized in that it comprises a monitoring step (E7), implemented by a monitoring unit, for recording the input data (A) measured by the plurality of sensors (4) over time, for comparing said input data (A) with predetermined threshold values, and, if at least one of said input data (A) is greater than said corresponding predetermined threshold value, for inhibiting the deployment of the control surface (3) of the aircraft (AC) controlled by the control order (C).

7. Device for generating a flutter phenomenon on at least one part of an aircraft, characterized in that it comprises at least: - a plurality of sensors (4) arranged on the aircraft (AC) and configured to measure input data (A); - an avionics computer (5) arranged on the aircraft (AC) and configured to determine, from said input data (A), at least one control order (C) for at least one control surface (3) of the aircraft (AC), the control order (C) being determined using a control law comprising at least one adjustable parameter corresponding to a gain value (G), said control law being configured to obtain a control order (C) making it possible to generate a flutter mode on at least one part (2) of the aircraft (AC); and - a control system (6) configured to control the control surface (3) of the aircraft (AC) so as to generate said floating mode using the control order (C) determined by the avionics computer (5).

8. Device according to claim 7, characterized in that it comprises an adjustment unit (7) comprising an interactive interface making it possible to adjust the adjustable parameter(s) of the control law in real time, so as to obtain a control order (C) of the control surface (3) of the aircraft (AC) making it possible to generate a predetermined floating mode.

9. Device according to one of claims 7 and 8, characterized in that it comprises a monitoring unit (8) configured to record the input data (A) measured by the plurality of sensors (4) over time, to compare said input data (A) with predetermined threshold values, and, if at least one of said input data (A) is greater than said corresponding predetermined threshold value, to inhibit the deployment of the control surface (3) of the aircraft (AC) controlled by the control order (C).

10. Aircraft, characterized in that it comprises at least one device (1) according to any one of claims 7 to 9.

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