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 a part of an aircraft. technical field

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

[0002] During flight, an aircraft is subjected to aerodynamic forces that deform its structure. These deformations, coupled with the airflow over the aircraft's external surfaces, can generate undesirable vibration phenomena, particularly in the wings. Indeed, at aircraft speeds exceeding a critical speed, aeroelastic instabilities appear, leading to an oscillation phenomenon known as flutter. This flutter phenomenon can be detrimental to the aircraft, thus limiting its maximum speed. Therefore, it is beneficial to have solutions to counteract the flutter phenomenon.

[0003] One method for suppressing flutter involves modifying the damping properties of the aircraft's (or at least certain parts of it) soft modes in flight by acting on control surfaces, for example, ailerons, flaps, or wing or tail spoilers. This method is developed using simulations and requires evaluation of its effectiveness under real-world conditions. To do this, a test aircraft is brought to conditions close to the flutter limit, i.e., at a speed close to its critical speed, in order to verify whether the damping enhancement capabilities allow for effective flutter suppression.

[0004] However, this method is not entirely satisfactory. Indeed, bringing an aircraft close to its critical speed presents risks. First, it assumes a high degree of confidence in knowing the precise critical speed of a given aircraft. Furthermore, this test method does not allow for the evaluation of the robustness of the flutter suppression, nor, where applicable, for the evaluation of the effectiveness of the self-adaptation algorithms associated with said flutter suppression. Moreover, in the event of unexpected behavior during the test, the only way to return to stable flight conditions is to reduce the aircraft's 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. Description of the invention

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

[0007] According to the invention, the method comprises at least the following sequence of successive steps, implemented repeatedly: - 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 determine, from the input data measured in the measurement step, at least one control command for at least one control surface of the aircraft, the control command 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 command enabling the generation of a flutter mode on at least a part of the aircraft; and - a command step to control the rudder in order to generate said flutter mode using the command 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 aircraft's behavior when it is subjected to a flutter phenomenon, under controlled and safe conditions. Indeed, in the event of an undesirable occurrence during the flutter generation, it is sufficient to stop the implementation of the process to almost instantaneously restore a stable situation for the aircraft.

[0009] Advantageously, the method includes a tuning step, implemented by a tuning unit, to adjust the adjustable parameter(s) of the control law in real time, so as to obtain a control order for the aircraft's rudder enabling the generation of a predetermined flutter mode.

[0010] Furthermore, advantageously, the method includes 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 due to rigid motion of the aircraft.

[0011] In a particular embodiment, the method includes 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 command order determined by the control law: a deployment amplitude limitation, a deployment speed limitation, a delay.

[0012] Moreover, advantageously, the control law includes at least one filtering chain.

[0013] In a particular embodiment, the method includes a monitoring step, implemented by a monitoring unit, 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 commanded by the command order.

[0014] The present invention also relates to a device for generating a fluttering phenomenon on at least a 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 command for at least one control surface of the aircraft, the control command 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 command enabling the generation of a flutter mode on at least a part of the aircraft; and - a control system configured to control said control surface in such a way as to generate said flutter mode using the control order determined by the avionics computer.

[0015] Advantageously, the device includes an adjustment unit comprising an interactive interface allowing adjustment of the adjustable parameter(s) of the control law in real time, so as to obtain a control order of the aircraft rudder enabling the generation of a predetermined flutter mode.

[0016] Furthermore, in a particular embodiment, the device includes 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 system 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 fluttering phenomenon as described above. Brief description of the figures

[0018] The accompanying figures will clearly illustrate how the invention can be implemented. In these figures, identical reference numerals designate similar elements.

[0019] Fig. 1 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 synoptic diagram of a method for generating a floating phenomenon according to a particular embodiment.

[0021] The [Fig.3] is a functional diagram of a control law used by the device for generating a floating phenomenon of the [Fig.1]. Detailed description

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

[0023] The term "float" or "flutter" refers to a vibrational resonance phenomenon caused by aeroelastic coupling that naturally occurs in an aircraft in flight when it reaches or exceeds a critical speed specific to it and dependent on its design. The objective of device 1 is to artificially generate (or induce) the occurrence of flutter in the aircraft AC in flight, without needing to exceed or approach its critical speed too closely.

[0024] Device 1 is particularly well-suited for studying the flutter behavior of an aircraft under controlled conditions. It can, for example, be used to test solutions for eliminating or reducing flutter. This is the purpose of a preferred application, which will be detailed later, in which aircraft AC is a test aircraft used to evaluate the performance of a flutter suppression function. In this preferred application, Device 1 is initially configured to generate flutter on aircraft AC. Then, a system (not described in detail herein), designed to counteract this flutter, is tested to evaluate its ability to effectively counteract the flutter.

[0025] In a particular embodiment, the device 1 is configured to perform measurements on a part 2 of the aircraft AC, on which a flutter is to be generated. From these measurements, the device 1 is capable of determining an order of command for control surfaces 3 of the AC aircraft. The deployment (or movement) of these control surfaces 3 allows the generation of a desired flutter mode on part 2, as explained below.

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

[0027] The term "flutter mode" refers to unstable vibration modes that occur naturally on certain parts of the AC aircraft when it exceeds its critical speed. Furthermore, "occurs naturally" means occurring under normal flight conditions, that is, without altering any element of the AC aircraft in a way that modifies its structural and aerodynamic properties in order to change its vibration response.

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

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

[0030] Furthermore, preferably, the accelerometers 4 are arranged near the control surfaces 3 used to generate the desired flutter mode. This makes it possible to efficiently 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 includes an accelerometer 4R configured to measure an acceleration associated with a particularly rigid motion of the aircraft AC. This acceleration, referred to as the rigid acceleration, corresponds to a rigid dynamic component of the overall motion of the aircraft AC in flight. It is suitable for use as a reference, as explained below. Preferably, the accelerometer 4R is arranged near the center of gravity of the aircraft AC or in an area whose dynamics are consistent with said center of gravity.

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

[0033] Furthermore, in the embodiment of [Fig. 1], 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 part 2, namely the wings 11. However, in other embodiments, the control surfaces 3 may correspond to other conventional control surfaces of the aircraft AC. The control surfaces 3 are chosen according to the application under consideration. Indeed, they must be capable, through their deployment during flight, of modifying the properties of the considered part 2, in particular its aerodynamic damping characteristics, so as to obtain the appearance of the targeted flutter mode.

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

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

[0036] The avionics computer 5 first performs operations on the data measured by the accelerometers 4, referred to as 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 4 accelerometers (without accelerometer 4R); a> is an acceleration measured by one of the accelerometers 4; 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 performs 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 global movement of the aircraft AC in the measurement of the vibration modes on 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 the 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 provides a control command to move the control surfaces 3 in a way that modifies the vibration mode of 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 part 2) are modified, and therefore its vibrational behavior. This leads to a change in the speed at which particular vibration modes appear, notably flutter modes.

[0040] Indeed, thanks to simulations based on previously established aeroelastic models of the AC aircraft, it is possible to identify the vibration modes of part 2 as a function of the AC aircraft's flight conditions, particularly its speed. It is also possible to determine how to modify certain properties of part 2, such as damping, to modify the properties of the resulting vibration mode for a given speed. Knowing this, it is possible to determine how to move the control surfaces 3 to modify these properties of part 2 as desired. The deployment of the control surfaces 3, which is controlled by the command signal, is directly related to the gain value G. Consequently, choosing the gain value G allows us to determine the resulting vibration mode under the considered flight conditions.

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

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

[0043] By way of non-limiting example, a function representing the filtering chain H can be expressed 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 floating mode that we seek to generate, in Hertz (Hz); and Am is a predetermined amplification factor.

[0044] The output signal Y obtained from the transfer function F corresponds to a raw control command that 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 perform data post-processing operations on the signal Y in order to refine it to obtain the final control command C, which is actually used to control the control surfaces 3. The data post-processing operations correspond to limitations denoted D, L1, and L2 in [Fig. 3]. These may include, but are not limited to, a limitation on the deployment amplitude (L1), a limitation on the deployment speed (L2), and / or a delay (D).

[0045] Thus, the control signal C used to control the control surfaces 3 can be adapted to avoid excessive or abrupt deployment of said control surfaces 3, which could be undesirable depending on the application, for example, in terms of the maneuverability of aircraft AC. A delay, on the other hand, can improve the synchronization between the deployment of the control surfaces 3 and the measurements taken by the accelerometers 4.

[0046] By way of non-limiting example, let us consider that the floating mode to be generated has a frequency ω (value given in Hz) and that, without the application of a delay D, there is 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 that allows obtaining a control command C in phase with the input signal U can be expressed in the form of the following equation: 360. /

[0047] Furthermore, as schematically represented in [Fig.1], the device 1 includes 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 schematically represented by dashed lines in [Fig.1], so as to be able to control them and generate the desired flutter mode.

[0048] Thus, thanks to device 1, it is possible to generate a floating mode on the Part 2 of the AC aircraft without needing 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, under controlled and safe conditions. Indeed, in the event of an undesirable occurrence during flutter generation, simply deactivating device 1 allows for an almost instantaneous return to a stable situation.

[0049] Furthermore, as schematically represented in [Fig. 1], the device 1 includes a control unit 7 comprising an interface for adjusting the control law's adjustable parameters. This interface is configured to allow a user to act on the control law's adjustable parameters by modifying them in real time, i.e., during the use of the device 1. For example, this interface may include a screen and input means such as a keyboard or a touchscreen.

[0050] The adjustment unit 7 allows for real-time modification of the flutter mode to be generated on part 2 of the aircraft AC, in particular by acting on the gain value G. It can also allow for adjustment of the performance of the control law, in particular by acting on the parameters of the filter chain H.

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

[0052] Furthermore, as schematically represented in [Fig. 1], the device 1 includes a monitoring unit 8 for monitoring the behavior of part 2 of the aircraft AC during the use of 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 said avionics computer 5 acquires it. The monitoring unit 8 stores the input data A thus received in memory so as to be able to establish and use a history of the measurements.

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

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

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

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

[0057] Step El 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 command for the control surfaces 3 to generate the desired flutter mode on part 2 of the aircraft AC. The data processing performed 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 data preprocessing and postprocessing steps, respectively, implemented by the avionics computer 5. Step E5 precedes step E2 to calculate the input signal U of the control law from the input data A, as described above. Step E6, on the other hand, 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 in order to generate the desired flutter mode using the control order C transmitted by the avionics computer 5.

[0061] The steps of the process P are implemented repetitively and continuously, thus forming an active control loop. Indeed, the vibrational behavior of part 2 allows for the determination of a control command C that moves the control surfaces 3. This deployment modifies the properties of the aircraft AC, which in turn modifies the structural response of part 2, ultimately leading to a new vibrational behavior. The new vibrational behavior is measured by the accelerometers 4, which allows for the determination of a new control command 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 allows the vibratory behavior of said part 2 to evolve until the desired flutter mode is generated.

[0063] Furthermore, the method P includes a tuning step E4, implemented by the tuning unit 7, for adjusting the control law parameters in real time. As detailed above, step E4 allows, in particular, targeting the desired flutter mode and adjusting the performance of the control law. order.

[0064] Furthermore, the method P includes 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 to the threshold values ​​not to be exceeded. If any 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 command signal determined in step E2.

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

[0066] In this application example, the design of the AC aircraft is considered to predispose it to a flutter mode at the wing level 11 at a certain critical speed. The aim is to prevent the occurrence of this flutter mode by implementing the aforementioned flutter suppression system. This would increase the usable speed range of the AC aircraft. Indeed, if the occurrence of the undesirable flutter mode is prevented, the critical speed could be used without risk to the AC aircraft.

[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, the wings 11 of aircraft AC are considered to have external ailerons 9 that influence the flutter pattern to be suppressed. The flutter suppression system is therefore configured to act on these external ailerons 9 in order to counteract the flutter pattern to be suppressed.

[0068] In this preferred application, it is desired to verify that the flutter suppression system is correctly configured by testing it under real-world conditions, without bringing the aircraft AC to a speed close to the critical speed. Therefore, 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 that allow the desired flutter mode to be generated on the wings 11, other than the external ailerons 9 (since these are already used by the flutter suppression system).

[0069] In this example, aircraft AC is assumed to have other ailerons, for example the inner ailerons 10, which can be used by device 1 to generate the desired flutter mode. Therefore, the accelerometers 4 on the wings 11 are arranged appropriately, close to the inner ailerons 10. Furthermore, one configures the control law of device 1 so as to determine a control order C for the inner 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 float mode can be to proceed in stages. First, the theoretical gain value G is determined to produce a control signal C that would generate the target float mode. Next, the control law is configured with a gain value G lower than that required, and then this value is gradually increased until the target float mode is actually generated.

[0071] Once the desired flutter pattern is generated on the wings 11, the flutter suppression system is implemented using the external ailerons 9. This allows the effectiveness of the flutter suppression system to be tested under controlled conditions. Furthermore, its robustness and, if present, its self-adaptive algorithms can also be tested. Indeed, thanks to the control unit 7, the control law can be adjusted during the test flight to vary the characteristics of the flutter pattern generated on the wings 11. This allows the flutter suppression system's ability to adapt to variations in the flutter pattern to be suppressed to be evaluated.

[0072] The device 1 for implementing process P as described above has many advantages. In particular: - it allows a floating phenomenon to be generated on the AC aircraft under controlled conditions without needing to approach 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 effectiveness and robustness of a flutter suppression device for flutter modes with varying characteristics, simply by modifying the adjustable parameters of the control law; and - it is particularly safe since it allows for an easy and quick return to a stable situation in the event of an undesirable event.

Claims

Demands

1. A method for generating a flutter phenomenon on at least one part of an aircraft, characterized in that it comprises at least the following sequence of successive steps, implemented repeatedly: - a measurement step (E1), 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 determine, from the input data (A) measured in the measurement step (E1), at least one control command (C) for at least one control surface (3) of the aircraft (AC), the control command (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 command (C) enabling the generation of a flutter mode on at least one part (2) of the aircraft (AC);and - a control step (E3), implemented by a control system (6), to control the rudder (3) so as to generate said flutter mode using the control order (C) determined in the data processing step (E2).

2. A method according to claim 1, characterized in that it comprises a setting step (E4), implemented by a setting unit (7), to set the adjustable parameter(s) of the control law in real time, so as to obtain a control order (C) for the aircraft (AC) rudder (3) enabling the generation of a predetermined flutter mode.

3. A 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), for calculating an input signal (U) for the control law from the input data (A) acquired in the acquisition step (E1), 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 related to a rigid motion of the aircraft (AC).

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

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

6. A 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) to 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 aircraft (AC) control surface (3) commanded by the command order (C).

7. A device for generating a flutter phenomenon on at least a 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 command (C) for at least one control surface (3) of the aircraft (AC), the control command (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 command (C) enabling the generation of a flutter mode on at least a part (2) of the aircraft (AC); and - a control system (6) configured to control the steers (3) of the aircraft (AC) so as to generate said flutter mode using the command order (C) determined by the avionics computer (5).

8. Device according to claim 7, characterized in that it comprises an adjustment unit (7) including an interactive interface allowing adjustment of the adjustable parameter(s) of the control law in real time, so as to obtain a control order (C) of the aircraft (AC) rudder (3) allowing the generation of a predetermined flutter mode.

9. Device according to any one of claims 7 and 8, characterized in that it comprises a monitoring unit (8) configured to record input data (A) measured by the plurality of sensors (4) over time, to compare said input data (A) to 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 aircraft (AC) control surface (3) commanded by the command 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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