Attitude-Elasticity-Aerodynamic Coupled Flutter Control Method for Flying Wing Aircraft

Abstract

The present invention discloses a method for attitude-elastic-aerodynamic coupled flutter control of an aircraft with a flying wing layout, which relates to technical fields such as mechanics, aeroelasticity, and control. The control method includes: integrating the Lagrangian equation in the general body axis system to establish a full-coupling dynamic model, and through small perturbation linearization, establishing a dynamic equation of perturbed motion for flight control system design; based on system identification flight tests, correcting the dynamic equation of perturbed motion; designing an attitude-elastic-aerodynamic coupled flutter suppression controller based on traditional or modern control theory and embedding it into the flight control system; verifying the effectiveness of the flutter suppression controller through flight tests. The present invention significantly solves the problems that existing numerical analysis and wind tunnel experiments cannot accurately reflect the complex coupling effects during flight, cannot effectively suppress the occurrence of flutter during flight, and cannot expand the flight envelope, etc., and can effectively improve the flight safety and stability of aircraft with a flying wing layout.

Description

Technical Field

[0001] This application relates to technical fields such as mechanics, aeroelasticity, and control, specifically to the attitude-elastic-aerodynamic coupled flutter control method for flying wing layout aircraft. Background Technique

[0002] Flying wing layout aircraft are widely used in modern aviation due to their high lift-to-drag ratio and stealth performance. However, due to the lack of traditional tail structures, flying wing layout aircraft have complex aeroelastic characteristics. In particular, during flight, they are prone to a rigid-flexible coupled flutter phenomenon caused by the coupling of flight mechanics modes and aeroelastic modes, which may lead to aircraft structural damage or even disintegration, seriously affecting flight safety and aircraft performance.

[0003] Traditional flutter suppression methods mainly rely on passive suppression techniques, such as increasing structural stiffness or installing dampers. However, these methods will increase the aircraft weight and reduce the aerodynamic performance. In recent years, active suppression techniques have gradually become a research hotspot, which suppress the occurrence of flutter through real-time control surface deflection. However, existing active suppression methods are mostly based on numerical simulation or ground tests, and it is difficult to comprehensively reflect the complex aeroelastic coupling effects in actual flight, and there is a lack of verification under actual flight conditions.

[0004] Therefore, researching the rigid-flexible coupled flutter suppression technology for flying wing layout aircraft, carrying out accurate modeling of the aeroservoelastic system, developing highly robust rigid-flexible coupled flutter suppression methods, effectively suppressing the occurrence of rigid-flexible coupled flutter under flight conditions, and expanding the flight envelope of flying wing layout aircraft have always been technical problems to be solved by those skilled in the art. Summary of the Invention

[0005] The purpose of this application is to provide an attitude-elastic-aerodynamic coupled flutter control method for flying wing layout aircraft. Through aeroservoelastic system modeling, attitude-elastic-aerodynamic coupled flutter suppression controller design, flight control system hardware development, and flutter suppression flight tests, it can effectively suppress the rigid-flexible coupled flutter of flying wing layout aircraft within the flight envelope and expand the flutter boundary.

[0006] The technical solution of this application is as follows:

[0007] The attitude-elastic-aerodynamic coupled flutter control method for flying wing layout aircraft includes the following steps:

[0008] Step 1: Synthesize the Lagrangian equation in the general body axis system, consider the high-fidelity finite element discrete model of the flying wing layout aircraft, the high-precision unsteady aerodynamic force model, etc., and establish a full-scale attitude-elastic-aerodynamic coupled dynamic model considering inertial coupling and aerodynamic coupling effects;

[0009] Step 2: Under typical operating conditions, perform static aeroelastic trim for the flying wing aircraft with pitch angle and elevator commands. Linearize the full-state attitude-elasticity-aerodynamics coupling dynamics model under this configuration for small perturbations, and establish the perturbation motion dynamics equation that can be used for flight control system design;

[0010] Step 3: Develop an aeroservoelastic system identification module and embed it into the flight control system hardware. Make the flying wing aircraft fly at a constant speed in a straight line along a specified route, activate the control surface sweep excitation and record the wing tip acceleration and airframe pitch rate responses. Based on the control surface commands, accelerations, and pitch rate responses obtained from the flight test, identify the frequency-damping characteristics of the aeroservoelastic system through the dynamic mode decomposition method with control (DMDc), and modify the frequency, damping, and phase of the established perturbation motion dynamics equation according to the results.

[0011] Step 4: Based on the modified perturbation motion dynamics equation, use traditional or modern control theories, such as proportional-derivative-integral control, robust control, or optimal control theory, to design an attitude-elasticity-aerodynamics coupling flutter suppression controller. The obtained flutter suppression controller is discretized at a sampling frequency of 300 Hz and then embedded into the flight control system hardware;

[0012] Step 5: Conduct a flutter suppression flight test. With the flying wing aircraft in the closed-loop of the flight control and flutter suppression controllers, increase the flight speed to the flutter speed and observe whether the aircraft becomes unstable. If the aircraft can fly stably, verify the effectiveness of the elastic coupling flutter suppression controller by switching the flutter suppression system switch.

[0013] Furthermore, for the modeling of the aeroservoelastic system, it is required to establish a full-state attitude-elasticity-aerodynamics coupling dynamics model considering inertial coupling and aerodynamic coupling effects based on the Lagrangian equation of the flying wing aircraft in the general body axis system:

[0014] ;

[0015] where , , , are the translational velocity, rotational angular velocity, modal displacement, and modal velocity of the aircraft respectively, , , are the momenta corresponding to the translational velocity, rotational angular velocity, and modal velocity of the aircraft respectively, , , are the generalized forces corresponding to the translational motion, rotational motion, and elastic deformation respectively, , , are the finite element nodes of the aircraft structure at ( is the total number of nodes), the modal shape matrix, the distributed mass, and the combined velocity, is the stiffness matrix of the structure, is the damping matrix of the structure, is the cross product matrix of the vector , refers to the derivative of the variable with respect to time, refers to the transpose of the matrix .

[0016] Furthermore, for the modeling of the aeroservoelastic system, it is required that the full-state attitude-elastic-aerodynamic coupling dynamics model performs static aeroelastic trim for the flying wing aircraft in terms of the pitch angle and elevator command under typical operating conditions. Under this configuration, the full-state nonlinear rigid-elastic coupling dynamics model is linearized by small perturbations to establish the perturbation motion dynamics equation that can be used for the design of the flight control system:

[0017] ;

[0018] wherein, , are the translational velocity and rotational angular velocity of the aircraft under the reference motion respectively, , are the momenta corresponding to the translational velocity and rotational angular velocity of the aircraft under the reference motion respectively, is the combined velocity at the finite element node of the aircraft structure under the reference motion; , , , are the translational velocity, rotational angular velocity, modal displacement, and modal velocity of the aircraft under the perturbation motion respectively, , , are the momenta corresponding to the translational velocity, rotational angular velocity, and modal velocity of the aircraft under the perturbation motion respectively, , , are the derivatives of each momentum of the aircraft under the perturbation motion respectively, , , are the generalized forces corresponding to the translational motion, rotational motion, and elastic deformation under the perturbation motion respectively, is the combined velocity at the finite element node of the aircraft structure under the perturbation motion.

[0019] Furthermore, for the modeling of the aeroservoelastic system, a constant-speed flight test is required to be carried out under the condition of a speed lower than the open-loop flutter speed of the aircraft. The wingtip acceleration and the body pitch rate response are recorded throughout the flight. After the aircraft reaches a stable condition of constant speed and straight flight, the control surface is swept-frequency excited by turning on the switch. The swept-frequency excitation signal of the control surface covering the required structural modal frequency range is:

[0020] ;

[0021] where is the control surface deflection command, is the control surface deflection amplitude, is the starting frequency of the swept frequency, is the cut-off frequency of the swept frequency, is the duration of the swept-frequency signal.

[0022] Furthermore, based on the control surface input command and the response output data under flight conditions, the transfer function of the aeroservoelastic system obtained from the flight test is:

[0023] ;

[0024] where is the transfer function of the subcritical aeroservoelastic system, is the cross-power spectrum of the output signal y with respect to the input signal , is the auto-power spectrum of the input signal .

[0025] Furthermore, for the modeling of the aeroservoelastic system, based on the control surface command and the acceleration response obtained from the flight test, the frequency-damping characteristics of the aeroservoelastic system are identified by the dynamic mode decomposition method with control (DMDc), and the frequency, damping, and phase of the established disturbance motion dynamics equation are corrected according to the results.

[0026] Furthermore, for the design of the attitude-elastic-aerodynamic coupled flutter suppression controller, a controller is designed based on the modified transfer function of the aeroservoelastic system. The wingtip acceleration and the body pitch rate response are used as the input signals of the controller, and the control surface deflection command is used as the output signal of the controller. The traditional or modern control theory, such as proportional-derivative-integral control, robust control, or optimal control theory, is adopted to design the attitude-elastic-aerodynamic coupled flutter suppression controller.

[0027] Further, for the development of the flight control system hardware, it will be developed based on the open-source ArduPilot flight control program. The aeroservoelastic system identification module therein includes the function implementation of the control surface sweep excitation and the opening and closing of the control surface sweep excitation controlled by a switch. The high-frequency data acquisition and processing module therein includes the data acquisition and recording of signals such as wingtip acceleration, airframe pitch rate, and control surface control commands at a sampling frequency of 300 Hz. The attitude-elastic-aerodynamic coupling flutter suppression module therein includes the implementation of the flutter suppression controller discretized at a sampling frequency of 300 Hz and the opening and closing of the flutter suppression controller controlled by a switch.

[0028] Further, for the flutter suppression flight test, it is required that the flying wing layout aircraft fly along a specified route under the condition that the attitude control and flutter control are both closed-loop, accelerate the aircraft beyond the flutter speed, observe whether the aircraft flight is stable, and verify the effectiveness of the attitude-elastic-aerodynamic coupling flutter suppression controller through the switching of the flutter suppression system switch under the condition of being greater than the flutter speed and enabling stable flight.

[0029] The beneficial effects of this application are as follows: By establishing a full-scale attitude-elastic-aerodynamic coupling dynamics model and correcting the model through aeroservoelastic flight tests, it is possible to more accurately determine the attitude-elastic-aerodynamic coupling flutter characteristics of the flying wing layout aircraft and design a more effective flutter suppression controller; Secondly, through the attitude-elastic-aerodynamic coupling flutter suppression flight test, it is possible to verify the effectiveness and reliability of the controller under actual flight conditions, further improving the flight performance and flight envelope of the aircraft; Finally, the flight control system is developed based on the open-source ArduPilot code, with high flexibility and scalability. Brief Description of the Drawings

[0030] Figure 1 is the technical roadmap of the attitude-elastic-aerodynamic coupling flutter control method for the flying wing layout aircraft of this application;

[0031] Figure 2 is the schematic diagram of the control surface and sensor layout of the flying wing layout aircraft in an embodiment of this application;

[0032] Figure 3 is the block diagram of the attitude-elastic-aerodynamic coupling flutter suppression control method based on robust control theory and considering flight control in an embodiment of this application;

[0033] Figure 4 is the result diagram of the flight test for verifying the effectiveness of the attitude-elastic-aerodynamic coupling flutter suppression controller in an embodiment of this application. Detailed Embodiment

[0034] To clearly illustrate the technical features of this application, the following provides a detailed elaboration of this application through specific implementation manners and in conjunction with its accompanying drawings.

[0035] This application is a method for attitude-elastic-aerodynamic coupled flutter control of a flying wing layout aircraft. The control method provided in this embodiment is applied to the establishment of an attitude-elastic-aerodynamic coupled model of a flying wing layout aircraft and the design of a flutter suppression controller, and actively suppresses attitude-elastic-aerodynamic coupled flutter during flight to expand the aircraft's flutter boundary and improve flight safety.

[0036] The control method provided in this application includes aeroservoelastic system modeling, attitude-elastic-aerodynamic coupled flutter suppression controller design, flight control system hardware development, and flutter suppression flight tests.

[0037] The aeroservoelastic system modeling is based on the Lagrangian equation in the general body axis system, and the aeroservoelastic system model is corrected using flight test data below the open-loop flutter speed. Based on traditional or modern control theories, such as proportional-derivative-integral control, robust control, or optimal control theory, the attitude-elastic-aerodynamic coupled flutter suppression controller is designed, and the flight control system hardware development is carried out. The attitude-elastic-aerodynamic coupled flutter suppression control law is embedded in the flight control system, and the effectiveness of the flutter suppression controller is verified through the flutter suppression flight test.

[0038] Its technical route is as Figure 1 shown. In combination with Figure 2 the flying wing layout aircraft described above, research on attitude-elastic-aerodynamic coupled flutter suppression technology is carried out. The outermost ailerons of the aircraft are the flutter suppression control surfaces, accelerometers are arranged at the wingtips to monitor the acceleration response in the vertical direction of the wingtips, and gyroscopes monitor the pitch rate of the airframe. The open-loop rigid-elastic coupled flutter speed of this aircraft is 39.2 m / s, and the flutter frequency is 3.40 Hz.

[0039] Step 1: Establish a high-fidelity finite element discrete model of a flying wing layout aircraft, and conduct a ground resonance test on the flying wing layout aircraft. The finite element model is corrected using the modal information obtained from the test. Combining the Lagrangian equation in the general body axis system and considering a high-precision unsteady aerodynamic model, a full-scale attitude-elastic-aerodynamic coupled dynamics model considering inertial coupling and aerodynamic coupling effects is established:

[0040] ;

[0041] where , , , are the translational velocity, rotational angular velocity, modal displacement, and modal velocity of the aircraft, respectively. , , are the momenta corresponding to the translational velocity, rotational angular velocity, and modal velocity of the aircraft respectively. is the finite element node of the aircraft structure at ( is the total number of nodes), the resultant velocity; , , are the generalized forces corresponding to translational motion, rotational motion, and elastic deformation respectively, obtained by integrating the aerodynamic force model, gravity, and engine thrust based on the high-fidelity aerodynamic shape calculation of the flying wing layout aircraft; , are the modal shape matrix and distributed mass at the finite element node of the aircraft structure respectively, is the stiffness matrix of the structure, is the damping matrix of the structure, obtained by performing modal analysis on the finite element model of the flying wing layout aircraft.

[0042] Step 2: Linearize the full-state attitude-elastic-aerodynamic coupling dynamics model under the static aeroelastic trim configuration to establish a perturbation motion dynamics equation that can be used for flight control system design. The solution provided in this embodiment is to select 36.0 m / s as the typical flight condition at a flutter speed lower than 39.2 m / s, perform static aeroelastic trim on the flying wing layout aircraft with the body pitch angle and inner elevator command, and linearize the full-state attitude-elastic-aerodynamic coupling dynamics model under this configuration to establish a perturbation motion dynamics equation that can be used for flight control system design:

[0043] ;

[0044] The structural mass distribution, modal shape, etc. in the above formula are obtained based on its finite element model, each generalized force is established based on the aerodynamic force, gravity, and engine thrust of the flying wing layout aircraft, and the meanings of each variable are the same as those in the previous formula. refers to the variable under the reference motion , refers to the variable under the perturbation motion . The zero-order variables related to the reference motion are obtained from the static aeroelastic trim solution under the given working conditions of the flying wing layout aircraft, and further a state-space model with each perturbation quantity as the state variable can be obtained.

[0045] Step 3: Develop an aeroservoelastic system identification module and embed it into the flight control system hardware. Make the flying wing aircraft fly in a straight line at a constant speed along a specified route, activate the swept-frequency excitation of the control surface and record the wingtip acceleration response. Obtain a dynamic model based on the test data through system identification methods, obtain the transfer function of the aeroservoelastic system identification, and correct the dynamic equation of the disturbance motion for the design of the flight control system according to the transfer function of the aeroservoelastic system identification.

[0046] In this embodiment, the flight condition of the flight test is 36.0 m / s. Conduct a swept-frequency flight test for aeroservoelastic system identification, develop an aeroservoelastic system identification module and embed it into the flight control system hardware. Set the straight-line flight distance of the aircraft to 1200 m, record the body pitch rate and wingtip acceleration response throughout the flight, and activate the swept-frequency excitation of the control surface through a switch. The swept-frequency excitation signal of the control surface is:

[0047] ;

[0048] where is the control surface deflection command, the control surface deflection amplitude , the starting frequency of the swept frequency , the cut-off frequency of the swept frequency , and the duration of the swept-frequency signal .

[0049] Preferably, according to the control surface input command and the output data of the wingtip vertical acceleration response under the flight condition of 36.0 m / s, identify the frequency damping characteristics of the aeroservoelastic system through the dynamic mode decomposition method with control DMDc, and correct the frequency, damping and phase of the established dynamic equation of the disturbance motion based on this result. The corrected transfer function of the aeroservoelastic system is:

[0050] ;

[0051] Step 4: Based on the corrected dynamic equation of the disturbance motion, design an attitude-elastic-aerodynamic coupled flutter suppression controller using traditional or modern control theory. After obtaining the discretized transfer function of the controller, embed it into the flight control system hardware.

[0052] Based on the corrected mathematical model of the aeroservoelastic system in the flight test under the flight condition of 36.0 m / s, using the wingtip acceleration response and the body pitch rate as the input signals of the controller, and the deflection command of the control surface as the output signal of the controller, in this embodiment, the robust control theory is used to design an attitude-elastic-aerodynamic coupled flutter suppression controller. The controller design block diagram is as Figure 3As shown, where r_dist, u_dist, and y_dist are the inputs of the external signal to the pilot input, aileron control input, and aircraft response on the generalized controlled plant respectively, and W r , W d , and W n are the weighting parameters corresponding to each input respectively; z1, z2, and z3 are the outputs of the aileron control input, aircraft response, and flutter mode speed on the generalized controlled plant respectively, and W u , W y , and W z are the weighting parameters corresponding to each output respectively. According to this robust controller design method, the transfer function of the designed attitude-elastic-aerodynamic coupling flutter suppression controller is:

[0053] ;

[0054] Preferably, the designed attitude-elastic-aerodynamic coupling flutter suppression controller is discretized at a sampling frequency of 300 Hz and then embedded into the flight control system.

[0055] Step Five: Conduct an attitude-elastic-aerodynamic coupling flutter suppression flight test. Select 40.0 m / s greater than the flutter speed as the flight condition for the flight test. The test data is as Figure 4 shown. During the whole takeoff process of the flying wing layout aircraft, keep the attitude control and flutter suppression controller turned on for the flight test. After 196 s, when the flutter suppression controller is turned off, the wing tip acceleration response and the body pitch rate of the aircraft will quickly diverge and flutter will occur. After 198 s, when the flutter suppression controller is turned on again, the wing tip acceleration response and the body pitch rate response are quickly suppressed, verifying the effectiveness of the attitude-elastic-aerodynamic coupling flutter suppression controller and the reliability of the attitude-elastic-aerodynamic coupling flutter control method for the flying wing layout aircraft.

[0056] The above embodiments are only for illustrating the technical concept and features of the present application, and their purpose is to enable those skilled in the art to understand the content of the present application and implement it accordingly. It should not be used to limit the protection scope of the present application. Any equivalent changes or modifications made according to the spirit of the present application should be covered within the protection scope of the present application.

Claims

1. A method for attitude-elastic-aerodynamic coupling flutter control of a flying wing layout aircraft, characterized in that It includes the following steps: Step 1: Establish a full - scale attitude - elasticity - aerodynamic coupling dynamic model considering inertial coupling and aerodynamic coupling effects based on the Lagrangian equation in the general body - axis system; Step 2: Linearize the small perturbation of the full - scale attitude - elasticity - aerodynamic coupling dynamic model in the static aeroelastic trim configuration to establish a dynamic equation of perturbation motion that can be used for flight control system design; Step 3: Develop an aerodynamic servo - elastic system identification module and embed it in the flight control system hardware. Make the flying - wing aircraft fly at a constant speed in a straight line along a specified route, turn on the control - surface sweep - frequency excitation and record the wing - tip acceleration response. Obtain a dynamic model based on experimental data through system identification methods, get the transfer function of the aerodynamic servo - elastic system identification, and correct the dynamic equation of perturbation motion for flight control system design according to the transfer function of the aerodynamic servo - elastic system identification; Step 4: Based on the corrected dynamic equation of perturbation motion, design an attitude - elasticity - aerodynamic coupling flutter suppression controller using traditional or modern control theory. After obtaining the discretized transfer function of the controller, embed it in the flight control system hardware; Step 5: Conduct a flutter suppression flight test. When the flying - wing aircraft is in a closed - loop state of both attitude control and flutter control, observe whether the aircraft will become unstable when the flight speed reaches the flutter speed, and verify the effectiveness of the controller when the aircraft has flutter through the fast - switch of the controller.

2. The attitude-elastic-aerodynamic coupled flutter control method for a flying wing layout aircraft according to claim 1, characterized in that The full - scale attitude - elasticity - aerodynamic coupling dynamic model considering inertial coupling and aerodynamic coupling effects is: ； Among them 、 、 、 are the translational velocity, rotational angular velocity, modal displacement, and modal velocity of the aircraft respectively, 、 、 are the momenta corresponding to the translational velocity, rotational angular velocity, and modal velocity of the aircraft respectively, 、 、 are the generalized forces corresponding to translational motion, rotational motion, and elastic deformation respectively, 、 、 are the modal shape matrix, distributed mass, and combined velocity at the finite element node of the aircraft structure respectively, is the stiffness matrix of the structure, is the damping matrix of the structure, is the cross product matrix of the vector , denotes the derivative of the variable with respect to time, denotes the transpose of the matrix .

3. The attitude-elasticity-aerodynamic coupling flutter control method for a flying wing layout aircraft according to claim 2, wherein, The specific meaning of linearizing the small perturbation of the full - scale attitude - elasticity - aerodynamic coupling dynamic model in the static aeroelastic trim configuration is: Divide the actual motion of the aircraft into reference motion and perturbation motion. Mark the reference motion with superscript 0 and the perturbation motion with superscript 1. Rewrite each parameter in the full - scale attitude - elasticity - aerodynamic coupling dynamic model considering inertial coupling and aerodynamic coupling effects as: ； Carry out static aeroelastic trim of the flying - wing aircraft with pitch angle and elevator command, solve the attitude angle, control - surface deflection angle, and structural elastic deformation trim amount, obtain the static aeroelastic trim configuration required for the aircraft to maintain the current maneuvering flight, and then linearize the small perturbation of the full - scale non - linear rigid - elastic coupling dynamic model in the current configuration: ； Among them, and are the translational velocity and rotational angular velocity under the reference motion of the aircraft respectively, and are the momenta corresponding to the translational velocity and rotational angular velocity under the reference motion of the aircraft respectively, is the resultant velocity at the finite element node of the aircraft structure under the reference motion; 、 、 、 are the translational velocity, rotational angular velocity, modal displacement, and modal velocity under the perturbation motion of the aircraft respectively, 、 、 are the momenta corresponding to the translational velocity, rotational angular velocity, and modal velocity under the perturbation motion of the aircraft respectively, 、 、 are the derivatives of each momentum under the perturbation motion of the aircraft respectively, 、 、 are the generalized forces corresponding to the translational motion, rotational motion, and elastic deformation under the perturbation motion respectively, is the resultant velocity at the finite element node of the aircraft structure under the perturbation motion.

4. The attitude-elasticity-aerodynamic coupling flutter control method for a flying wing layout aircraft according to claim 1, characterized in that In the flutter suppression flight test, conduct a constant - speed flight test under the condition of lower than the flutter speed of the aircraft. The output signal is the wing - tip acceleration response, and the control - surface excitation input signal covering the required structural modal frequency range is: ； wherein is the rudder surface deflection command, is the amplitude of the rudder surface deflection, is the start frequency of the frequency sweep, is the stop frequency of the frequency sweep, is the duration of the frequency sweep signal.

5. The attitude-elasticity-aerodynamic coupling flutter control method for a flying wing layout aircraft according to claim 1, characterized in that The transfer function of the aerodynamic servo - elastic system identification is: ； wherein, is the transfer function of the aeroelastic servo system, is the cross-power spectrum of the output signal y with respect to the input signal , is the auto-power spectrum of the input signal .

6. The attitude-elasticity-aerodynamic coupling flutter control method for the flying wing layout aircraft according to claim 1, characterized in that The specific method of correcting the dynamic equation of perturbation motion for flight control system design according to the transfer function of the aerodynamic servo - elastic system identification is: Based on the control - surface command, acceleration, and pitch - rate response obtained from flight tests, identify the frequency - response function of the aerodynamic servo - elastic system through the dynamic mode decomposition method with control DMDc, and correct the frequency, damping, and phase of the established dynamic equation of perturbation motion according to this result.

7. The attitude-elasticity-aerodynamic coupling flutter control method for a flying wing layout aircraft according to claim 1, characterized in that The transfer function of the attitude - elasticity - aerodynamic coupling flutter suppression controller is: ； Where s is the Laplace operator.

8. The attitude-elastic-aerodynamic coupled flutter control method for a flying wing layout aircraft according to claim 1, characterized in that The hardware development of the flight control system includes an aerodynamic servo-elastic system identification module, a high-frequency data acquisition and processing module, and an attitude-elastic-aerodynamic coupling flutter suppression module based on the open-source ArduPilot flight control program. By switching different switches, the opening and closing of the aerodynamic elastic system identification and the rigid-flexible coupling flutter suppression system can be realized respectively. The wingtip acceleration, body pitch rate, and control command signals are collected, recorded, and saved at a set sampling frequency.

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