Attitude-elastic-pneumatic coupling flutter control method of flying wing layout aircraft

Abstract

The invention discloses an attitude-elasticity-pneumatic coupling flutter control method of a flying wing layout aircraft, and relates to the technical fields of mechanics, aeroelasticity, control and the like. The control method comprises the steps that a Lagrangian equation under a general body shafting is synthesized, a full-quantity coupling dynamic model is established, and a disturbance motion dynamic equation used for flight control system design is established through small disturbance linearization; correcting the disturbance motion kinetic equation based on a system identification flight test; designing an attitude-elastic-pneumatic coupling flutter suppression controller based on a traditional or modern control theory, and embedding the attitude-elastic-pneumatic coupling flutter suppression controller into a flight control system; the effectiveness of the flutter suppression controller is verified through a flight test. The problems that existing numerical analysis and wind tunnel experiments cannot accurately reflect the complex coupling effect in flight, flutter cannot be effectively restrained in flight, and the flight envelope cannot be expanded are solved, and the flight safety and stability of the flying wing layout aircraft can be effectively improved.

Description

Technical Field

[0001] The present application relates to technical fields such as mechanics, aeroelasticity, and control, and specifically to an attitude-elasticity-aerodynamic coupled flutter control method for a flying wing aircraft. Background Art

[0002] Flying wing aircraft are widely used in modern aviation due to their high lift-to-drag ratio and stealth performance. However, due to the lack of a traditional tail structure, flying wing aircraft have complex aeroelastic characteristics. In particular, during flight, a rigid-elastic coupling flutter phenomenon, which is a coupling between flight mechanics mode and aeroelastic mode, is prone to occur, which may cause damage to the aircraft structure 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, but these methods increase aircraft weight and reduce aerodynamic performance. In recent years, active suppression technology has gradually become a research hotspot, which suppresses flutter by real-time control surface deflection. However, existing active suppression methods are mostly based on numerical simulation or ground tests, which are difficult to fully reflect the complex aeroelastic coupling effects in actual flight, and lack verification under actual flight conditions.

[0004] Therefore, it has always been a technical challenge to be solved by technical personnel in this field to study the rigid-elastic coupling flutter suppression technology of flying wing layout aircraft, carry out accurate aerodynamic servoelastic system modeling, develop highly robust rigid-elastic coupling flutter suppression methods, effectively suppress the occurrence of rigid-elastic coupling flutter under flight conditions, and expand the flight envelope of flying wing layout aircraft. Summary of the invention

[0005] The purpose of the present application is to provide an attitude-elastic-aerodynamic coupled flutter control method for a flying wing aircraft. By modeling a pneumatic servoelastic system, designing an attitude-elastic-aerodynamic coupled flutter suppression controller, developing flight control system hardware, and conducting flutter suppression flight tests, the method can effectively suppress the rigid-elastic coupled flutter of a flying wing aircraft within the flight envelope and expand the flutter boundary.

[0006] The technical solution of this application is as follows: The attitude-elastic-aerodynamic coupled flutter control method for a flying wing configuration aircraft comprises the following steps: Step 1: Integrate the Lagrangian equations under the general body axis system, consider the high-fidelity finite element discrete model of the flying wing aircraft, the high-precision unsteady aerodynamic model, etc., and establish a full-scale attitude-elastic-aerodynamic coupling dynamics model that considers the inertial coupling and aerodynamic coupling effects; Step 2: Under typical working conditions, the static aerodynamic trim of the flying wing aircraft is carried out with pitch angle and elevator commands. Under this configuration, the full attitude-elastic-aerodynamic coupling dynamic model is linearized with small disturbances, and the perturbation motion dynamics equation that can be used for the design of the flight control system is established. Step 3: Develop an aero-servoelastic system identification module and embed it into the flight control system hardware, so that the flying-wing aircraft can fly in a straight line at a constant speed along the specified route, turn on the sweep frequency excitation of the rudder surface and record the wingtip acceleration and the pitch rate response of the fuselage. Based on the rudder instructions, acceleration and pitch rate responses obtained from the flight test, the frequency damping characteristics of the aero-servoelastic system are identified through the controlled dynamic mode decomposition method (DMDc), and the frequency, damping and phase of the established perturbation motion dynamics equations are corrected based on the results.

[0007] Step 4: Based on the modified disturbance motion dynamics equation, traditional or modern control theories, such as proportional-differential-integral control, robust control or optimal control theory, are used to design an attitude-elastic-aerodynamic coupled 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. Step 5: Conduct a flutter suppression flight test. With the flight control and flutter suppression controller in a closed loop, make the flying wing aircraft reach the flutter speed to observe whether the aircraft becomes unstable. If the aircraft can fly stably, verify the effectiveness of the missile-coupled flutter suppression controller by switching the flutter suppression system switch.

[0008] Furthermore, the aerodynamic servoelastic system modeling requires the establishment of a full attitude-elastic-aerodynamic coupling dynamic model that takes into account the inertial coupling and aerodynamic coupling effects for the Lagrangian equations of the flying wing aircraft under the general body axis system: ; in , , , are the aircraft's translational velocity, angular velocity, modal displacement, and modal velocity, respectively. , , are the momentum corresponding to the aircraft's translational velocity, angular velocity, and modal velocity, respectively. , , They are the generalized forces corresponding to translational motion, rotational motion, and elastic deformation, respectively. , , They are respectively the aircraft structure finite element nodes Department ( is the total number of nodes), the modal vibration matrix, the distributed mass, and the total velocity, is the stiffness matrix of the structure, is the damping matrix of the structure, Directivity The cross product matrix of Refers to variables The time derivative of Refers to the matrix The transpose of .

[0009] Furthermore, the aerodynamic servoelastic system modeling requires that the full-volume attitude-elastic-aerodynamic coupling dynamics model be used to carry out static aerodynamic trim of a flying wing aircraft with pitch angle and elevator commands under typical working conditions. Under this configuration, the full-volume nonlinear rigid-elastic coupling dynamics model is linearized with small disturbances to establish a disturbance motion dynamics equation that can be used for flight control system design: ; in, , are respectively the translational velocity and angular velocity of the aircraft under the reference motion, , They are respectively the momentum corresponding to the translational velocity and the angular velocity of the aircraft under the reference motion. The finite element nodes of the aircraft structure under the reference motion The total speed at , , , They are respectively the translational velocity, angular velocity, modal displacement, and modal velocity under the aircraft disturbance motion. , , They are respectively the momentum corresponding to the translational velocity, angular velocity, and modal velocity of the aircraft under disturbance motion, , , are the derivatives of each momentum under the aircraft disturbance motion, , , They are the generalized forces corresponding to translational motion, rotational motion, and elastic deformation under disturbance motion, respectively. The finite element nodes of the aircraft structure under disturbance motion The combined speed at.

[0010] Furthermore, the aerodynamic servoelastic system modeling requires that a constant-speed flight test be carried out under conditions lower than the open-loop flutter speed of the aircraft, and the wingtip acceleration and the body pitch rate response be recorded during the entire flight process. After the aircraft reaches a stable constant-speed and constant-straight flight condition, the rudder sweep frequency excitation is turned on by a switch, and the rudder sweep frequency excitation signal covering the required structural modal frequency range is: ; in is the rudder deflection command, is the rudder surface deflection amplitude, is the sweep start frequency, is the sweep cutoff frequency, is the duration of the sweep signal.

[0011] Furthermore, according to the input command and response output data of the control surface under flight conditions, the transfer function of the aerodynamic servoelastic system obtained based on the flight test is: ; in is the transfer function of the subcritical aeroservoelastic system, The output signal y is the input signal The cross power spectrum of For input signal The autopower spectrum of .

[0012] Furthermore, the aero-servoelastic system modeling is based on the control surface instructions and acceleration responses obtained from the flight test, and the frequency damping characteristics of the aero-servoelastic system are identified through the controlled dynamic mode decomposition method (DMDc), and the frequency, damping and phase of the established disturbance motion dynamics equations are corrected based on the results.

[0013] Furthermore, the attitude-elasticity-aerodynamic coupling flutter suppression controller is designed based on a modified aerodynamic servoelastic system transfer function, with wingtip acceleration and body pitch rate response as input signals of the controller, and the deflection command of the rudder surface as the output signal of the controller. Traditional or modern control theories, such as proportional-differential-integral control, robust control or optimal control theory, are used to design the attitude-elasticity-aerodynamic coupling flutter suppression controller.

[0014] Furthermore, the flight control system hardware development will be developed based on the ArduPilot open source flight control program, wherein the aerodynamic servoelastic system identification module includes the function implementation of the rudder sweep frequency excitation and the start and end of the rudder sweep frequency excitation controlled by a switch, wherein the high-frequency data acquisition and processing module includes data acquisition and record preservation of signals such as wingtip acceleration, body pitch rate and rudder control instructions at a sampling frequency of 300 Hz, and wherein the attitude-elastic-aerodynamic coupled flutter suppression module includes the implementation of a flutter suppression controller discretized at a sampling frequency of 300 Hz and the start and end of the flutter suppression controller controlled by a switch.

[0015] Furthermore, the flutter suppression flight test requires the flying wing aircraft to fly along a specified route with attitude control and flutter control in closed loop at the same time, so that the aircraft is accelerated to exceed the flutter speed, and whether the aircraft flight is stable is observed. Under the condition that the speed is greater than the flutter speed and stable flight can be achieved, the effectiveness of the attitude-elastic-aerodynamic coupled flutter suppression controller is verified by switching the flutter suppression system switch.

[0016] The beneficial effects of the present application are as follows: by establishing a full-scale flight attitude-elasticity-aerodynamic coupling dynamics model and correcting the model through aerodynamic servoelasticity flight tests, the attitude-elasticity-aerodynamic coupling flutter characteristics of a flying wing layout aircraft can be determined more accurately and a more effective flutter suppression controller can be designed; secondly, through attitude-elasticity-aerodynamic coupling flutter suppression flight tests, the effectiveness and reliability of the controller can be verified 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 and has high flexibility and scalability. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 It is the technical roadmap of the attitude-elastic-aerodynamic coupled flutter control method of the flying wing layout aircraft of the present application; Figure 2 This is a schematic diagram of the arrangement of control surfaces and sensors of a flying wing aircraft in accordance with an embodiment of the present application; Figure 3 It is a block diagram of an attitude-elastic-aerodynamic coupled flutter suppression control method based on robust control theory and considering flight control according to an embodiment of the present application; Figure 4 This is a diagram of the results of a flight test to verify the effectiveness of an attitude-elasticity-aerodynamic coupled flutter suppression controller according to an embodiment of the present application. DETAILED DESCRIPTION

[0018] In order to clearly illustrate the technical features of the present application, the present application is described in detail below through specific implementation methods and in conjunction with the accompanying drawings.

[0019] The present application is a method for controlling attitude-elasticity-aerodynamic coupling flutter of a flying wing aircraft. The control method provided in this embodiment is applied to the establishment of an attitude-elasticity-aerodynamic coupling model of a flying wing aircraft and the design of a flutter suppression controller, and active suppression of attitude-elasticity-aerodynamic coupling flutter is achieved during flight to expand the aircraft's flight flutter envelope and improve flight safety.

[0020] The control method provided in the present application includes aerodynamic servoelastic system modeling, attitude-elastic-aerodynamic coupled flutter suppression controller design, flight control system hardware development and flutter suppression flight test.

[0021] The pneumatic servoelastic system modeling is based on the Lagrangian equation under the general body axis system, and the pneumatic servoelastic system model is corrected by flight test data below the open-loop flutter speed. Based on traditional or modern control theories, such as proportional-differential-integral control, robust control or optimal control theory, the attitude-elastic-pneumatic coupled flutter suppression controller is designed, and the flight control system hardware development is carried out. The attitude-elastic-pneumatic 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.

[0022] Its technical route is as follows Figure 1 As shown. Combined Figure 2 The flying wing layout aircraft carries out attitude-elastic-aerodynamic coupling flutter suppression technology research. The outermost aileron of the aircraft is a flutter suppression control surface. Acceleration sensors are arranged at the wingtips to monitor the vertical acceleration response of the wingtips. Gyroscopes monitor the pitch rate of the fuselage. The open-loop rigid-elastic coupling flutter speed of this aircraft is 39.2 m / s, and the flutter frequency is 3.40 Hz.

[0023] Step 1: Establish a high-fidelity finite element discrete model of the flying wing aircraft, and carry out ground resonance tests on the flying wing aircraft. Use the modal information obtained from the test to modify the finite element model. Combine the Lagrangian equations under the general body axis system, consider the high-precision unsteady aerodynamic model, and establish a full-scale attitude-elastic-aerodynamic coupling dynamic model that considers inertial coupling and aerodynamic coupling effects: ; in , , , are the aircraft's translational velocity, angular velocity, modal displacement, and modal velocity, respectively. , , are the momentum corresponding to the aircraft's translational velocity, angular velocity, and modal velocity, respectively. Aircraft structural finite element node Department ( is the total number of nodes); , , They 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 calculated based on the high-fidelity aerodynamic shape of a flying wing aircraft; , They are respectively the aircraft structure finite element nodes The modal shape matrix and distributed mass at is the stiffness matrix of the structure, is the damping matrix of the structure, which is obtained by modal analysis based on the finite element model of the flying wing layout aircraft.

[0024] Step 2: Under the static aerodynamic trim configuration, the full-scale attitude-elastic-aerodynamic coupling dynamic model is linearized with small disturbances 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 a typical flight condition under a flutter speed lower than 39.2 m / s, perform static aerodynamic trim of a flying wing layout aircraft with the body pitch angle and the inner elevator command, and under this configuration, the full-scale attitude-elastic-aerodynamic coupling dynamics model is linearized with small disturbances to establish a perturbation motion dynamics equation that can be used for flight control system design: ; The structural mass distribution, modal vibration shape, etc. in the above formula are obtained based on its finite element model. The generalized forces are established based on the aerodynamic force, gravity, and engine thrust of the flying wing layout aircraft. The meaning of each variable is consistent with the above formula. Refers to the variable under the reference motion , Refers to the variable under the disturbance motion The 0th-order variables related to the reference motion are obtained through the static aerodynamic balance solution of the flying wing aircraft under given working conditions, and further the state space model with each disturbance as the state variable can be obtained.

[0025] Step three, develop an aerodynamic servoelastic system identification module and embed it into the flight control system hardware, make the flying wing layout aircraft fly in a straight line at a constant speed along the specified route, turn on the control surface sweep excitation and record the wingtip acceleration response, obtain a dynamic model based on the test data through the system identification method, obtain the transfer function of the aerodynamic servoelastic system identification, and modify the perturbation motion dynamics equation used for the flight control system design according to the transfer function of the aerodynamic servoelastic system identification.

[0026] In this embodiment, 36.0 m / s is used as the flight condition of the flight test, and a pneumatic servoelastic system identification sweep frequency flight test is carried out. An aerodynamic servoelastic system identification module is developed and embedded in the flight control system hardware. The aircraft is set to fly a straight line at a distance of 1200 m. The body pitch rate and wingtip acceleration response are recorded during the entire flight process. The sweep frequency excitation of the rudder surface is turned on by a switch. The sweep frequency excitation signal of the rudder surface is: ; in is the rudder deflection command, the rudder deflection amplitude , sweep start frequency , sweep cutoff frequency , sweep signal duration .

[0027] Preferably, according to the control surface input command and the wing tip vertical acceleration response output data under the flight condition of 36.0 m / s, the frequency damping characteristics of the aero-servoelastic system are identified by the controlled dynamic mode decomposition method DMDc, and the established disturbance motion dynamics equation is corrected in frequency, damping and phase according to the result, and the corrected aero-servoelastic system transfer function is: ; Step 4: Based on the modified disturbance motion dynamics equation, the attitude-elastic-aerodynamic coupled flutter suppression controller is designed using traditional or modern control theory, and the discretized transfer function of the controller is obtained and embedded into the flight control system hardware.

[0028] Based on the mathematical model of the aerodynamic servoelastic system modified by the flight test under the flight condition of 36.0 m / s, the wingtip acceleration response and the body pitch rate are used as the input signals of the controller, and the deflection command of the control surface is used as the output signal of the controller. In this embodiment, the attitude-elastic-aerodynamic coupling flutter suppression controller is designed using the robust control theory. The controller design block diagram is shown in the figure. Figure 3 As shown, r_dist, u_dist and y_dist are the inputs of external signals to the pilot input, aileron control input and aircraft response on the generalized controlled object, respectively. r , W d and W n are the weighted parameters corresponding to each input; 1 、z 2 and z 3 are the aileron control input, aircraft response and flutter mode velocity output on the generalized controlled object, respectively, u , W y and W z are weighted parameters corresponding to each output. According to this robust controller design method, the transfer function of the designed attitude-elastic-aerodynamic coupled flutter suppression controller is: ; Preferably, the designed attitude-elastic-aerodynamic coupled flutter suppression controller is discretized at a sampling frequency of 300 Hz and then embedded into the flight control system.

[0029] Step 5: Conduct attitude-elastic-aerodynamic coupling flutter suppression flight test. Select a flutter speed greater than 40.0m / s as the flight condition for the flight test. The test data is as follows: Figure 4As shown in the figure, the attitude control and flutter suppression controller of the flying wing aircraft are kept turned on during the whole takeoff process for flight test. After 196 seconds, the flutter suppression controller is turned off, and the wingtip acceleration response and the body pitch rate of the aircraft will diverge rapidly and flutter. After 198 seconds, the flutter suppression controller is turned on again, and the wingtip acceleration response and the body pitch rate response are quickly suppressed, which verifies the effectiveness of the attitude-elasticity-aerodynamic coupling flutter suppression controller and the reliability of the attitude-elasticity-aerodynamic coupling flutter control method of the flying wing aircraft.

[0030] The above embodiments are only for illustrating the technical concept and features of the present application, and their purpose is to enable people familiar with the technology to understand the content of the present application and implement it accordingly, and they cannot 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 included in the protection scope of the present application.

Claims

1. A method for attitude-elastic-aerodynamic coupled flutter control of a flying wing aircraft, characterized in that: The following steps are involved: Step 1: Based on the Lagrangian equation of the general body axis system, a full attitude-elastic-aerodynamic coupling dynamic model considering the inertial coupling and aerodynamic coupling effects is established; Step 2: performing small disturbance linearization on the full-scale attitude-elastic-aerodynamic coupling dynamics model under the static air-bomb trim configuration, and establishing a disturbance motion dynamics equation that can be used for flight control system design; Step 3: Develop an aero-servoelastic system identification module and embed it into the flight control system hardware, make the flying wing layout aircraft fly straight and at a constant speed along the specified route, turn on the sweep frequency excitation of the control surface and record the wing tip acceleration response, obtain a dynamic model based on the test data through a system identification method, obtain the transfer function of the aero-servoelastic system identification, and modify the perturbation motion dynamics equation used for flight control system design according to the transfer function of the aero-servoelastic system identification; Step 4: Based on the modified disturbance motion dynamics equation, the attitude-elastic-aerodynamic coupling flutter suppression controller is designed using traditional or modern control theory, and the controller discretized transfer function is obtained and embedded into the flight control system hardware; Step 5: Conduct a flutter suppression flight test. When the attitude control and flutter control of the flying wing aircraft are closed-loop, the flight speed reaches the flutter speed to observe whether the aircraft becomes unstable. The effectiveness of the controller when flutter occurs is verified by rapid on-off switching of the controller.

2. The attitude-elastic-aerodynamic coupled flutter control method for a flying wing aircraft according to claim 1, characterized in that: The full attitude-elastic-aerodynamic coupling dynamics model considering the inertial coupling and aerodynamic coupling effects is: ； in , , , are the aircraft's translational velocity, angular velocity, modal displacement, and modal velocity, respectively. , , are the momentum corresponding to the aircraft's translational velocity, angular velocity, and modal velocity, respectively. , , They are the generalized forces corresponding to translational motion, rotational motion, and elastic deformation, respectively. , , They are respectively the aircraft structure finite element nodes The modal shape matrix, distributed mass, and combined velocity at is the stiffness matrix of the structure, is the damping matrix of the structure, Directivity The cross product matrix of Refers to variables The time derivative of Refers to the matrix The transpose of .

3. The attitude-elastic-aerodynamic coupled flutter control method for a flying wing aircraft according to claim 2, characterized in that: The small-disturbance linearization of the full-volume attitude-elastic-aerodynamic coupling dynamics model under the static air-bullet trim configuration specifically refers to: The actual motion of the aircraft is divided into reference motion and disturbance motion. The reference motion is marked with superscript 0 and the disturbance motion is marked with superscript 1. The parameters in the full attitude-elastic-aerodynamic coupling dynamics model considering the inertial coupling and aerodynamic coupling effects are rewritten as follows: ； The static aerodynamic trim of the flying wing aircraft is carried out with pitch angle and elevator commands, and the attitude angle, rudder deflection angle, and structural elastic deformation trim are solved to obtain the static aerodynamic trim configuration required for the aircraft to maintain the current maneuvering flight. Then, the full nonlinear rigid-elastic coupling dynamic model is linearized with small disturbances under the current configuration: ； in, , are respectively the translational velocity and angular velocity of the aircraft under the reference motion, , They are respectively the momentum corresponding to the translational velocity and the angular velocity of the aircraft under the reference motion. The finite element nodes of the aircraft structure under the reference motion The total speed at , , , They are respectively the translational velocity, angular velocity, modal displacement, and modal velocity under the aircraft disturbance motion. , , They are respectively the momentum corresponding to the translational velocity, angular velocity, and modal velocity under the aircraft disturbance motion, , , are the derivatives of each momentum under the aircraft disturbance motion, , , They are the generalized forces corresponding to translational motion, rotational motion, and elastic deformation under disturbance motion, respectively. The finite element nodes of the aircraft structure under disturbance motion The combined speed at.

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

5. The attitude-elastic-aerodynamic coupled flutter control method for a flying wing aircraft according to claim 1, characterized in that: The transfer function identified by the aero-servoelastic system is: ； in, is the transfer function of the aeroservoelastic system, The output signal y is the input signal The cross power spectrum of For input signal The autopower spectrum of .

6. The attitude-elastic-aerodynamic coupled flutter control method for a flying wing aircraft according to claim 1, characterized in that: The perturbation motion dynamics equation used for flight control system design is corrected according to the transfer function identified by the aerodynamic servoelastic system. Specifically, based on the control surface instructions, acceleration and pitch rate responses obtained from flight tests, the frequency response function of the aerodynamic servoelastic system is identified by a controlled dynamic mode decomposition method DMDc, and the established perturbation motion dynamics equation is corrected in frequency, damping and phase according to the result.

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

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

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