Method, device, equipment and medium for time domain analysis of aircraft body freedom flutter

Abstract

The present application discloses a method, device, equipment and medium for time domain analysis of the flutter of the freedom degree of an aircraft body, and relates to the field of dynamic aeroelastic analysis of an aircraft, including: using an aircraft aerodynamic grid model and an aircraft structural grid model to determine the initial steady state of the aircraft through static balancing; performing motion deformation coupling analysis on the aircraft after the initial disturbance is applied under the initial steady state condition, and determining the displacement of the relative equilibrium position of the nodes on the surface of the aircraft based on the analysis results; updating the aerodynamic grid through displacement, obtaining the aerodynamic force on the surface of the aircraft based on the updated aerodynamic grid, and mapping the aerodynamic force to the structural grid using the aerodynamic structure coupling interface to obtain the aerodynamic load; performing motion deformation coupling iteration and aerodynamic structure coupling iteration based on the aerodynamic load, and performing coupling time advancement when the coupling residual meets the preset conditions, judging the dynamic stability of the aircraft according to the displacement during the time advancement process, and completing the time domain analysis of the flutter of the freedom degree of the aircraft body. High-precision analysis of the aircraft is achieved.

Description

Technical Field

[0001] The present invention relates to aerodynamic analysis of aircraft, and in particular to a method, device, equipment and medium for analyzing the flutter time domain of the freedom degree of an aircraft body. Background Art

[0002] As high-altitude, long-flight aircraft gradually develop towards a large aspect ratio, flexible materials, and flying wing layout, their high flexibility will make the natural frequency of their deformation mode lower and close to the rigid body modal frequency of the entire aircraft, which will lead to a significant coupling relationship between the space motion and flexible deformation of the aircraft. Under the action of aerodynamic forces, this type of aircraft will present a very complex dynamic response, and at a certain speed may induce dynamic instability different from traditional flutter. This dynamic instability phenomenon is called body freedom flutter (abbreviated as body flutter).

[0003] The body flutter problem belongs to the aerodynamic-motion-deformation three-field coupling problem. However, analysis shows that the influence of aerodynamics on motion and deformation is reflected in the form of aerodynamics, while the influence of motion and deformation on aerodynamics can be characterized by changes in the surface of the object. Therefore, the three-field coupling problem can be regarded as a nested layer of coupling with aerodynamics on the basis of motion-deformation coupling. At the level of aerodynamic-structure coupling, body flutter is similar to conventional flutter analysis methods, and the main difference is reflected in the level of motion-deformation coupling.

[0004] As for the coupling of motion and deformation (usually called rigid-flexible coupling in structural dynamics), the modal method is usually used in the existing body flutter analysis methods. It mainly extracts the flexible mode of the aircraft as the generalized coordinates, and uses the Lagrange equation method and the generalized coordinates corresponding to the rigid body motion freedom to form the dynamic equation of the rigid-flexible coupling system of the aircraft. By selecting a suitable flexible modal coordinate system, the coupling terms of the rigid-flexible coupling equation can be greatly simplified to ensure high computational efficiency; however, the coordinate system selected at this time is usually the average axis system, which is different from the body coordinate system in traditional flight mechanics, and is more troublesome in the practical application of technologies such as active flutter suppression.

[0005] At the same time, modal theory is a theoretical method based on linearization. It inevitably ignores some high-order small quantities in the process of constructing flexible modes, which may lead to deviations in the results of such methods when the deformation is large. For example, when the aircraft is deformed, its mass distribution changes, and its moment of inertia relative to each axis will also change. For flexible aircraft with large aspect ratio, this deformation is usually large, which makes the change of moment of inertia more obvious, which leads to obvious deviation between the actual dynamic response and the calculation results based on the modal method. Therefore, how to achieve high-precision analysis of aircraft body vibration under large deformation conditions is an urgent problem to be solved. Summary of the invention

[0006] In view of this, the purpose of the present invention is to provide a method, device, equipment and medium for time domain analysis of aircraft body freedom flutter, which can achieve high-precision analysis of aircraft body flutter under large deformation conditions. The specific scheme is as follows:

[0007] In a first aspect, the present application discloses a method for time domain analysis of aircraft body freedom flutter, comprising:

[0008] Constructing an aircraft aerodynamic mesh model and an aircraft structural mesh model according to the geometric shape, structure, size and material of the aircraft, extracting the aerodynamic-structural coupling interface of the aircraft, and determining the initial steady state of the aircraft by static trim using the aircraft aerodynamic mesh model and the aircraft structural mesh model;

[0009] Applying an initial disturbance to the aircraft, and performing a motion-deformation coupling analysis on the aircraft after the initial disturbance is applied under the initial steady-state condition, and determining the displacement of the surface nodes of the aircraft relative to the equilibrium position based on the corresponding motion-deformation coupling analysis results;

[0010] The aerodynamic grid corresponding to the aircraft is updated by the displacement, and the aerodynamic force on the surface of the aircraft at the current time step is obtained based on the corresponding updated aerodynamic grid, and the aerodynamic force is mapped to the structural grid of the aircraft by using the aerodynamic-structure coupling interface and the coupling data transfer matrix to obtain the aerodynamic load of the aircraft;

[0011] Based on the aerodynamic load, motion deformation coupling iteration and aerodynamic structure coupling iteration are performed, and when the coupling residual meets the preset conditions, coupling time advancement is performed. The dynamic stability of the aircraft is judged according to the displacement of the aircraft surface nodes relative to the equilibrium position during the time advancement process, and the time domain analysis of the flutter of the aircraft body degrees of freedom is completed.

[0012] Optionally, the determining the initial steady state of the aircraft by static trim using the aircraft aerodynamic grid model and the aircraft structural grid model includes:

[0013] Performing aeroelastic analysis based on pre-predicted trim parameters and flight parameters corresponding to the aircraft to obtain lift and pitch moment corresponding to the aircraft; the trim parameters include an elevation angle and an elevator deflection angle of the aircraft;

[0014] Iteratively adjusting the elevation angle based on lift-weight balance using the lift force to obtain an adjusted elevation angle;

[0015] Iteratively adjusting the elevator deflection angle based on the pitching moment balance using the pitching moment to obtain an adjusted elevator deflection angle;

[0016] The initial steady state is determined according to the adjusted pitch angle and the adjusted elevator deflection angle.

[0017] Optionally, determining the displacement of the aircraft surface node relative to the equilibrium position based on the corresponding motion-deformation coupling analysis result includes:

[0018] The displacement of the aircraft surface node relative to the equilibrium position is determined using a preset displacement calculation formula based on the corresponding motion deformation coupling analysis results; the preset displacement calculation formula is:

[0019] ;

[0020] in, is the displacement of the node on the surface of the aircraft relative to the equilibrium position; is the displacement vector of the aircraft relative to the equilibrium position; is the transformation matrix from the body coordinate system to the inertial system; is the vector corresponding to the node in the body coordinate system when the aircraft is in an undeformed state; is the deformation displacement vector of the node relative to the body coordinate system.

[0021] Optionally, updating the aerodynamic grid corresponding to the aircraft by using the displacement includes:

[0022] Based on the topological connection relationship between the grid nodes, the displacement of the aircraft surface nodes relative to the equilibrium position is transmitted to the space grid to update the aerodynamic grid corresponding to the aircraft.

[0023] Optionally, extracting the aerodynamic structure coupling interface of the aircraft includes:

[0024] The outer surface of the aircraft is determined as the aerodynamic-structural coupling interface so as to extract the surface nodes of the aircraft, select representative nodes from the surface nodes, and construct the coupling data transfer matrix based on the representative nodes; the representative nodes include the object surface mesh nodes in the aerodynamic mesh and the surface mesh nodes in the structural mesh.

[0025] Optionally, mapping the aerodynamic force to the structural grid of the aircraft by using the aerodynamic-structural coupling interface and the coupling data transfer matrix to obtain the aerodynamic load of the aircraft includes:

[0026] Extracting the aerodynamic force of each interpolation point of the surface grid in the aerodynamic grid corresponding to the aircraft, and determining the aerodynamic force of each interpolation point of the structural grid corresponding to the aircraft based on the dot product result of the aerodynamic force of each interpolation point of the surface grid and the coupling data transfer matrix;

[0027] The aerodynamic forces of each interpolation point of the structural grid are interpolated to obtain the aerodynamic load of the structural grid corresponding to the aircraft.

[0028] Optionally, performing motion deformation coupling iteration and aerodynamic structure coupling iteration based on the aerodynamic load includes:

[0029] Performing motion-deformation coupling iteration based on the aerodynamic load to obtain updated displacements of the aircraft surface nodes relative to equilibrium positions;

[0030] The updated displacement of the aircraft surface nodes relative to the equilibrium position is used to jump back to the step of updating the aerodynamic mesh corresponding to the aircraft through the displacement, the aerodynamic load of the aircraft is obtained, and the step of performing motion deformation coupling iteration based on the aerodynamic load is jumped again.

[0031] Optionally, judging the dynamic stability of the aircraft according to the displacement of the surface node of the aircraft relative to the equilibrium position during the time advancement process includes:

[0032] Determining the reciprocating motion amplitude of the aircraft according to the displacement of the surface nodes of the aircraft relative to the equilibrium position during the time advancement process;

[0033] If the reciprocating motion amplitude meets the preset reduction rule, it is determined that the aircraft has the dynamic stability;

[0034] If the reciprocating motion amplitude does not meet the preset reduction rule, it is determined that the aircraft does not have the dynamic stability.

[0035] In a second aspect, the present application discloses a device for analyzing the flutter time domain of an aircraft body degree of freedom, comprising:

[0036] An initial steady state determination module is used to construct an aircraft aerodynamic grid model and an aircraft structural grid model according to the geometric shape, structure, size and material of the aircraft, extract the aerodynamic-structural coupling interface of the aircraft, and determine the initial steady state of the aircraft by static trim using the aircraft aerodynamic grid model and the aircraft structural grid model;

[0037] a displacement determination module, used for applying an initial disturbance to the aircraft, performing a motion-deformation coupling analysis on the aircraft after the initial disturbance is applied under the initial steady-state condition, and determining the displacement of the surface nodes of the aircraft relative to the equilibrium position based on the corresponding motion-deformation coupling analysis results;

[0038] an aerodynamic force and load determination module, configured to update the aerodynamic mesh corresponding to the aircraft through the displacement, obtain the aerodynamic force on the surface of the aircraft at the current time step based on the corresponding updated aerodynamic mesh, and map the aerodynamic force to the structural mesh of the aircraft using the aerodynamic structure coupling interface and the coupling data transfer matrix to obtain the aerodynamic load of the aircraft;

[0039] The dynamic stability judgment module is used to perform motion deformation coupling iteration and aerodynamic structure coupling iteration based on the aerodynamic load, and when the coupling residual meets the preset conditions, perform coupling time advancement, judge the dynamic stability of the aircraft according to the displacement of the aircraft surface node relative to the equilibrium position during the time advancement process, and complete the time domain analysis of the flutter of the aircraft body degree of freedom.

[0040] In a third aspect, the present application discloses an electronic device, comprising:

[0041] Memory, used to store computer programs;

[0042] A processor is used to execute the computer program to implement the above-mentioned aircraft body freedom flutter time domain analysis method.

[0043] In a fourth aspect, the present application discloses a computer-readable storage medium for storing a computer program, wherein the computer program, when executed by a processor, implements the aforementioned aircraft body degree of freedom flutter time domain analysis method.

[0044] When performing time domain analysis on the body degree of freedom flutter of an aircraft, the present application first constructs an aircraft aerodynamic mesh model and an aircraft structural mesh model according to the geometric shape, structure, size and material of the aircraft, extracts the aerodynamic-structural coupling interface of the aircraft, and uses the aircraft aerodynamic mesh model and the aircraft structural mesh model to determine the initial steady state of the aircraft through static balancing; then, an initial disturbance is applied to the aircraft, and a motion-deformation coupling analysis is performed on the aircraft after the initial disturbance is applied under the initial steady-state condition, and the displacement of the surface nodes of the aircraft relative to the equilibrium position is determined based on the corresponding motion-deformation coupling analysis results; and then, the aerodynamic-structural coupling interface of the aircraft is extracted by using the aircraft aerodynamic mesh model and the aircraft structural mesh model to determine the initial steady state of the aircraft. The displacement is used to update the aerodynamic grid corresponding to the aircraft, and the aerodynamic force on the surface of the aircraft at the current time step is obtained based on the corresponding updated aerodynamic grid, and the aerodynamic force is mapped to the structural grid of the aircraft using the aerodynamic structure coupling interface and the coupling data transfer matrix to obtain the aerodynamic load of the aircraft; finally, motion deformation coupling iteration and aerodynamic structure coupling iteration are performed based on the aerodynamic load, and when the coupling residual meets the preset conditions, coupling time advancement is performed, and the dynamic stability of the aircraft is judged according to the displacement of the surface nodes of the aircraft relative to the equilibrium position during the time advancement process, and the time domain analysis of the flutter of the aircraft body freedom degree is completed. It can be seen that the present application ensures the solution accuracy of complex configuration flight through motion deformation coupling solution, and can handle the body flutter problem under large deformation conditions; at the same time, combined with the aerodynamic structure coupling solution method, through the dual coupling of aerodynamic structure and motion deformation, high-precision analysis of aircraft body flutter under large deformation conditions is achieved, and the solution efficiency is improved. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the provided drawings without paying creative work.

[0046] Figure 1 A flow chart of a method for analyzing the flutter of a degree of freedom of an aircraft body disclosed in the present application;

[0047] Figure 2 This is a schematic diagram of the appearance of a large aspect ratio flying wing layout aircraft model disclosed in this application;

[0048] Figure 3 A schematic diagram of an aerodynamic grid for a large aspect ratio flying wing layout aircraft disclosed in this application;

[0049] Figure 4 A schematic diagram of a structural grid of a large aspect ratio flying wing layout aircraft disclosed in this application;

[0050] Figure 5 A schematic diagram of the pitch angle of an aircraft disclosed in the present application varying with time;

[0051] Figure 6 A flowchart of a specific method for analyzing the flutter time domain of a specific degree of freedom of an aircraft disclosed in the present application;

[0052] Figure 7 This is a schematic diagram of the structure of a time-domain analysis device for the degree of freedom flutter of an aircraft disclosed in the present application;

[0053] Figure 8 This is a structural diagram of an electronic device disclosed in this application. DETAILED DESCRIPTION

[0054] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0055] In the existing body flutter analysis methods, the modal method is usually used, which mainly extracts the flexible mode of the aircraft as a generalized coordinate, and uses the Lagrange equation method and the generalized coordinates corresponding to the rigid body motion freedom to form the dynamic equation of the rigid-flexible coupling system of the aircraft. By selecting a suitable flexible modal coordinate system, the coupling terms of the rigid-flexible coupling equation can be greatly simplified to ensure high computational efficiency; but the coordinate system selected at this time is usually the average axis system, which is different from the body coordinate system in traditional flight mechanics, and is more troublesome in the practical application of technologies such as active flutter suppression. At the same time, modal theory is a theoretical method based on linearization, which inevitably ignores some high-order small quantities in the process of constructing flexible modes, which may cause deviations in the results of this type of method when the deformation is large. For example, when the aircraft is deformed, its mass distribution changes, and its moment of inertia relative to each axis will also change. For flexible aircraft with a large aspect ratio, this deformation is usually large, which makes the change in the moment of inertia more obvious, which leads to a significant deviation between the actual dynamic response and the calculation results based on the modal method. In order to solve the above technical problems, the present application discloses a method, device, equipment and medium for time-domain analysis of aircraft body freedom flutter, which can achieve high-precision analysis of aircraft body flutter under large deformation conditions.

[0056] See also Figure 1 As shown, an embodiment of the present invention discloses a method for analyzing the flutter of an aircraft body in a time domain, including:

[0057] Step S11, constructing an aircraft aerodynamic grid model and an aircraft structural grid model according to the geometric shape, structure, size and material of the aircraft, extracting the aerodynamic-structural coupling interface of the aircraft, and determining the initial steady state of the aircraft by static balancing using the aircraft aerodynamic grid model and the aircraft structural grid model.

[0058] In this embodiment, before conducting the analysis, it is first necessary to establish a solution model based on the actual shape, structure, size, etc. of the aircraft to be analyzed for subsequent analysis, including establishing an aerodynamic analysis grid model, establishing a structural finite element grid model, and extracting the aerodynamic-structural coupling interface. When establishing the aircraft aerodynamic grid model, an aerodynamic solution grid is established based on the geometric shape and size of the aircraft to be analyzed. The specific establishment process is the same as the conventional CFD (computational fluid dynamics) solution grid. However, in order to adapt to the subsequent aerodynamic grid update caused by the rigid body motion of the aircraft relative equilibrium position, the solution boundary size of the external flow field should be larger than the conventional CFD solution. It is recommended to take 1.5 times the conventional CFD solution boundary size. When establishing the aircraft structural grid model, a structural grid model is established based on the geometric structure, size, material and other properties of the aircraft to be analyzed. The specific grid division, unit attribute definition, material attribute definition and other processes are the same as the conventional finite element grid model. In order to improve the efficiency of subsequent calculations, the origin of the coordinate system should preferably select the center of mass of the aircraft, and the direction of the coordinate axis should preferably select the principal axis of inertia of the aircraft. As shown in the following example Figure 2 As an example, the high aspect ratio flying wing aircraft model shown in the figure is constructed as follows. Figure 3 The aerodynamic mesh model of the aircraft shown is similar to that of Figure 4 The aircraft structure mesh model is shown.

[0059] Then it is necessary to extract the aerodynamic structure coupling interface of the aircraft. The coupling interface is extracted based on the outer surface of the aircraft to be analyzed. The coupling interface is the interface between the fluid and the solid. In the specific aerodynamic structure coupling analysis combined with CFD / FEM (Finite Element Method), the actual extraction is the representative nodes on the interface grid, including the object surface grid nodes in the aerodynamic grid and the surface grid nodes in the structural grid. The two groups of representative nodes can be used to construct the coupling data transfer matrix, and at the same time, they are used as interpolation points to map the transferred physical quantities to the solution grid. The specific extraction process is the same as the conventional aerodynamic structure coupling analysis process. That is, the outer surface of the aircraft is determined as the aerodynamic structure coupling interface, so as to extract the surface nodes of the aircraft, select representative nodes from the surface nodes, and construct the coupling data transfer matrix based on the representative nodes; the representative nodes include the object surface grid nodes in the aerodynamic grid and the surface grid nodes in the structural grid. In a specific embodiment, the coupling interface of the large aspect ratio aircraft is the outer surface of the aircraft, all surface nodes of the model are extracted and representative nodes are selected from them as interpolation points, and the coupling data transfer matrix is ​​constructed using the obtained representative nodes. The point set extracted from the structural model is: , the number of which is ; The point set extracted from the aerodynamic model is: , the number of which is , then the corresponding coupling data transfer matrix is , which is a OK A matrix of columns for quantities from the aerodynamic grid, such as aerodynamic forces , the physical quantity of the interpolation point corresponding to the structured grid can be obtained by dot multiplication of the matrix: ; For the physical quantities from the structured grid, the inverse matrix of this matrix is ​​used to realize the transfer. After the physical quantities are transferred to the interpolation points, the complete physical quantity distribution of the grid interface can be interpolated.

[0060] When the initial steady state of the aircraft is determined by static balancing using the aircraft aerodynamic grid model and the aircraft structural grid model, the present application performs a static aeroelastic analysis based on the pre-predicted balancing parameters and the flight parameters corresponding to the aircraft to obtain the lift and pitch moment corresponding to the aircraft; the balancing parameters include the elevation angle and elevator deflection angle of the aircraft; the lift is used to iteratively adjust the elevation angle based on the lift-to-weight balance to obtain the adjusted elevation angle; the pitch moment is used to iteratively adjust the elevator deflection angle based on the pitch moment balance to obtain the adjusted elevator deflection angle; the initial steady state is determined based on the adjusted elevation angle and the adjusted elevator deflection angle.

[0061] Step S12: applying an initial disturbance to the aircraft, and performing a motion-deformation coupling analysis on the aircraft after the initial disturbance is applied under the initial steady-state condition, and determining the displacement of the surface nodes of the aircraft relative to the equilibrium position based on the corresponding motion-deformation coupling analysis results.

[0062] In this embodiment, after the aforementioned iterative adjustment, the trim parameters of the aircraft under the set of flight conditions are obtained, which correspond to the equilibrium state of aerodynamics, motion and deformation during flight. Body flutter analysis can be regarded as a dynamic stability analysis of the equilibrium state; and based on the definition of dynamic stability, an initial disturbance can be applied to the equilibrium state to observe whether it can eventually return to the equilibrium state. For symmetrical large aspect ratio flying wings, the body flutter phenomenon is usually induced by the coupling of the pitch of the whole machine and the first-order bending. Therefore, for such aircraft, considering that the application of the first-order bending disturbance is more troublesome, the initial disturbance should be applied to the pitch degree of freedom first; and the specific application method is recommended to apply it in the form of initial velocity, that is, in the solution of the 6-degree-of-freedom rigid body motion in the rigid-flexible coupling analysis, the initial velocity of the pitch axis is set.

[0063] Then, the motion deformation coupling analysis is performed on the aircraft after the initial disturbance is applied under the initial steady-state conditions. Based on the motion state of the aircraft in the initial or previous time step, and the resultant force and torque of the aircraft in the current time step, the 6-degree-of-freedom motion displacement, velocity and acceleration of the aircraft in the current time step are solved. The solution process is similar to the analysis of the 6-degree-of-freedom motion of a traditional single rigid body space. Then, based on the 6-degree-of-freedom motion state of the aircraft obtained by the motion solution, the inertial acceleration field in the internal body coordinate system of the current time step is calculated. , and based on the inertial acceleration field, the inertial load of each particle in the body coordinate system is obtained: ,in is the mass of the particle element. Then, based on the deformation state of the aircraft in the previous time step and the load in the current time step, the inertial load obtained by the above-mentioned inertial acceleration field is superimposed, and the finite element time marching solution is performed to obtain the grid node displacement, velocity, and acceleration corresponding to the deformation of the aircraft in the current time step. The current position of each node is obtained by using the displacement of the aircraft grid node in the current time step, and the inertia tensor of the aircraft in the deformation state in the current time step is calculated accordingly. Finally, based on the corresponding accuracy requirements, after the above-mentioned steps based on the motion state of the aircraft at the initial or previous time step to calculate the inertia tensor of the aircraft in the deformation state in the current time step are completed, the result is returned to the step based on the motion state of the aircraft at the initial or previous time step, and the steps based on the motion state of the aircraft at the initial or previous time step to calculate the inertia tensor of the aircraft in the deformation state in the current time step are repeated, and the result is returned to the step based on the motion state of the aircraft at the initial or previous time step, that is, the coupled iteration of motion and deformation is performed. When the residual, that is, the difference between two adjacent iterations, is less than a predetermined value, it is determined that the coupled iteration is completed, the obtained result can be passed to the subsequent aerodynamic solution, and preparation is made for the motion deformation coupled time advancement of the next time step.

[0064] After completing the motion-deformation coupling analysis, the displacement of the aircraft surface nodes relative to the equilibrium position can be determined using the preset displacement calculation formula based on the corresponding motion-deformation coupling analysis results; the preset displacement calculation formula is:

[0065] ;

[0066] in, is the displacement of the node on the aircraft surface relative to the equilibrium position; is the displacement vector of the aircraft relative to the equilibrium position; is the transformation matrix from the body coordinate system to the inertial system; is the vector corresponding to the node in the body coordinate system when the aircraft is not deformed; is the deformation displacement vector of the node relative to the body coordinate system.

[0067] Step S13, updating the aerodynamic grid corresponding to the aircraft through the displacement, and obtaining the aerodynamic force on the surface of the aircraft at the current time step based on the corresponding updated aerodynamic grid, and mapping the aerodynamic force to the structural grid of the aircraft using the aerodynamic structure coupling interface and the coupling data transfer matrix to obtain the aerodynamic load of the aircraft.

[0068] In this embodiment, a mesh deformation method is used to update the aerodynamic mesh, that is, after obtaining the aforementioned surface mesh node displacement, the topological connection relationship between the mesh nodes is retained, and the displacement of the object surface nodes is orderly transferred to the spatial mesh, and the updated mesh is obtained using the displacement of each node. Based on the updated aerodynamic mesh, a time-marching solution of the CFD flow field is performed to obtain the spatial flow field and the aerodynamic force distribution on the aircraft surface at the current time step. Then, the aerodynamic force of each interpolation point of the surface mesh in the aerodynamic mesh corresponding to the aircraft is extracted, and the aerodynamic force of each interpolation point of the structural mesh corresponding to the aircraft is determined based on the dot product result of the aerodynamic force of each interpolation point of the surface mesh and the coupling data transfer matrix; the aerodynamic force of each interpolation point of the structural mesh is interpolated to obtain the aerodynamic load of the structural mesh corresponding to the aircraft. Specifically, based on the obtained aerodynamic force and the obtained coupling interface, the aerodynamic force is mapped to the structural mesh. First, the aerodynamic force of each interpolation point of the surface mesh in the aerodynamic mesh is extracted. , the aerodynamic force value of each interpolation point of the structure grid is obtained by dot multiplication of the coupling transfer matrix: , then interpolate the point aerodynamics based on the structural grid , the complete aerodynamic loads acting on the surface grid of the structure are interpolated for subsequent structural time-marching solution.

[0069] Step S14, performing motion deformation coupling iteration and aerodynamic structure coupling iteration based on the aerodynamic load, and performing coupling time advancement when the coupling residual meets the preset conditions, judging the dynamic stability of the aircraft according to the displacement of the aircraft surface nodes relative to the equilibrium position during the time advancement process, and completing the time domain analysis of the flutter of the aircraft body degrees of freedom.

[0070] In this embodiment, the body tremor coupling iteration is a three-field double coupling iteration, including coupling at the aerodynamic structure level and coupling at the motion deformation level. Therefore, the motion deformation coupling iteration is performed based on the aerodynamic load to obtain the updated displacement of the aircraft surface node relative to the equilibrium position; the updated displacement of the aircraft surface node relative to the equilibrium position is used to jump back to the step of updating the aerodynamic grid corresponding to the aircraft through the displacement, to obtain the aerodynamic load of the aircraft, and then jump again to the step of performing motion deformation coupling iteration based on the aerodynamic load. Specifically, firstly, the coupling iteration at the motion deformation level is performed, that is, the step based on the initial or previous time step of the aircraft motion state is repeatedly executed to calculate the aircraft inertia tensor under the deformation state of the current time step based on this step. After completion, the result is returned to the step based on the initial or previous time step of the aircraft motion state until the coupling residual meets the accuracy requirements. At this time, the obtained structural surface mesh displacement can be used for aerodynamic structure coupling iteration; then the aerodynamic structure coupling iteration is performed, that is, after the motion deformation coupling analysis is repeatedly executed, the preset displacement calculation formula can be used to determine the displacement of the aircraft surface node relative to the equilibrium position based on the corresponding motion deformation coupling analysis results to the aerodynamic force based on the structural mesh interpolation point. , interpolate to obtain the complete aerodynamic load on the surface grid of the structure for the subsequent structural time advancement solution step. In each repeated process, that is, in the aerodynamic structure coupling iteration step, after obtaining the aerodynamic load of the current iteration step, it is necessary to return to the motion deformation coupling solution to perform the motion deformation coupling iterative solution of the corresponding aerodynamic load to obtain the updated object surface grid displacement for the next aerodynamic structure coupling iteration step, until the coupling residual meets the accuracy requirements. At this time, the coupling time advancement can be performed to start the time solution of the next time step.

[0071] Finally, the amplitude of the reciprocating motion of the aircraft is determined based on the displacement of the aircraft surface nodes relative to the equilibrium position during the advancement of time; if the reciprocating motion amplitude meets the preset reduction rule, it is determined that the aircraft has dynamic stability; if the reciprocating motion amplitude does not meet the preset reduction rule, it is determined that the aircraft does not have the said dynamic stability. Body tremor analysis belongs to dynamic stability analysis, and its main purpose is to determine whether the aircraft can eventually come to rest in the equilibrium position, that is, return to the equilibrium state, after the corresponding rigid body motion or flexible deformation degree of freedom is deviated from the equilibrium position by a small disturbance. In general, the aircraft will reciprocate near the equilibrium position after the degree of freedom to be analyzed deviates from the equilibrium position. For the displacement of the aircraft's motion and deformation degrees of freedom relative to the equilibrium position under body tremor conditions, the following should be done. Figure 5 As shown in the figure, it is a vibration with a certain amplitude and frequency. The lowest frequency is taken as the fundamental frequency. To ensure the reliability of the judgment, the total time of the aerodynamic structure coupling solution should not be less than 5 times of the period. If the amplitude of the reciprocating motion gradually decreases, it is called convergence. At this time, the aircraft will return to the equilibrium state within a finite time, indicating that the aircraft has dynamic stability in this degree of freedom; if the amplitude of the reciprocating motion gradually increases, it is called divergence. At this time, it will not be able to return to the equilibrium state, indicating that the aircraft does not have dynamic stability in this degree of freedom; in addition, there is a critical state where the amplitude remains constant. Finally, the three-field coupling problem is realized through the dual coupling of aerodynamic structure and motion deformation, which improves the solution efficiency while ensuring the solution accuracy.

[0072] In summary, when performing time domain analysis on the body degree of freedom flutter of an aircraft, the present application first constructs an aircraft aerodynamic mesh model and an aircraft structural mesh model according to the geometric shape, structure, size and material of the aircraft, extracts the aerodynamic-structural coupling interface of the aircraft, and uses the aircraft aerodynamic mesh model and the aircraft structural mesh model to determine the initial steady state of the aircraft through static balancing; then, an initial disturbance is applied to the aircraft, and a motion-deformation coupling analysis is performed on the aircraft after the initial disturbance is applied under the initial steady state condition, and the displacement of the surface nodes of the aircraft relative to the equilibrium position is determined based on the corresponding motion-deformation coupling analysis results; and then The aerodynamic grid corresponding to the aircraft is updated through the displacement, and the aerodynamic force on the surface of the aircraft at the current time step is obtained based on the corresponding updated aerodynamic grid, and the aerodynamic force is mapped to the structural grid of the aircraft using the aerodynamic structure coupling interface and the coupling data transfer matrix to obtain the aerodynamic load of the aircraft; finally, motion deformation coupling iteration and aerodynamic structure coupling iteration are performed based on the aerodynamic load, and when the coupling residual meets the preset conditions, coupling time advancement is performed, and the dynamic stability of the aircraft is judged according to the displacement of the aircraft surface node relative to the equilibrium position during the time advancement process, and the time domain analysis of the aircraft body degree of freedom flutter is completed. It can be seen that the present application ensures the solution accuracy of complex configuration flight through motion deformation coupling solution, and can handle body flutter problems under large deformation conditions; at the same time, combined with the aerodynamic structure coupling solution method, through the dual coupling of aerodynamic structure and motion deformation, high-precision analysis of aircraft body flutter under large deformation conditions is achieved, and the solution efficiency is improved.

[0073] Based on the above embodiment, it can be known that the present application can determine the initial steady state of the aircraft by static trimming using the aircraft aerodynamic grid model and the aircraft structural grid model. Next, the static aerodynamic trimming process for the initial state will be described in detail. Figure 6 As shown, the present application discloses a specific method for time-domain analysis of aircraft body freedom flutter.

[0074] When solving the initial steady state, the present application performs aerostatic trim analysis of the aircraft to obtain a stable flight state under given flight conditions such as airspeed and altitude, which serves as the initial steady state for subsequent dynamic stability analysis, including initial trim parameter prediction, whole-machine aerostatic trim analysis, lift-weight balance iteration, and moment trim iteration.

[0075] When making the initial trim parameter prediction, we first make an initial estimate of the trim parameters based on the angle of attack and rudder angle of the aircraft under given flight parameters such as airspeed and altitude based on the lift coefficient curve and pitch moment coefficient curve of the aircraft obtained from previous work such as flight / wind tunnel tests, static aeroelastic analysis, and empirical estimation. ,in is the air density near the aircraft, is the speed of the aircraft relative to the atmosphere, referred to as airspeed. is the dynamic pressure, is the reference area of ​​the aircraft, usually the wing area, is the lift coefficient. In addition to the properties of the wing airfoil, it is mainly affected by the aircraft elevation angle, that is, the angle between the aircraft axis and the airspeed direction. The specific relationship with the elevation angle is given by the lift coefficient curve. In the commonly used elevation angle range, the lift coefficient is usually linearly related to the elevation angle, that is, the relationship between the lift coefficient and the elevation angle can be written as: ,in is the slope of the lift line, is the current elevation angle, is zero lift angle; during the flight of an aircraft, lift is mainly used to overcome gravity, so lift With the whole machine gravity At the same time, during normal flight, the lift center and the center of gravity often do not coincide at a given elevation angle, resulting in a pitch moment: ,in is the average aerodynamic chord length, is the pitch moment coefficient, which is mainly affected by the aircraft's pitch angle. The specific relationship is given by the pitch moment coefficient curve. The pitch moment is determined by the elevator control moment. Balance, the control torque is determined by the elevator deflection angle Determined, so the moment balance: This gives an initial estimate of the elevator angle.

[0076] Then, when performing the whole-machine static aeroelastic analysis, based on the given flight parameters (speed, altitude, etc.) and trim parameters (pitch angle, rudder deflection angle), the CFD static aeroelastic analysis is performed on the aircraft model to obtain the actual deformation and aerodynamic force of the aircraft under this set of parameter conditions. The specific method is the same as that of conventional static aeroelastics: first, perform the whole-machine CFD calculation under the given flight and trim parameters to obtain the aerodynamic force distribution on the aircraft surface in the form of pressure, and then use the coupling interface to map the aerodynamic force to the structural finite element mesh to solve the flexible deformation of the aircraft under the action of the aerodynamic force. The flexible deformation should use inertial release and other methods to eliminate rigid body motion. Then, based on the displacement of the object surface mesh nodes corresponding to the flexible deformation, the aerodynamic mesh is updated, and CFD calculation is performed again in the updated aerodynamic mesh. Repeat this iterative process until the difference in the deformation displacement of the aircraft in two adjacent iterations is small and meets the accuracy requirements, and the deformation and aerodynamic results of the aircraft under the current flight and trim parameters can be obtained.

[0077] Lift-weight balance iteration means that for the elevation angle given by the initial prediction, the actual lift corresponding to it is not necessarily balanced with the gravity, and the elevation angle needs to be iteratively adjusted until the deviation between gravity and lift is small and meets the accuracy requirements. At the same time, the moment balance iteration is for the rudder angle given by the initial prediction, which is also not guaranteed to be balanced with the pitch moment. At the same time, the corresponding pitch moment will also change after the elevation angle is updated, so the rudder angle also needs to be iteratively adjusted. The specific iteration process is similar to the lift-weight balance iteration, and will not be repeated here.

[0078] In summary, the present application performs aeroelastic analysis based on pre-predicted trim parameters and flight parameters corresponding to the aircraft to obtain lift and pitch moment corresponding to the aircraft; the trim parameters include the pitch angle and elevator deflection angle of the aircraft; the lift is used to iteratively adjust the pitch angle based on lift-to-weight balance to obtain the adjusted pitch angle; the pitch moment is used to iteratively adjust the elevator deflection angle based on pitch moment balance to obtain the adjusted elevator deflection angle; finally, the initial steady state is determined based on the adjusted pitch angle and the adjusted elevator deflection angle.

[0079] In this way, the motion-deformation coupled dynamic response of the structural model after the initial disturbance under the initial steady-state conditions can be further solved, thereby ensuring the solution accuracy of complex configuration flight and dealing with body tremor problems under large deformation conditions.

[0080] See also Figure 7 As shown, an embodiment of the present invention discloses a device for analyzing the flutter time domain of an aircraft body degree of freedom, comprising:

[0081] The initial steady state determination module 11 is used to construct an aircraft aerodynamic grid model and an aircraft structural grid model according to the geometric shape, structure, size and material of the aircraft, extract the aerodynamic-structural coupling interface of the aircraft, and determine the initial steady state of the aircraft by static trim using the aircraft aerodynamic grid model and the aircraft structural grid model;

[0082] a displacement determination module 12, configured to apply an initial disturbance to the aircraft, and perform a motion-deformation coupling analysis on the aircraft after the initial disturbance is applied under the initial steady-state condition, and determine the displacement of the surface nodes of the aircraft relative to the equilibrium position based on the corresponding motion-deformation coupling analysis results;

[0083] an aerodynamic force and load determination module 13, configured to update the aerodynamic grid corresponding to the aircraft through the displacement, obtain the aerodynamic force on the surface of the aircraft at the current time step based on the corresponding updated aerodynamic grid, and map the aerodynamic force to the structural grid of the aircraft using the aerodynamic structure coupling interface and the coupling data transfer matrix to obtain the aerodynamic load of the aircraft;

[0084] The dynamic stability judgment module 14 is used to perform motion deformation coupling iteration and aerodynamic structure coupling iteration based on the aerodynamic load, and when the coupling residual meets the preset conditions, perform coupling time advancement, judge the dynamic stability of the aircraft according to the displacement of the aircraft surface node relative to the equilibrium position during the time advancement process, and complete the time domain analysis of the flutter of the aircraft body degree of freedom.

[0085] It can be seen that the present application ensures the solution accuracy of complex configuration flight through motion deformation coupling solution, and can handle the body flutter problem under large deformation conditions; at the same time, combined with the aerodynamic structure coupling solution method, through the dual coupling of aerodynamic structure and motion deformation, high-precision analysis of aircraft body flutter under large deformation conditions is achieved, and the solution efficiency is improved.

[0086] In some specific embodiments, the initial steady state determination module 11 may specifically include:

[0087] A lift and pitch moment acquisition unit, configured to perform aeroelastic analysis based on pre-predicted trim parameters and flight parameters corresponding to the aircraft, so as to obtain lift and pitch moment corresponding to the aircraft; the trim parameters include an elevation angle and an elevator deflection angle of the aircraft;

[0088] A first adjustment unit, configured to iteratively adjust the elevation angle based on lift-weight balance using the lift force to obtain an adjusted elevation angle;

[0089] A second adjustment unit is used to iteratively adjust the elevator deflection angle based on the pitching moment balance by using the pitching moment to obtain an adjusted elevator deflection angle;

[0090] An initial steady state determining unit is used to determine the initial steady state according to the adjusted pitch angle and the adjusted elevator deflection angle.

[0091] In some specific embodiments, the displacement determination module 12 may specifically include:

[0092] The displacement determination unit is used to determine the displacement of the aircraft surface node relative to the equilibrium position based on the corresponding motion deformation coupling analysis results using a preset displacement calculation formula; the preset displacement calculation formula is:

[0093] ;

[0094] in, is the displacement of the node on the surface of the aircraft relative to the equilibrium position; is the displacement vector of the aircraft relative to the equilibrium position; is the transformation matrix from the body coordinate system to the inertial system; is the vector corresponding to the node in the body coordinate system when the aircraft is in an undeformed state; is the deformation displacement vector of the node relative to the body coordinate system.

[0095] In some specific embodiments, the aerodynamic force and load determination module 13 may specifically include:

[0096] The aerodynamic grid updating unit is used to transfer the displacement of the surface nodes of the aircraft relative to the equilibrium position to the spatial grid based on the topological connection relationship between the grid nodes, so as to update the aerodynamic grid corresponding to the aircraft.

[0097] In some specific embodiments, the initial steady state determination module 11 may specifically include:

[0098] A coupling interface determination unit is used to determine the outer surface of the aircraft as the aerodynamic structure coupling interface, so as to extract the surface nodes of the aircraft, select representative nodes from the surface nodes, and construct the coupling data transfer matrix based on the representative nodes; the representative nodes include the object surface grid nodes in the aerodynamic grid and the surface grid nodes in the structural grid.

[0099] In some specific embodiments, the aerodynamic force and load determination module 13 may specifically include:

[0100] an aerodynamic force determination unit, configured to extract the aerodynamic force of each interpolation point of the surface grid in the aerodynamic grid corresponding to the aircraft, and determine the aerodynamic force of each interpolation point of the structural grid corresponding to the aircraft based on the dot product result of the aerodynamic force of each interpolation point of the surface grid and the coupling data transfer matrix;

[0101] The aerodynamic load acquisition unit is used to interpolate the aerodynamic force of each interpolation point of the structural grid to obtain the aerodynamic load of the structural grid corresponding to the aircraft.

[0102] In some specific embodiments, the dynamic stability judgment module 14 may specifically include:

[0103] A displacement acquisition unit, used for performing motion-deformation coupling iteration based on the aerodynamic load to acquire updated displacements of the nodes on the surface of the aircraft relative to the equilibrium position;

[0104] A jump unit is used to jump again to the step of updating the aerodynamic mesh corresponding to the aircraft by using the updated displacement of the aircraft surface node relative to the equilibrium position, obtain the aerodynamic load of the aircraft, and jump again to the step of performing motion deformation coupling iteration based on the aerodynamic load.

[0105] In some specific embodiments, the dynamic stability judgment module 14 may specifically include:

[0106] a reciprocating motion amplitude determination unit, configured to determine the reciprocating motion amplitude of the aircraft according to the displacement of the surface nodes of the aircraft relative to the equilibrium position during the time advancement process;

[0107] A first determination unit, configured to determine that the aircraft has the dynamic stability if the reciprocating motion amplitude meets a preset reduction rule;

[0108] The second determination unit is configured to determine that the aircraft does not have the dynamic stability if the reciprocating motion amplitude does not meet a preset reduction rule.

[0109] Furthermore, the present application also discloses an electronic device. Figure 8 This is a structural diagram of an electronic device 20 according to an exemplary embodiment. The content in the diagram cannot be regarded as any limitation on the scope of use of the present application.

[0110] Figure 8 A schematic diagram of the structure of an electronic device 20 provided in an embodiment of the present application. The electronic device 20 may specifically include: at least one processor 21, at least one memory 22, a power supply 23, a communication interface 24, an input / output interface 25, and a communication bus 26. The memory 22 is used to store a computer program, which is loaded and executed by the processor 21 to implement the relevant steps in the time domain analysis method for the degree of freedom flutter of an aircraft body disclosed in any of the aforementioned embodiments. In addition, the electronic device 20 in this embodiment may specifically be an electronic computer.

[0111] In this embodiment, the power supply 23 is used to provide working voltage for each hardware device on the electronic device 20; the communication interface 24 can create a data transmission channel between the electronic device 20 and the external device, and the communication protocol it follows is any communication protocol that can be applied to the technical solution of the present application, and is not specifically limited here; the input and output interface 25 is used to obtain external input data or output data to the outside world, and its specific interface type can be selected according to specific application needs and is not specifically limited here.

[0112] In addition, the memory 22 as a carrier for resource storage may be a read-only memory, a random access memory, a disk or an optical disk, etc. The resources stored thereon may include an operating system 221, a computer program 222, etc., and the storage method may be temporary storage or permanent storage.

[0113] The operating system 221 is used to manage and control the hardware devices and computer program 222 on the electronic device 20, and may be Windows Server, Netware, Unix, Linux, etc. In addition to the computer program that can be used to complete the time-domain analysis method for the degree of freedom flutter of the aircraft body executed by the electronic device 20 disclosed in any of the aforementioned embodiments, the computer program 222 may further include computer programs that can be used to complete other specific tasks.

[0114] Furthermore, the present application also discloses a computer-readable storage medium for storing a computer program; wherein, when the computer program is executed by a processor, the aforementioned disclosed method for analyzing the flutter time domain of the aircraft body freedom is implemented. The specific steps of the method can refer to the corresponding contents disclosed in the aforementioned embodiments, and will not be repeated here.

[0115] In this specification, each embodiment is described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the embodiments can be referred to each other. For the device disclosed in the embodiment, since it corresponds to the method disclosed in the embodiment, the description is relatively simple, and the relevant parts can be referred to the method part.

[0116] Professionals may further appreciate that the units and algorithm steps of each example described in conjunction with the embodiments disclosed herein can be implemented in electronic hardware, computer software, or a combination of the two. In order to clearly illustrate the interchangeability of hardware and software, the composition and steps of each example have been generally described in the above description according to function. Whether these functions are performed in hardware or software depends on the specific application and design constraints of the technical solution. Professionals and technicians may use different methods to implement the described functions for each specific application, but such implementation should not be considered to be beyond the scope of this application.

[0117] The steps of the method or algorithm described in conjunction with the embodiments disclosed herein may be implemented directly using hardware, a software module executed by a processor, or a combination of the two. The software module may be placed in a random access memory (RAM), a memory, a read-only memory (ROM), an electrically programmable ROM, an electrically erasable programmable ROM, a register, a hard disk, a removable disk, a CD-ROM, or any other form of storage medium known in the art.

[0118] Finally, it should be noted that, in this article, relational terms such as first and second, etc. are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, the elements defined by the sentence "comprise a ..." do not exclude the presence of other identical elements in the process, method, article or device including the elements.

[0119] The technical solution provided by the present application is introduced in detail above. Specific examples are used in this article to illustrate the principles and implementation methods of the present application. The description of the above embodiments is only used to help understand the method of the present application and its core idea. At the same time, for general technicians in this field, according to the idea of ​​the present application, there will be changes in the specific implementation methods and application scope. In summary, the content of this specification should not be understood as a limitation on the present application.

Claims

1. A method for analyzing the flutter of an aircraft body in the time domain, characterized in that: include: Constructing an aircraft aerodynamic mesh model and an aircraft structural mesh model according to the geometric shape, structure, size and material of the aircraft, extracting the aerodynamic-structural coupling interface of the aircraft, and determining the initial steady state of the aircraft by static trim using the aircraft aerodynamic mesh model and the aircraft structural mesh model; Applying an initial disturbance to the aircraft, and performing a motion-deformation coupling analysis on the aircraft after the initial disturbance is applied under the initial steady-state condition, and determining the displacement of the surface nodes of the aircraft relative to the equilibrium position based on the corresponding motion-deformation coupling analysis results; The aerodynamic grid corresponding to the aircraft is updated by the displacement, and the aerodynamic force on the surface of the aircraft at the current time step is obtained based on the corresponding updated aerodynamic grid, and the aerodynamic force is mapped to the structural grid of the aircraft by using the aerodynamic-structure coupling interface and the coupling data transfer matrix to obtain the aerodynamic load of the aircraft; Based on the aerodynamic load, motion deformation coupling iteration and aerodynamic structure coupling iteration are performed, and when the coupling residual meets the preset conditions, coupling time advancement is performed. The dynamic stability of the aircraft is judged according to the displacement of the aircraft surface nodes relative to the equilibrium position during the time advancement process, and the time domain analysis of the flutter of the aircraft body degrees of freedom is completed.

2. The method for analyzing the flutter of an aircraft body in time domain according to claim 1, characterized in that: The method of determining the initial steady state of the aircraft by static trimming using the aircraft aerodynamic grid model and the aircraft structural grid model includes: Performing aeroelastic analysis based on pre-predicted trim parameters and flight parameters corresponding to the aircraft to obtain lift and pitch moment corresponding to the aircraft; the trim parameters include an elevation angle and an elevator deflection angle of the aircraft; Iteratively adjusting the elevation angle based on lift-weight balance using the lift force to obtain an adjusted elevation angle; Iteratively adjusting the elevator deflection angle based on the pitching moment balance using the pitching moment to obtain an adjusted elevator deflection angle; The initial steady state is determined according to the adjusted pitch angle and the adjusted elevator deflection angle.

3. The method for analyzing the flutter of an aircraft body in time domain according to claim 1, characterized in that: Determining the displacement of the aircraft surface node relative to the equilibrium position based on the corresponding motion deformation coupling analysis results includes: The displacement of the aircraft surface node relative to the equilibrium position is determined based on the corresponding motion deformation coupling analysis results using a preset displacement calculation formula; the preset displacement calculation formula is: ； in, is the displacement of the node on the surface of the aircraft relative to the equilibrium position; is the displacement vector of the aircraft relative to the equilibrium position; is the transformation matrix from the body coordinate system to the inertial system; is the vector corresponding to the node in the body coordinate system when the aircraft is in an undeformed state; is the deformation displacement vector of the node relative to the body coordinate system.

4. The method for analyzing the flutter of an aircraft body in time domain according to claim 1, characterized in that: The updating of the aerodynamic grid corresponding to the aircraft by using the displacement includes: Based on the topological connection relationship between the grid nodes, the displacement of the aircraft surface nodes relative to the equilibrium position is transmitted to the space grid to update the aerodynamic grid corresponding to the aircraft.

5. The method for analyzing the flutter of an aircraft body in time domain according to claim 1, characterized in that: The extracting of the aerodynamic structure coupling interface of the aircraft comprises: The outer surface of the aircraft is determined as the aerodynamic-structural coupling interface so as to extract the surface nodes of the aircraft, select representative nodes from the surface nodes, and construct the coupling data transfer matrix based on the representative nodes; the representative nodes include the object surface mesh nodes in the aerodynamic mesh and the surface mesh nodes in the structural mesh.

6. The method for analyzing the flutter of an aircraft body in time domain according to claim 5, characterized in that: The step of mapping the aerodynamic force to the structural grid of the aircraft by using the aerodynamic-structural coupling interface and the coupling data transfer matrix to obtain the aerodynamic load of the aircraft includes: Extracting the aerodynamic force of each interpolation point of the surface grid in the aerodynamic grid corresponding to the aircraft, and determining the aerodynamic force of each interpolation point of the structural grid corresponding to the aircraft based on the dot product result of the aerodynamic force of each interpolation point of the surface grid and the coupling data transfer matrix; The aerodynamic forces of each interpolation point of the structural grid are interpolated to obtain the aerodynamic load of the structural grid corresponding to the aircraft.

7. The method for analyzing the flutter of an aircraft body in time domain according to claim 1, characterized in that: The performing motion deformation coupling iteration and aerodynamic structure coupling iteration based on the aerodynamic load includes: Performing motion-deformation coupling iteration based on the aerodynamic load to obtain updated displacements of the aircraft surface nodes relative to equilibrium positions; The updated displacement of the aircraft surface nodes relative to the equilibrium position is used to jump back to the step of updating the aerodynamic mesh corresponding to the aircraft through the displacement, the aerodynamic load of the aircraft is obtained, and the step of performing motion deformation coupling iteration based on the aerodynamic load is jumped again.

8. The method for analyzing the flutter of an aircraft body in time domain according to any one of claims 1 to 7, characterized in that: The determining the dynamic stability of the aircraft according to the displacement of the aircraft surface node relative to the equilibrium position during the time advancement process includes: Determining the reciprocating motion amplitude of the aircraft according to the displacement of the surface nodes of the aircraft relative to the equilibrium position during the time advancement process; If the reciprocating motion amplitude meets the preset reduction rule, it is determined that the aircraft has the dynamic stability; If the reciprocating motion amplitude does not meet the preset reduction rule, it is determined that the aircraft does not have the dynamic stability.

9. An aircraft body freedom flutter time domain analysis device, characterized in that: include: An initial steady state determination module is used to construct an aircraft aerodynamic grid model and an aircraft structural grid model according to the geometric shape, structure, size and material of the aircraft, extract the aerodynamic-structural coupling interface of the aircraft, and determine the initial steady state of the aircraft by static trim using the aircraft aerodynamic grid model and the aircraft structural grid model; a displacement determination module, used for applying an initial disturbance to the aircraft, performing a motion-deformation coupling analysis on the aircraft after the initial disturbance is applied under the initial steady-state condition, and determining the displacement of the surface nodes of the aircraft relative to the equilibrium position based on the corresponding motion-deformation coupling analysis results; an aerodynamic force and load determination module, configured to update the aerodynamic mesh corresponding to the aircraft through the displacement, obtain the aerodynamic force on the surface of the aircraft at the current time step based on the corresponding updated aerodynamic mesh, and map the aerodynamic force to the structural mesh of the aircraft using the aerodynamic structure coupling interface and the coupling data transfer matrix to obtain the aerodynamic load of the aircraft; The dynamic stability judgment module is used to perform motion deformation coupling iteration and aerodynamic structure coupling iteration based on the aerodynamic load, and when the coupling residual meets the preset conditions, perform coupling time advancement, judge the dynamic stability of the aircraft according to the displacement of the aircraft surface node relative to the equilibrium position during the time advancement process, and complete the time domain analysis of the flutter of the aircraft body degree of freedom.

10. An electronic device, characterized in that: include: Memory, used to store computer programs; A processor is used to execute the computer program to implement the time domain analysis method for the degree of freedom flutter of an aircraft body as claimed in any one of claims 1 to 8.

11. A computer-readable storage medium, characterized in that: Used to store a computer program, wherein when the computer program is executed by a processor, the method for time-domain analysis of aircraft body degree of freedom flutter according to any one of claims 1 to 8 is implemented.

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