Chatter analysis modeling method

Abstract

This invention provides a flutter analysis modeling method, comprising: establishing a structural model of the entire aircraft; performing finite element discretization on the structural model of the entire aircraft to obtain a modal analysis model of the entire aircraft; selecting fuselage interpolation nodes on the fuselage and wing interpolation nodes on the wings of the entire aircraft; determining the analysis conditions based on a set Mach number, determining control parameters, aerodynamic parameters, and coordinate system, and generating post-processing keywords; sequentially selecting key nodes of the wing-type finite element model in the modal analysis model of the entire aircraft, establishing wing-type aerodynamic surfaces, and using wing interpolation nodes to interpolate the wing-type aerodynamic surfaces to complete the modeling of the wing-type aerodynamic surfaces; establishing body-type aerodynamic surfaces, and using fuselage interpolation nodes to interpolate the body-type aerodynamic surfaces to complete the modeling of the body-type aerodynamic surfaces. The technical solution of this invention solves the technical problem in the prior art where modeling in aircraft modal analysis and flutter analysis is independent.

Description

Technical Field

[0001] This invention relates to the field of aeroelastic analysis technology for aircraft, and in particular to a flutter analysis modeling method. Background Technology

[0002] Flutter is a self-excited vibration phenomenon in aircraft caused by disturbances, which in turn induces aerodynamic excitation. Flutter often leads to rapid structural failure within seconds, seriously affecting aircraft safety. Therefore, flutter has always been a key aeroelastic issue in aircraft design.

[0003] The fundamental data for aircraft flutter analysis includes modal data such as aircraft frequencies / mode shapes, unsteady aerodynamic forces, and flight envelope. The modal data, such as frequencies / mode shapes, are obtained through modal analysis. Modal analysis results are the foundational data for flutter analysis; to obtain these results, a modal analysis model must first be established. Modal analysis is based on the aircraft's detailed structure, center of mass, and moment of inertia data; the modal analysis model represents the aircraft's inherent properties. The unsteady aerodynamic forces are modeled based on the aircraft's external shape. Simultaneously, in flutter analysis, interpolation is performed between the structural mesh nodes and the unsteady aerodynamic force model mesh nodes to transfer loads.

[0004] Aircraft modal analysis typically involves modeling in the pre-processing program Hypermesh. Flutter analysis, on the other hand, is modeled by inputting flutter analysis keyword cards (including control parameters, aerodynamic parameters, fuselage / wing aerodynamic surfaces, flutter analysis settings, etc.). In particular, body aerodynamic surface modeling involves a significant amount of manual sorting, and interpolation node data needs to be manually extracted from the modal analysis model. Currently, flutter and modal analysis modeling are performed in different software, requiring manual interaction for interpolation node data, making it impossible to unify modal and flutter analysis modeling. Therefore, a more convenient flutter analysis modeling method and system is urgently needed. Summary of the Invention

[0005] This invention provides a flutter analysis modeling method that can solve the technical problem that modeling in aircraft modal analysis and flutter analysis is independent in the prior art.

[0006] This invention provides a flutter analysis modeling method, which includes: establishing a structural model of the entire aircraft; performing finite element discretization on the structural model of the entire aircraft to obtain a modal analysis model of the entire aircraft; based on the modal analysis model of the entire aircraft, selecting fuselage interpolation nodes on the fuselage and wing interpolation nodes on the wing surfaces of the entire aircraft; determining the analysis conditions according to the set Mach number, determining the control parameters, aerodynamic parameters, and coordinate system, and generating post-processing keywords; sequentially selecting key nodes of the wing-type finite element model in the modal analysis model of the entire aircraft, establishing wing-type aerodynamic surfaces according to the selected key nodes and the set spanwise and chordwise node numbers, and using wing interpolation nodes to interpolate the wing-type aerodynamic surfaces to complete the modeling of the wing-type aerodynamic surfaces; establishing body-type aerodynamic surfaces, and using fuselage interpolation nodes to interpolate the body-type aerodynamic surfaces to complete the modeling of the body-type aerodynamic surfaces; and constructing a flutter analysis model of the entire aircraft based on the wing-type aerodynamic surface model and the body-type aerodynamic surface model.

[0007] Furthermore, the selection of interpolation nodes on the entire aircraft fuselage specifically includes: selecting the intersection generatrix of the first plane of the entire aircraft and the uppermost side of the fuselage as the first generatrix node; selecting the intersection generatrix of the second plane of the entire aircraft and the leftmost side of the fuselage as the second generatrix node; and selecting the intersection generatrix of the second plane of the entire aircraft and the rightmost side of the fuselage as the third generatrix node, wherein the first plane and the second plane are perpendicular.

[0008] Furthermore, selecting wing interpolation nodes on the entire wing surface of the aircraft specifically includes: selecting nodes on the entire wing surface whose stiffness exceeds a set stiffness threshold as wing interpolation nodes.

[0009] Furthermore, the control parameters include modal input format and mode output format, the aerodynamic parameters include reference span, reference chord length and reference area, and the coordinate system includes local coordinate system and Cartesian coordinate system; the nodes on the entire aircraft's wing surface whose stiffness exceeds the set stiffness threshold include wing frame nodes and stiffener nodes.

[0010] Furthermore, post-processing keywords include Vg plot, Vf plot, flutter mode, and interpolation mode.

[0011] Furthermore, the key nodes of the finite element model of the wing include the root chord leading edge point, the root chord trailing edge point, the pointed chord leading edge point, and the pointed chord trailing edge point.

[0012] Furthermore, the creation of volumetric aerodynamic surfaces specifically includes: creating volumetric aerodynamic surfaces using arbitrary cross-sections, or in circular or elliptical shapes.

[0013] Furthermore, after completing the flutter analysis model, the flutter analysis modeling method also includes: completing the unsteady aerodynamic forces and flutter solution of the entire aircraft based on the flutter analysis model.

[0014] Furthermore, when the Mach number Ma is less than 1, the ZONA6 method is used to solve the unsteady aerodynamic forces; when the Mach number Ma is greater than 1, the ZONA7U method is used to solve the unsteady aerodynamic forces; the flutter analysis modeling method uses the G method to solve the flutter equation.

[0015] Furthermore, the finite element discretization of the structural model of the entire aircraft to obtain the modal analysis model of the entire aircraft specifically includes: based on the structural model, mass center of mass and moment of inertia data of the entire aircraft, the finite element discretization of the structural model of the entire aircraft to obtain the modal analysis model of the entire aircraft.

[0016] This invention provides a flutter analysis modeling method. This method integrates flutter analysis and modal analysis in the same software environment, and incorporates interpolation nodes and other information from modal analysis. This enables seamless integration of modal and flutter analysis, facilitating data interaction between the two models, improving data transmission and modeling efficiency, and avoiding manual modification and modeling of flutter analysis cards. Therefore, the flutter analysis modeling method provided by this invention solves the problems of manual modification of cards during flutter modeling, particularly manual sorting of aerodynamic surfaces and manual extraction of interpolation node data. It unifies modal and flutter modeling, achieving a graphical and procedural approach to flutter analysis, effectively addressing the technical problem of independent modeling in existing aircraft modal and flutter analyses. Attached Figure Description

[0017] The accompanying drawings, which form part of this specification, are provided to further illustrate embodiments of the invention and, together with the textual description, explain the principles of the invention. It is obvious that the drawings described below are merely some embodiments of the invention, and those skilled in the art can obtain other drawings based on these drawings without any creative effort.

[0018] Figure 1 A schematic diagram of the flutter analysis and modeling method provided according to a specific embodiment of the present invention is shown;

[0019] Figure 2 A flowchart illustrating the process of determining control parameters, pneumatic parameters, and operating conditions according to a specific embodiment of the present invention is shown.

[0020] Figure 3 A flowchart illustrating the process of creating the aerodynamic surfaces of an airfoil component according to a specific embodiment of the present invention is shown;

[0021] Figure 4A schematic diagram of the aerodynamic surface modeling rules for an aircraft body provided according to a specific embodiment of the present invention is shown. Detailed Implementation

[0022] It should be noted that, unless otherwise specified, the embodiments and features described in this application can be combined with each other. The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of them. The following description of at least one exemplary embodiment is merely illustrative and is in no way intended to limit the present invention or its application or use. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0023] It should be noted that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to limit the exemplary embodiments according to this application. As used herein, the singular form is intended to include the plural form as well, unless the context clearly indicates otherwise. Furthermore, it should be understood that when the terms "comprising" and / or "including" are used in this specification, they indicate the presence of features, steps, operations, devices, components, and / or combinations thereof.

[0024] Unless otherwise specifically stated, the relative arrangement, numerical expressions, and values ​​of the components and steps set forth in these embodiments do not limit the scope of the invention. It should also be understood that, for ease of description, the dimensions of the various parts shown in the drawings are not drawn to actual scale. Techniques, methods, and devices known to those skilled in the art may not be discussed in detail, but where appropriate, such techniques, methods, and devices should be considered part of the specification. In all examples shown and discussed herein, any specific values ​​should be interpreted as merely exemplary and not as limitations. Therefore, other examples of exemplary embodiments may have different values. It should be noted that similar reference numerals and letters in the following figures denote similar items; therefore, once an item is defined in one figure, it need not be further discussed in subsequent figures.

[0025] like Figures 1 to 4As shown in the figure, a flutter analysis modeling method is provided according to a specific embodiment of the present invention. The flutter analysis modeling method includes: establishing a structural model of the entire aircraft; performing finite element discretization on the structural model of the entire aircraft to obtain a modal analysis model of the entire aircraft; based on the modal analysis model of the entire aircraft, selecting fuselage interpolation nodes on the fuselage and wing interpolation nodes on the wing surfaces of the entire aircraft; determining the analysis conditions according to the set Mach number, determining the control parameters, aerodynamic parameters, and coordinate system, and generating post-processing keywords; sequentially selecting key nodes of the wing-type finite element model in the modal analysis model of the entire aircraft, establishing wing-type aerodynamic surfaces according to the selected key nodes and the set spanwise and chordwise node numbers, and interpolating the wing-type aerodynamic surfaces using wing interpolation nodes to complete the modeling of the wing-type aerodynamic surfaces; establishing body-type aerodynamic surfaces, and interpolating the body-type aerodynamic surfaces using fuselage interpolation nodes to complete the modeling of the body-type aerodynamic surfaces; and completing the construction of the flutter analysis model of the entire aircraft based on the wing-type aerodynamic surface model and the body-type aerodynamic surface model.

[0026] This configuration provides a flutter analysis modeling method. Flutter analysis and modal analysis are performed within the same software environment. Furthermore, flutter analysis modeling can incorporate interpolation nodes and other information from modal analysis, enabling seamless integration between the two methods. This facilitates data interaction between the modal and flutter models, improving data transmission and modeling efficiency, and avoiding manual modification and modeling of flutter analysis model cards. Therefore, the flutter analysis modeling method provided by this invention solves the problems of manual modification of cards during flutter modeling, particularly manual sorting of aerodynamic surfaces and manual extraction of interpolation node data. It unifies modal and flutter modeling, achieving a graphical and procedural approach to flutter analysis, effectively addressing the technical problem of independent modeling in existing aircraft modal and flutter analyses.

[0027] In this invention, to complete the flutter analysis modeling, it is first necessary to establish a structural model of the entire aircraft. This structural model is then discretized using the finite element method (FEM) to obtain a modal analysis model of the entire aircraft. As a specific embodiment of this invention, the structural model of the entire aircraft is a 3D structural model, which can be constructed using software such as UG. The structural model includes structural components and equipment. Based on the structural model, mass center of mass, and moment of inertia data of the entire aircraft, the structural model is discretized using the finite element method in Hypermesh, transforming it into a finite element model suitable for finite element analysis, i.e., the modal analysis model of the entire aircraft. After constructing the modal analysis model of the entire aircraft, modal solution parameters are set, and modal solving is performed.

[0028] Furthermore, after obtaining the modal analysis model of the entire aircraft, interpolation nodes can be selected on the fuselage and wing surfaces based on the model. In this invention, selecting interpolation nodes on the fuselage specifically includes: selecting the intersection generatrix of the first plane of the entire aircraft with the uppermost side of the fuselage as the first generatrix node; selecting the intersection generatrix of the second plane of the entire aircraft with the leftmost side of the fuselage as the second generatrix node; and selecting the intersection generatrix of the second plane of the entire aircraft with the rightmost side of the fuselage as the third generatrix node, with the first and second planes perpendicular to each other. In this invention, a coordinate system is established with the aircraft's center of mass O as the origin, where the Z-axis is the vertical axis, the X-axis is the axis along the aircraft's axial direction, and the Y-axis is the horizontal axis, with the X, Y, and Z axes being mutually perpendicular. The first plane refers to the plane formed by the X-axis and Z-axis, and the second plane refers to the plane formed by the X-axis and Y-axis. Selecting interpolation nodes on the entire aircraft's wing surface specifically includes selecting nodes on the wing surface whose stiffness exceeds a set stiffness threshold as interpolation nodes. As a specific embodiment of the invention, the nodes on the wing surface whose stiffness exceeds the set stiffness threshold include wing frame nodes and stiffening rib nodes. After determining the airframe interpolation nodes and wing surface interpolation nodes, multiple node SET sets can be numbered sequentially according to the order 31002, 32002, 33002, etc.

[0029] Furthermore, after determining the structural interpolation nodes, the analysis conditions can be determined based on the set Mach number, along with the control parameters, aerodynamic parameters, and coordinate system, generating post-processing keywords. In this invention, each Mach number Ma represents an analysis condition. After determining the Mach number, the analysis condition and its output are automatically determined, generating execution control parameters. Specifically, the control parameters include modal input formats and mode output formats. Modal input formats include MSC, FREE, and ABAQUS, while mode output formats include .FO6, .DET, and *.UNV. Aerodynamic parameters include reference span, reference chord length, and reference area. Coordinate systems include local coordinate systems and Cartesian coordinate systems. Post-processing keywords include the output Vg (velocity-damping) / Vf (velocity-frequency) plots, flutter modes, interpolation modes, etc.

[0030] After determining the control parameters, aerodynamic parameters, and operating conditions, key nodes of the finite element model of the airfoil in the modal analysis model of the entire aircraft can be selected sequentially. Based on the selected key nodes and the set number of spanwise and chordwise nodes, airfoil aerodynamic surfaces are established. Airfoil interpolation nodes are used to interpolate the airfoil aerodynamic surfaces to complete the airfoil aerodynamic surface modeling. As a specific embodiment of this invention, the key nodes of the airfoil finite element model include the root chord leading edge point, root chord trailing edge point, sharp chord leading edge point, and sharp chord trailing edge point. A list file of the airfoil structure (containing identifiers and the number of spanwise / chordwise aerodynamic mesh nodes) is established, with the data format shown in Table 1. The key nodes and the number of spanwise and chordwise nodes set in Table 1 are used to establish the airfoil aerodynamic surfaces. Simultaneously, airfoil interpolation nodes are read in Hypermesh and interpolated with the corresponding airfoil aerodynamic surfaces to complete the airfoil aerodynamic surface modeling. The code related to the airfoil aerodynamic surface module is as follows: Figure 2 As shown.

[0031] Table 1 Example of a list file for wing-type structures

[0032] N number Identifier Number of nodes Number of chord nodes 1 zdy1 10 5 2 ydy1 10 5 3 zpw1 15 12 。 。。 。。 。。

[0033] Furthermore, after completing the modeling of the airfoil aerodynamic surfaces, the body aerodynamic surfaces can be established. The body interpolation nodes are used to interpolate the body aerodynamic surfaces to complete their modeling. Specifically, in this invention, body aerodynamic surfaces are established using arbitrary cross-sections (ITYPE=3), circles (ITYPE=1), or ellipses (ITYPE=2). As a specific embodiment of this invention, all body aerodynamic surfaces are modeled using arbitrary cross-sections (ITYPE=3). In programmatic implementation, directly using arbitrary cross-sections (ITYPE=3) can accommodate all cross-sectional shapes, providing universality. When defining cross-sectional coordinates, they are defined counter-clockwise along the -X-axis, with the first point located at the maximum position on the -Z-axis. A finite element model is established based on the aircraft's aerodynamic shape, with each cross-section having the same number of nodes. By identifying the node coordinates of identical cross-sections, the body aerodynamic surfaces are sorted according to the above rules. The relevant code for the body aerodynamic surface modeling module is as follows: Figure 3 As shown. After completing the modeling of the body-type aerodynamic surfaces, the flutter analysis model of the entire aircraft was constructed based on the wing-type aerodynamic surface model and the body-type aerodynamic surface model.

[0034] Furthermore, in this invention, after constructing the flutter analysis model, the flutter analysis modeling method also includes: completing the unsteady aerodynamic forces and flutter solution for the entire aircraft based on the flutter analysis model. Specifically, in this invention, the ZONA6 method is used to solve the unsteady aerodynamic forces; when the Mach number Ma is greater than 1, the ZONA7U method is used to solve the unsteady aerodynamic forces; the flutter analysis modeling method uses the G method to solve the flutter equations.

[0035] To gain a further understanding of the present invention, the following description is provided in conjunction with... Figures 1 to 4 The flutter analysis and modeling method provided by this invention will be described in detail.

[0036] like Figures 1 to 4 As shown, a flutter analysis and modeling method is provided according to a specific embodiment of the present invention. The method is run in Hypermesh using TCL / TK language and includes a module for executing control parameters, aerodynamic parameters and operating conditions, a module for modeling airfoil aerodynamic surfaces, a module for modeling body aerodynamic surfaces, a module for unsteady cleanup and flutter solution, and a post-processing module. The method specifically includes the following steps.

[0037] Step 1: Full-vehicle dynamics modeling. A structural model of the entire aircraft is established, and this model is discretized using the finite element method (FEM) to obtain a modal analysis model. In this embodiment, the structural model is a 3D model, which can be built using software such as UG. The structural model includes structural components and equipment. Based on the structural model, mass center of mass, and moment of inertia data of the entire aircraft, the structural model is discretized using the finite element method in Hypermesh, transforming it into a finite element model suitable for finite element analysis, i.e., the modal analysis model of the entire aircraft. After constructing the modal analysis model, modal solution parameters are set, and modal solutions are performed.

[0038] Step two: Determine the structural interpolation nodes. Based on the full aircraft modal model, establish a set of structural interpolation nodes (SETs). For the fuselage, select nodes from three busbars: the upper busbar, left busbar, and right busbar. For the wing surfaces, select nodes with good wing stiffness, such as the positions of the wing frame and stiffeners. Multiple structural interpolation node sets are numbered sequentially according to a set order, for example, 31002, 32002, 33002, etc.

[0039] Step 3: Determine the control parameters, aerodynamic parameters, and operating conditions. Based on the input Mach data, determine the analysis operating conditions, including transonic point conditions, maximum dynamic pressure point conditions, etc. Each Mach number (Ma) represents one analysis operating condition. Generate the analysis operating conditions, execution control and aerodynamic parameters, and coordinate system for the flutter analysis, and generate post-processing keywords. In this embodiment, control parameters include modal input formats and mode output formats. Modal input formats include MSC, FREE, and ABAQUS; mode output formats include .FO6, .DET, and *.UNV. Aerodynamic parameters include reference span, reference chord length, and reference area; coordinate systems include local coordinate systems and Cartesian coordinate systems. Post-processing keywords include the output Vg plot (velocity-damping plot) / Vf plot (velocity-frequency plot), flutter modes, interpolation modes, etc.

[0040] Step 4: Establish the aerodynamic surface model of the airfoil component. Create a list file of the airfoil structure (including identifiers and the number of spanwise / chordwise aerodynamic mesh nodes), as shown in Table 1. Select the key nodes (root chord leading edge, root chord trailing edge, sharp chord leading edge, and sharp chord trailing edge) of the airfoil finite element model one by one. Use these key nodes and the spanwise and chordwise node numbers set in Table 1 to establish the airfoil aerodynamic surfaces. Simultaneously, read the airfoil interpolation nodes in Hypermesh and interpolate them with the corresponding airfoil aerodynamic surfaces to complete the airfoil aerodynamic surface modeling.

[0041] Step 5: Establish the body-type aerodynamic surfaces. All body-type aerodynamic surfaces are modeled using arbitrary cross-sections (ITYPE=3). When defining the cross-section coordinates, the view is along the -X-axis, defined counter-clockwise, with the first point located at the maximum position on the -Z-axis. A finite element model is established based on the aircraft's aerodynamic shape. Each cross-section has the same number of nodes. By identifying the node coordinates of identical cross-sections, the body-type aerodynamic surfaces are ordered according to the above rules. The relevant code for the body-type aerodynamic surface modeling module is as follows: Figure 3 As shown.

[0042] Step six: Construct the flutter analysis model of the entire aircraft based on the airfoil aerodynamic surface model and the body aerodynamic surface model. Solve the unsteady aerodynamic forces and flutter problems of the entire aircraft based on the flutter analysis model. Specifically, in this embodiment, the ZONA6 method is used to solve the unsteady aerodynamic forces; when the Mach number Ma is greater than 1, the ZONA7U method is used to solve the unsteady aerodynamic forces; the G-method is used to solve the flutter equations for flutter analysis modeling.

[0043] In summary, this invention proposes a flutter analysis modeling method. This method integrates flutter analysis and modal analysis in the same software environment, and incorporates interpolation nodes and other information from modal analysis. This enables seamless integration of modal and flutter analysis, facilitating data interaction between the two models, improving data transmission and modeling efficiency, and avoiding manual modification and modeling of flutter analysis models. Furthermore, this method generates body-like aerodynamic surfaces based on the finite element mesh of the aircraft's aerodynamic shape, allowing for rapid generation and node sorting of these surfaces. Therefore, the flutter analysis modeling method provided by this invention solves the problems of manual modification of cards during flutter modeling, especially manual sorting of aerodynamic surfaces of body types and manual extraction of interpolation node data. During flutter modeling, it automatically reads interpolation node information and generates control parameters, aerodynamic parameters, fuselage / wing surface aerodynamic surfaces, flutter analysis settings, etc., realizing the unification of modal modeling and flutter modeling, realizing the interface and programmability of flutter analysis, and effectively solving the technical problem that modeling in aircraft modal analysis and flutter analysis is independent in the prior art.

[0044] For ease of description, spatial relative terms such as "above," "on top of," "on the upper surface of," "above," etc., are used herein to describe the spatial positional relationship of a device or feature as shown in the figures to other devices or features. It should be understood that spatial relative terms are intended to encompass different orientations in use or operation beyond the orientation of the device as described in the figures. For example, if the device in the figures were inverted, a device described as "above" or "on top of" other devices or structures would subsequently be positioned as "below" or "under" other devices or structures. Thus, the exemplary term "above" can include both "above" and "below." The device may also be positioned in other different ways (rotated 90 degrees or in other orientations), and the spatial relative descriptions used herein will be interpreted accordingly.

[0045] Furthermore, it should be noted that the use of terms such as "first" and "second" to define components is merely for the purpose of distinguishing the corresponding components. Unless otherwise stated, the above terms have no special meaning and therefore should not be construed as limiting the scope of protection of this invention.

[0046] The above description is merely a preferred embodiment of the present invention and is not intended to limit the invention. Various modifications and variations can be made to the present invention by those skilled in the art. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A flutter analysis and modeling method, characterized in that, The flutter analysis and modeling method includes: A structural model of the entire aircraft is established, and the structural model of the entire aircraft is discretized by finite element method to obtain the modal analysis model of the entire aircraft. Based on the modal analysis model of the entire aircraft, fuselage interpolation nodes are selected on the fuselage of the entire aircraft, and wing surface interpolation nodes are selected on the wing surface of the entire aircraft. The analysis conditions are determined based on the set Mach number, the control parameters, aerodynamic parameters and coordinate system are determined, and post-processing keywords are generated. Key nodes of the finite element model of the wing type in the modal analysis model of the whole aircraft are selected in sequence. The wing type aerodynamic surface is established according to the selected key nodes and the set spanwise and chordwise node numbers. The wing type aerodynamic surface is interpolated using the wing interpolation nodes to complete the modeling of the wing type aerodynamic surface. A body-type aerodynamic surface is established, and the body interpolation node is used to interpolate the body-type aerodynamic surface to complete the modeling of the body-type aerodynamic surface; Based on the airfoil aerodynamic surface model and the body aerodynamic surface model, the flutter analysis model of the entire aircraft is constructed. The selection of interpolation nodes on the aircraft fuselage specifically includes: selecting the intersection generatrix of the first plane of the entire aircraft with the uppermost side of the fuselage as the first generatrix node; selecting the intersection generatrix of the second plane of the entire aircraft with the leftmost side of the fuselage as the second generatrix node; selecting the intersection generatrix of the second plane of the entire aircraft with the rightmost side of the fuselage as the third generatrix node. The first plane is perpendicular to the second plane. A coordinate system is established with the aircraft's center of mass O as the origin, where the Z-axis is the vertical axis, the X-axis is the axis along the aircraft's axis, and the Y-axis is the horizontal axis. The X, Y, and Z axes are mutually perpendicular. The first plane refers to the plane formed by the X and Z axes, and the second plane refers to the plane formed by the X and Y axes.

2. The flutter analysis and modeling method according to claim 1, characterized in that, Selecting wing surface interpolation nodes on the wing surface of the entire aircraft specifically includes: selecting nodes on the wing surface of the entire aircraft whose stiffness exceeds a set stiffness threshold as wing surface interpolation nodes.

3. The method of claim 2, wherein, The control parameters include modal input format and mode output format; the aerodynamic parameters include reference span, reference chord length, and reference area; the coordinate system includes a local coordinate system and a Cartesian coordinate system; the nodes on the wing surface of the entire aircraft whose stiffness exceeds the set stiffness threshold include wing frame nodes and stiffener nodes.

4. The flutter analysis and modeling method according to claim 3, characterized in that, The post-processing keywords include Vg graph, Vf graph, flutter mode, and interpolation mode.

5. The flutter analysis and modeling method according to any one of claims 1 to 4, characterized in that, The key nodes of the finite element model of the wing include the root chord leading edge point, the root chord trailing edge point, the pointed chord leading edge point, and the pointed chord trailing edge point.

6. The method of chattering analysis modeling according to claim 5, wherein, The creation of a volume-type aerodynamic surface specifically includes: creating the volume-type aerodynamic surface using an arbitrary cross-section, or in a circular or elliptical shape.

7. The method of chattering analysis modeling according to claim 6, wherein, After completing the flutter analysis model, the flutter analysis modeling method further includes: solving the unsteady aerodynamic forces and flutter of the entire aircraft based on the flutter analysis model.

8. The flutter analysis and modeling method according to claim 7, characterized in that, When the Mach number Ma is less than 1, the ZONA6 method is used to solve the unsteady aerodynamic forces; when the Mach number Ma is greater than 1, the ZONA7U method is used to solve the unsteady aerodynamic forces; the flutter analysis modeling method uses the G method to solve the flutter equation.

9. The flutter analysis and modeling method according to claim 8, characterized in that, The process of discretizing the structural model of the entire aircraft using finite element method to obtain the modal analysis model of the entire aircraft specifically includes: based on the structural model, mass center of mass, and moment of inertia data of the entire aircraft, performing finite element discretization on the structural model of the entire aircraft to obtain the modal analysis model of the entire aircraft.

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