A topology-size sequential aeroelastic optimization method for aircraft structures

Abstract

This invention belongs to the field of aircraft optimization technology and discloses a method for sequential aeroelastic optimization of aircraft structure topology and dimensions. The method comprises: constructing an initial structural model based on aircraft design parameters, inputting the initial flight state, aerodynamic shape, and a preset temperature field; obtaining aerodynamic force data through aeroelastic analysis as a load input to the model; then performing topological optimization, adjusting the grid cell density, and updating the structure file; iterating this process a predetermined number of times to obtain a preliminary optimized structure; and finally, performing dimensional optimization for beam and rib thicknesses and fine-tuning the structure to obtain the optimal aircraft structure. Through multiple rounds of iterative optimization, the technical solution described in this invention significantly reduces structural compliance and achieves a more reasonable structural distribution that meets engineering requirements.

Description

Technical Field

[0001] The present invention belongs to the technical field of aircraft optimization, and in particular relates to an aircraft structure topology-size sequential aeroelastic optimization method. Background Art

[0002] A hypersonic aircraft refers to an aircraft that flies at a Mach number exceeding five times the speed of sound. Its main activity space is the atmosphere and trans-atmospheric space. This type of aircraft has huge economic and military value.

[0003] Structural optimization design, involving computational mechanics, mathematical programming, computer science, and other engineering disciplines, is a key research area in modern structural design. Hypersonic vehicle structural design must meet both strength and stiffness requirements and mass and volume constraints, resulting in multiple optimization objectives and complex constraints. However, traditional aircraft structural optimization methods involve complex aerodynamic calculations and ignore the effects of temperature on the structure, resulting in unrealistic structural distributions. Summary of the Invention

[0004] The purpose of the present invention is to provide a method for sequential aeroelastic optimization of aircraft structure topology and size to solve the problems existing in the above-mentioned prior art.

[0005] To achieve the above objectives, the present invention provides a method for sequential aeroelastic optimization of aircraft structure topology and dimensions, comprising:

[0006] Step S1: constructing an initial structural model of the aircraft based on the design parameters of the aircraft; inputting initial condition data and preset temperature field data into the initial structural model, wherein the initial condition data includes the initial flight state and aerodynamic shape parameters of the aircraft;

[0007] Step S2: performing aeroelastic analysis on the initial structural model, and outputting current aerodynamic force data as input load of the initial structural model;

[0008] Step S3: After the aeroelastic analysis is completed, topology optimization is performed on the initial structural model to obtain a current cell density value of each grid cell in the initial structural model, and the current cell density value that meets the preset rules is screened to update the structure file of the initial structural model;

[0009] Repeat steps S2 to S3 to perform iterative optimization until the preset number of iterations is reached and then stop the iteration to output preliminary optimized structural data of the aircraft;

[0010] Step S4: Optimizing the beam thickness and rib thickness in the preliminary optimized structure data to obtain the best optimized structure of the aircraft.

[0011] Optionally, step S2 specifically includes:

[0012] Step S21: Calculating surface aerodynamic data of the aircraft surface based on the initial condition data and the preset temperature field data in combination with the third-order piston theory;

[0013] Step S22: calculating the inertial load data of the initial structural model, and calculating the node aerodynamic data of the initial structural model based on the surface aerodynamic data and the inertial load data;

[0014] Step S23: performing a static analysis on the initial structural model based on the node aerodynamic data to obtain current node displacement data;

[0015] Step S24: updating the aerodynamic mesh of the initial structural model based on the current node displacement data, and calculating the updated aerodynamic force data and model structure displacement data;

[0016] Repeat steps S23 to S24 to perform iterative updates until the preset iterative conditions are met and then stop the iteration and output the current aerodynamic data.

[0017] Optionally, step S22 specifically includes:

[0018] The surface aerodynamic force data is loaded into the initial structural model in the form of nodal force using the inertial load data as a load boundary condition to obtain nodal aerodynamic force data.

[0019] Optionally, step S3 specifically includes:

[0020] Step S31: taking the minimum structural compliance as the topology optimization objective function;

[0021] Step S32: performing topology optimization on the initial structural model based on the topology optimization objective function and the corresponding constraint conditions to obtain the current cell density value of each grid cell in the initial structural model.

[0022] Step S33: screening current unit density values ​​greater than a preset threshold, and updating the structure file of the initial structure model based on the screened current unit density values;

[0023] Step S34: After the structural file is updated, return to step S2 to perform aeroelastic analysis, repeat steps S2 to S3 for iterative optimization, and stop iteration after reaching a preset number of iterations, and output preliminary optimized structural data of the aircraft.

[0024] Optionally, the constraints corresponding to the topology optimization objective function include volume ratio constraints, minimum structure size constraints and displacement constraints.

[0025] Optionally, before performing step S4, the process further includes performing smoothing processing on the preliminary optimized structural data, specifically including:

[0026] The preliminary optimized structural data is smoothed based on engineering experience, and size optimization is performed based on the smoothed preliminary optimized structural data.

[0027] Optionally, step S4 specifically includes:

[0028] The objective function of size optimization is to minimize the total mass of the structure;

[0029] Based on the size optimization objective function and the corresponding constraint variables, the beam thickness and rib thickness in the preliminary optimized structural data are iteratively optimized until the iteration is stopped after reaching a preset number of iterations, and the optimal optimized structure of the aircraft is output.

[0030] Optionally, the constraint variables corresponding to the size optimization objective function include deformation constraints, stress constraints and strain constraints.

[0031] The technical effects of the present invention are:

[0032] In terms of structural topology optimization, this invention significantly reduces structural compliance through multiple rounds of iterative optimization, resulting in a more reasonable structural distribution that meets engineering requirements. After completing topology optimization, this invention optimizes beam and rib thickness to achieve more detailed and reasonable structural dimensions, reducing structural weight and manufacturing costs while meeting flight performance requirements. By setting a suitable structural temperature field to simulate the aerodynamic heating effects of sustained flight, this invention can significantly reduce calculation time while also accounting for the impact of temperature on the maximum structural deformation. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0034] The accompanying drawings, which constitute part of this application, are intended to provide a further understanding of this application. The exemplary embodiments and descriptions of this application are intended to explain this application and do not constitute an improper limitation on this application. In the accompanying drawings:

[0035] Figure 1 Schematic diagram of the optimization process in an embodiment of the present invention;

[0036] Figure 2 Schematic diagram of the optimization model in the embodiment of the present invention; wherein, Figure 2(a) is a schematic diagram of a rudder model in an embodiment of the present invention, Figure 2 (a) is a schematic diagram of an airfoil model in an embodiment of the present invention;

[0037] Figure 3 is the topology optimization result in the embodiment of the present invention; wherein, Figure 3 (a) is a schematic diagram of the rudder topology optimization results in an embodiment of the present invention. Figure 3 (a) is a schematic diagram of the airfoil topology optimization results in an embodiment of the present invention;

[0038] Figure 4 is a diagram of the structure after smoothing in an embodiment of the present invention; wherein, Figure 4 (a) is a schematic diagram of the rudder surface smoothing result in an embodiment of the present invention, Figure 4 (a) is a schematic diagram of the airfoil smoothing result in an embodiment of the present invention;

[0039] Figure 5 Schematic diagram of design variables in the embodiment of the present invention; wherein, Figure 5 (a) is a schematic diagram of the rudder design variables in an embodiment of the present invention. Figure 5 (a) is a schematic diagram of airfoil design variables in an embodiment of the present invention;

[0040] Figure 6 is a structural deformation diagram before optimization in an embodiment of the present invention; wherein, Figure 6 (a) is a deformation diagram of the rudder surface before optimization in an embodiment of the present invention. Figure 6 (a) is a deformation diagram of the airfoil before optimization in an embodiment of the present invention;

[0041] Figure 7 is a deformation diagram of the optimized structure in the embodiment of the present invention; wherein, Figure 7 (a) is a deformation diagram of the rudder surface after optimization in an embodiment of the present invention. Figure 7 (a) is a deformation diagram of the airfoil after optimization in an embodiment of the present invention. DETAILED DESCRIPTION

[0042] Various exemplary embodiments of the present invention will now be described in detail. This detailed description should not be considered as limiting the present invention, but rather as a more detailed description of certain aspects, features, and embodiments of the present invention.

[0043] It should be understood that the terms described herein are intended only to describe particular embodiments and are not intended to limit the present invention. In addition, for numerical ranges herein, it should be understood that each intermediate value between the upper and lower limits of the range is also specifically disclosed. Each smaller range between any intermediate value within a stated value or stated range and any other stated value or intermediate value within the stated range is also encompassed by the present invention. The upper and lower limits of these smaller ranges may be independently included or excluded within the scope.

[0044] Unless otherwise indicated, all technical and scientific terms used herein have the same meaning as commonly understood by those skilled in the art. Although only preferred methods are described herein, any method similar or equivalent to that described herein may also be used in the practice or testing of the present invention. All documents mentioned in this specification are incorporated by reference to disclose and describe the methods associated with the documents. In the event of any conflict with any incorporated document, the contents of this specification shall prevail.

[0045] It will be apparent to those skilled in the art that various modifications and variations may be made to the specific embodiments of the present invention without departing from the scope or spirit of the invention. Other embodiments will be apparent to those skilled in the art from the present invention. The present description and examples are intended to be illustrative only.

[0046] The words “include,” “including,” “have,” “contain,” etc. used in this article are open-ended terms, meaning including but not limited to.

[0047] It should be noted that, in the absence of conflict, the embodiments and features of the embodiments in this application can be combined with each other. The present application will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0048] Example 1

[0049] like Figure 1 - Figure 7 As shown, this embodiment provides a method for sequential aeroelastic optimization of aircraft structure topology and size, including:

[0050] Step S1: constructing an initial structural model of the aircraft based on the design parameters of the aircraft; inputting initial condition data and preset temperature field data into the initial structural model, wherein the initial condition data includes the initial flight state and aerodynamic shape parameters of the aircraft;

[0051] Step S2: performing aeroelastic analysis on the initial structural model, and outputting current aerodynamic force data as input load of the initial structural model;

[0052] Step S3: After the aeroelastic analysis is completed, topology optimization is performed on the initial structural model to obtain a current cell density value of each grid cell in the initial structural model, and the current cell density value that meets the preset rules is screened to update the structure file of the initial structural model;

[0053] Repeat steps S2 to S3 to perform iterative optimization until the preset number of iterations is reached and then stop the iteration to output preliminary optimized structural data of the aircraft;

[0054] Step S4: Optimizing the beam thickness and rib thickness in the preliminary optimized structure data to obtain the best optimized structure of the aircraft.

[0055] The aeroelastic topology optimization method for hypersonic vehicles mainly involves three aspects: aerodynamic force, structural static analysis, and structural topology optimization. The two major disciplines of aerodynamic force analysis and structural static analysis constitute the aeroelastic analysis module. In aeroelastic analysis, only static aeroelastic analysis is considered; at the level of calculation methods, this method adopts a more flexible loose coupling method; in terms of coupling relationships, since the influence of weak coupling on three-field coupling is negligible, this method only considers strong coupling relationships. In order to ensure the accuracy of the analysis and improve the analysis speed, for aerodynamic force calculation, an engineering method combining shock wave / expansion wave theory and local piston theory is selected to quickly give the hypersonic flow field parameters; the aerodynamic force at each node of the aerodynamic grid is obtained by interpolation method, and the static aeroelastic deformation is obtained by the static analysis module of Nastran. Then, the aerodynamic force and aeroelastic deformation are recalculated until the displacement converges, and the aerodynamic force at this time is output.

[0056] Topology optimization analysis can be performed using Nastran sol200. Minimum structural compliance is set as the objective, and the load is the aerodynamic load output from the aeroelastic analysis. Appropriate volume constraints are pre-set, along with parameters such as the maximum displacement constraint, minimum structural dimensions, initial density, and density variation amplitude. After obtaining the results of a round of topology optimization, the aerodynamic forces are recalculated and used as new load inputs to achieve iterative optimization. After topology optimization is complete, component size optimization is performed based on the smoothed structure. Nastran sol200 allows for direct definition of the objective function, initial values ​​and upper and lower limits of design variables, design constraints, and the maximum number of iterations, allowing for the optimal size to be achieved through multiple rounds of iteration.

[0057] like Figure 1As shown in the optimization flow chart, the aerodynamic calculation part is entered at the beginning. First, the geometric shape and flight state (altitude, Mach number, angle of attack, etc.) of the hypersonic component are determined. Taking into account that hypersonic flight will have a certain aerodynamic heating effect, the initial temperature field of the structure is pre-set, and the aerodynamic pressure on the surface of the hypersonic aircraft component is calculated based on the third-order piston theory. The inertial load is calculated according to the pre-set overloads in the x, y, and z directions. The aerodynamic load is converted from the surface pressure to the node, and the aerodynamic force is obtained by multiplying the pressure by the area formed by the four nodes, and then the force is evenly distributed to the four nodes.

[0058] Aerodynamic calculation method: Assume that in the time dt, the speed of the piston is v n Change to v n +dv n , the distance the disturbance propagates is a·dt, then the mass of the disturbed gas is paSdt (ρ is the air density, a is the speed of sound, S is the piston area), and the momentum change of the gas is ρaSdt·dv n On the other hand, the pressure change is dp, and the impulse generated by the piston during the time is Sdp·dt. According to the principle of conservation of momentum, we have:

[0059] Sdp·dt=ρaSdt·dv n

[0060] Right now:

[0061] dp=ρadv n

[0062] When the piston moves forward in the cylinder at a speed of |V n |< ∞ When (which is satisfied in most cases), the disturbance generated is negligible compared with the propagation process of the airflow, and the piston motion process can be considered as an isentropic process.

[0063] From the isentropic formula and the local sonic formula, we can get:

[0064]

[0065] Where γ is the gas specific heat ratio (generally taken as 1.4);

[0066] Integrating both sides of the above equation yields:

[0067]

[0068] Where C is the integration constant, and from the far-field boundary condition, p=p ∞ , V ∞ =0, then the integration constant C can be determined as:

[0069] ​

[0070] Substituting back into the above formula, the local instantaneous pressure acting on the piston surface can be expressed as:

[0071]

[0072] Taylor expansion is performed on the above formula. When only the first-order term is retained, it is called the first-order piston theory. When the second-order, third-order and higher-order terms are retained, they are called the second-order, third-order and higher-order piston theories. The expression of the pressure coefficient of the third-order piston theory can be obtained as follows:

[0073]

[0074] When encountering strong three-dimensional effects, such as large relative thickness, large object surface inclination, or large calculated attack angle, the local airflow parameters on the object surface obtained by steady flow calculation can be used to replace the far-field airflow parameters in the piston theory, which is the so-called local flow piston theory:

[0075]

[0076] The subscript local represents the local airflow parameters.

[0077] Then, the node displacement under this aerodynamic force is obtained through static analysis. Since the node displacement will affect the geometric shape of the component surface, the change in geometric shape will affect the aerodynamic force calculation results. Therefore, the aerodynamic force and displacement are recalculated. After the iteration converges, the aeroelastic analysis module ends and the aerodynamic force at this time is output.

[0078] Statics analysis method:

[0079] The static equilibrium equation of a three-dimensional linear elastic structure is:

[0080]

[0081] The matrix format is abbreviated as: LTσ+b=0;

[0082] Where, is L T Differential operator, σ is the stress vector, b is the external force vector, and the corresponding expression is:

[0083]

[0084] σ=[σ xx σ yy σ zz σ xy σ xz σ yz ] T

[0085] b=[b x by b z ] T

[0086] Based on the above equilibrium equations, and according to the principle of virtual work or the Galerkin weighted residual method, the finite element overall equation for structural static analysis can be obtained as follows:

[0087] KU=F

[0088] Where K is the overall stiffness matrix, U is the overall displacement vector, and F is the overall load vector obtained after integrating various boundary conditions.

[0089] The overall load vector F is mainly composed of the following three aspects:

[0090]

[0091] Among them, N is the element shape function, Q is the volume force vector, and b is the surface force vector.

[0092] The overall stiffness matrix K is mainly obtained by assembling the element stiffness matrix of each element:

[0093]

[0094] Where B is the strain matrix, which can be expressed by the geometric equation ε=Lu=LNu e =Bu e get;

[0095]

[0096] ε=[ε xx ε yy ε zz γ xy γ xz γ yz ] T

[0097] u e =[u 1c u 1y u 1z … u nx u ny u nz ] T

[0098]

[0099] Among them, u e is the displacement of n nodes in the unit, φ1 to φ n is the n elements that make up the unit shape function. D is the stress-strain relationship matrix in the constitutive equation σ=Dε:

[0100]

[0101] Where E is Young's modulus and v is Poisson's ratio.

[0102] After solving the overall displacement results, the strain and stress distribution of the structure can be obtained according to the above geometric equations and constitutive equations.

[0103] After completing the aeroelastic analysis, enter the topology optimization analysis. In the TOPVAR card, pre-set the lower limit, initial value, and maximum variation of the unit density. In the master control file, set the optimization objective function, volume ratio constraint, minimum structural size, displacement constraint and other parameters. By continuously adjusting the unit density, the structure file is updated and the aerodynamic force is recalculated as the input load. After multiple rounds of iterations, the optimal force transmission path is obtained, such as Figure 3 As shown in the figure. After topological optimization, a new structure is obtained. The thickness of the beam and rib sections of the new structure is selected for size optimization. The objective function is set to minimize the total mass of the structure while satisfying certain displacement, stress, and strain constraints, and finally the optimal structural form is obtained. The optimized structure can reduce the total mass of the structure by more than 10% while ensuring that the maximum deformation of the load-bearing structure remains unchanged. The structural deformation diagrams before and after optimization are shown in the figure. Figure 6 and Figure 7 shown.

[0104] The calculation process of topology optimization analysis problem:

[0105] Assume that the density vector of each unit in the problem domain is x and the volume vector of each unit is v e , the compliance is C(x), then the corresponding topology optimization model is:

[0106]

[0107] Where F is the overall displacement vector, K is the overall stiffness matrix, F is the overall load vector; N is the total number of elements, p is the penalty factor, and U is the i and K i are the unit displacement vector and stiffness matrix respectively; v is the unit volume vector, V0 is the total volume of the entire problem domain when the density is 1, is the volume ratio constraint value; x min is the minimum element density, which is used to avoid the singularity of the overall stiffness matrix when the element density is 0. Generally, x min =0.001.

[0108] Size optimization problem description:

[0109] The mathematical model of the optimization problem can be expressed as:

[0110] minF(v)

[0111] stgj (v)≤0j=1,2,…n c ,

[0112] n c (v i ) lower ≤v i ≤(v i ) upper i=1,2,…,n d

[0113] Where F is the objective function of the optimization model, i.e., the lightest structural weight; v is the optimized design variable; g is the expression of the aeroelastic inequality constraint; n c is the number of inequality constraints; n d is the number of structural design variables; the subscripts lower and upper represent the lower and upper bounds of the structural design variable value range, respectively.

[0114] Based on the analytical flow chart, the aerodynamic / structural coupled analysis and optimization framework for hypersonic vehicle components primarily consists of an aeroelastic analysis module (aerodynamic force analysis and structural static analysis), a topology optimization module, and a sizing optimization module. The specific analysis begins by inputting the initial flight state and aerodynamic shape parameters into the Matlab program. Considering aerodynamic heating, a suitable temperature field must be pre-set. This can be achieved by writing the temperature of each node in a pch file. Introducing a temperature field has a certain impact on material parameters, primarily manifesting in a reduction in the material's Young's modulus. Once the input file is prepared, the inner loop of the aeroelastic module performs an iterative coupled analysis. The aerodynamic forces on the vehicle component surface are calculated based on third-order piston theory, and the structural displacements are calculated using Nastran Sol 101. After a single calculation, the node displacements in the generated pch file are read, the model's aerodynamic mesh is updated based on the node displacements, and the aerodynamic forces and structural displacements are recalculated until the structural displacements converge, completing the aeroelastic analysis module.

[0115] After completing the aeroelastic analysis module, enter the topology optimization module. The input files include the topology optimization master control file, the model structure file, and the topology variable file. The constraints and initial density input need to be set. After a round of topology optimization, the new density value of each unit is obtained. The unit density value in the DES file is read and filtered in the Matlab program. Units with density values ​​greater than the threshold will be retained and rewritten to the structure file. Then, re-enter the aeroelastic analysis module and update the aerodynamic load card. After repeated iterations, the preliminary form of the structure is obtained. Based on engineering experience, the structure after filtering the units is smoothed into a configuration that is more in line with engineering practice, and then enter the size optimization module.

[0116] Size optimization is mainly aimed at optimizing the thickness of beams and ribs of the smoothed structure (such as Figure 4 (as shown in the figure), the objective function is set to minimize the total structural mass while satisfying constraints such as deformation, stress, and strain, ultimately achieving the optimal structural form. Nastran sol200 is selected to define the design variables, objective function, and constraints for dimensional optimization. The maximum number of iterations is set by setting DOPTPRMDESMAX. When defining design variables, the initial size, upper and lower limits, and maximum variation of the design variables must be defined. After optimization is complete, search the generated F06 file for the keyword SUMMARY OF DESIGN CYCLEHISTORY to read the objective function values ​​and final values ​​of the design variables after each optimization step.

[0117] Comparing this embodiment with the CFD calculation results, the pressure, deformation, and stress calculation results of the hypersonic vehicle obtained using the engineering algorithm are basically consistent, proving the correctness and effectiveness of the aerodynamic calculation.

[0118] In terms of structural topology optimization, this embodiment significantly reduces structural flexibility through multiple rounds of iterative optimization, and can obtain a more reasonable structural distribution that conforms to engineering practice.

[0119] After the topology optimization is completed in this embodiment, more detailed and reasonable structural dimensions can be obtained by optimizing the thickness of the beams and ribs, which can reduce the structural weight and save manufacturing costs while meeting the flight performance.

[0120] This embodiment simulates the aerodynamic heating effect during continuous flight by setting an appropriate structural temperature field, which can significantly save calculation time and also consider the effect of temperature on the maximum deformation of the structure.

[0121] The above description is merely a preferred embodiment of the present application, but the scope of protection of the present application is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in the present application should be included in the scope of protection of the present application. Therefore, the scope of protection of the present application should be based on the scope of protection of the claims.

Claims

1. A method for sequential aeroelastic optimization of aircraft structure topology and size, characterized by: include: Step S1: constructing an initial structural model of the aircraft based on the design parameters of the aircraft; Inputting initial condition data and preset temperature field data into the initial structural model, wherein the initial condition data includes an initial flight state and aerodynamic shape parameters of the aircraft; Step S2: performing aeroelastic analysis on the initial structural model, and outputting current aerodynamic force data as input load of the initial structural model; The step S2 specifically includes: Step S21: Calculating surface aerodynamic data of the aircraft surface based on the initial condition data and the preset temperature field data in combination with the third-order piston theory; Step S22: calculating the inertial load data of the initial structural model, and calculating the node aerodynamic data of the initial structural model based on the surface aerodynamic data and the inertial load data; Step S23: performing a static analysis on the initial structural model based on the node aerodynamic data to obtain current node displacement data; Step S24: updating the aerodynamic mesh of the initial structural model based on the current node displacement data, and calculating the updated aerodynamic force data and model structure displacement data; Repeat steps S23 to S24 to perform iterative updates until a preset iterative condition is reached, and then stop the iteration and output the current aerodynamic data; Step S3: After the aeroelastic analysis is completed, topology optimization is performed on the initial structural model to obtain the current cell density value of each grid cell in the initial structural model, and the current cell density value that meets the preset rules is screened to update the structure file of the initial structural model; Repeat steps S2 to S3 to perform iterative optimization until the preset number of iterations is reached and then stop the iteration to output preliminary optimized structural data of the aircraft; Step S4: Optimizing the beam thickness and rib thickness in the preliminary optimized structure data to obtain the best optimized structure of the aircraft.

2. The method for sequential aeroelastic optimization of aircraft structure topology and size according to claim 1, characterized in that: The step S22 specifically includes: The surface aerodynamic force data is loaded into the initial structural model in the form of nodal force using the inertial load data as a load boundary condition to obtain nodal aerodynamic force data.

3. The method for sequential aeroelastic optimization of aircraft structure topology and size according to claim 1, characterized in that: The step S3 specifically includes: Step S31: taking the minimum structural compliance as the topology optimization objective function; Step S32: performing topology optimization on the initial structural model based on the topology optimization objective function and the corresponding constraint conditions to obtain the current cell density value of each grid cell in the initial structural model. Step S33: screening current unit density values ​​greater than a preset threshold, and updating the structure file of the initial structure model based on the screened current unit density values; Step S34: After the structural file is updated, return to step S2 to perform aeroelastic analysis, repeat steps S2 to S3 for iterative optimization, and stop iteration after reaching a preset number of iterations, and output preliminary optimized structural data of the aircraft.

4. The method for sequential aeroelastic optimization of aircraft structure topology and size according to claim 1, characterized in that: The constraints corresponding to the topology optimization objective function include volume ratio constraints, minimum structure size constraints and displacement constraints.

5. The method for sequential aeroelastic optimization of aircraft structure topology and size according to claim 1, characterized in that: Before step S4, the preliminary optimized structural data is smoothed, specifically including: The preliminary optimized structural data is smoothed based on engineering experience, and size optimization is performed based on the smoothed preliminary optimized structural data.

6. The method for sequential aeroelastic optimization of aircraft structure topology and size according to claim 1, characterized in that: The step S4 specifically includes: The objective function of size optimization is to minimize the total mass of the structure; Based on the size optimization objective function and the corresponding constraint variables, the beam thickness and rib thickness in the preliminary optimized structural data are iteratively optimized until the iteration is stopped after reaching a preset number of iterations, and the optimal optimized structure of the aircraft is output.

7. The method for sequential aeroelastic optimization of aircraft structure topology and size according to claim 6, characterized in that: The constraint variables corresponding to the size optimization objective function include deformation constraints, stress constraints and strain constraints.

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