Aircraft structure topology-size sequential aeroelasticity optimization method

Abstract

The invention belongs to the technical field of aircraft optimization, and discloses an aircraft structure topology-size sequential aeroelasticity optimization method, which comprises the following steps: constructing an initial structure model based on aircraft design parameters, and inputting an initial flight state, an aerodynamic configuration and a preset temperature field; aerodynamic force data is obtained through aeroelasticity analysis and serves as a load input model; then topological optimization is carried out, the density of grid cells is adjusted, and a structure file is updated; iterating the process to a preset number of times to obtain a preliminary optimization structure; and finally, carrying out size optimization aiming at the thicknesses of the beams and the ribs, and finely adjusting the structure to obtain the optimal structure of the aircraft. According to the technical scheme, through multiple rounds of iterative optimization, the structure flexibility is remarkably reduced, reasonable structure distribution can be obtained, and engineering practice is met.

Description

Technical Field

[0001] The 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 greater than five times the speed of sound. Its main activity space is the atmosphere and across the atmosphere. This type of aircraft has huge economic and military value.

[0003] Structural optimization design involves computational mechanics, mathematical programming, computer science and other engineering disciplines, and is an important research direction in the field of modern structural design. The structural design of hypersonic aircraft needs to meet both strength and stiffness design requirements and mass and volume constraints, and has the characteristics of multiple optimization targets and complex optimization constraints. However, the traditional aircraft structural optimization design method has complex aerodynamic calculations and ignores the influence of temperature on the structure, resulting in a structural distribution that does not conform to reality. Summary of the invention

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

[0005] To achieve the above object, the present invention provides a method for sequential aeroelastic optimization of aircraft structure topology-size, 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 an 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, the initial structural model is topologically optimized 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;

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

[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, the step S2 specifically includes:

[0012] Step S21: Calculating the 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 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 updating until the preset iterative condition is reached, then stop the iteration and output the current aerodynamic data.

[0017] Optionally, the step S22 specifically includes:

[0018] The surface aerodynamic 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 data.

[0019] Optionally, the step S3 specifically includes:

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

[0021] Step S32: performing topology optimization on the initial structure 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 structure 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 preliminary optimized structure data is further smoothed, specifically including:

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

[0027] Optionally, the 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 best 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, the present invention significantly reduces the structural compliance through multiple rounds of iterative optimization, and can obtain a more reasonable structural distribution that conforms to engineering practice. After the topology optimization is completed, the present invention can obtain more detailed and reasonable structural dimensions by optimizing the dimensions of the beam and rib thickness, which can reduce the structural weight and save manufacturing costs while meeting the flight performance. The present invention can greatly save calculation time by setting a suitable structural temperature field to simulate the aerodynamic heating effect during continuous flight, while also considering the effect of temperature on the maximum deformation of the structure. 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 drawings required for use in the embodiments will be briefly introduced below. 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 creative work.

[0034] The drawings constituting a part of the present application are used to provide a further understanding of the present application. The illustrative embodiments and descriptions of the present application are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

[0035] Figure 1 It is a schematic diagram of the optimization process in an embodiment of the present invention;

[0036] Figure 2 is a schematic diagram of an optimization model in an 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 topology optimization result of the control surface in an embodiment of the present invention, Figure 3 (a) is a schematic diagram of the airfoil topology optimization result 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 smoothing result of the rudder surface 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 control surface 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 should be understood as a more detailed description of certain aspects, features, and embodiments of the present invention.

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

[0044] Unless otherwise indicated, all technical and scientific terms used herein have the same meanings as those of ordinary skill in the art generally understood by the present invention. Although the present invention describes only preferred methods, any method similar or equivalent to that described herein may also be used in the implementation or testing of the present invention. All documents mentioned in this specification are incorporated by reference to disclose and describe the methods related to the documents. In the event of a conflict with any incorporated document, the content 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 description without departing from the scope or spirit of the present invention. Other embodiments derived from the present invention description will be apparent to those skilled in the art. The present application description and examples are exemplary 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 in the embodiments of the present 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] Embodiment 1

[0049] like Figure 1 - Figure 7 As shown, this embodiment provides a method for sequential aeroelastic optimization of aircraft structure topology-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 an 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, the initial structural model is topologically optimized 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;

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

[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; in terms of calculation methods, this method uses 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 on each node of the aerodynamic grid is obtained by the interpolation method, and the static aeroelastic deformation is obtained by the static analysis module of Nastran, and 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 achieved through Nastran sol200. The minimum compliance of the structure is set as the goal. The load is the aerodynamic load output in the aeroelastic analysis. The more appropriate volume constraints are set in advance, and the maximum displacement constraints, minimum structural dimensions, initial density values, density change amplitude and other parameters are set. After obtaining the results of a round of topology optimization, the aerodynamic force is recalculated and used as a new load input to achieve iterative optimization. After the topology optimization is completed, the component size is optimized according to the smoothed structure. Nastran sol200 can directly define the objective function, the initial value and upper and lower limits of the design variables, the design constraints, the maximum number of iterations, etc., and the optimal size is obtained through multiple rounds of iterations.

[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 components are determined. Taking into account that hypersonic flight will have a certain aerodynamic heating effect, the initial temperature field of the structure is set in advance, and the aerodynamic pressure on the surface of the hypersonic aircraft components is calculated according to 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 the conversion of 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 within the time dt, the speed of the piston is v n Change to v n +dv n , the distance of disturbance propagation is a·dt, then the mass of the disturbed gas is paSdt (ρ is air density, a is speed of sound, S is 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 in 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 sound speed formula, we can get:

[0064]

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

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

[0067]

[0068] In the formula, C is the integration constant, and according to the far-field boundary condition, p = p ∞ , V ∞ =0, then the integral 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, it is called the second-order, third-order and higher-order piston theory. The expression of the pressure coefficient of the third-order piston theory is 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 of 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 displacement of the node will affect the geometric shape of the component surface, the change of the geometric shape will affect the aerodynamic force calculation result, so the aerodynamic force and displacement are recalculated. After the iteration converges, the aeroelastic analysis module is terminated 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, 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, according to the virtual work principle or 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 unit 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 unit stiffness matrix of each unit:

[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 It is the n elements that constitute the unit shape function. D is the stress-strain relationship matrix in the constitutive equation σ=Dε:

[0100]

[0101] Among them, 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 change 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. After topological optimization, the new structure is obtained. The thickness of the beam and rib section 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 best 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 diagram before and after optimization is shown in 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 unit density, which is used to avoid the singularity of the overall stiffness matrix when the unit 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] Among them, F is the objective function of the optimization model, that is, the lightest structure 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 value range of the structural design variables, respectively.

[0114] Based on the analysis idea of ​​the flowchart, the aerodynamic / structural coupling analysis and optimization framework of hypersonic aircraft components is mainly composed of aeroelastic analysis module (aerodynamic force analysis, structural static analysis), topology optimization module and size optimization module. In the specific analysis, the initial flight state and aerodynamic shape parameters must first be input into the Matlab program. Considering the aerodynamic heating effect, it is necessary to pre-set the appropriate temperature field, which can be achieved by writing the temperature of each node in the pch file. The introduction of the temperature field will have a certain impact on the material parameters, which is mainly reflected in the reduction of the Young's modulus of the material. After the input file is prepared, the inner loop iterative coupling analysis of the aeroelastic module is carried out. The aerodynamic force on the surface of the aircraft component is calculated based on the third-order piston theory, and the structural displacement is calculated by Nastran sol101. After one calculation is completed, the node displacement in the generated pch file is read, the aerodynamic mesh of the model is updated according to the node displacement, and the aerodynamic force and structural displacement are recalculated until the structural displacement converges. At this point, the aeroelastic analysis module is completed.

[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 screened in the Matlab program. The 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. According to engineering experience, the structure after the units are screened 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 structure after smoothing (such as Figure 4 As shown in the figure, the objective function is set to minimize the total mass of the structure, while satisfying the constraints of deformation, stress, strain, etc., and finally obtaining the optimal form of the structure. Select Nastran sol200 to define the design variables, objective functions and constraints for size optimization, and select the maximum number of iterations by setting DOPTPRMDESMAX. When defining the design variables, it is necessary to define the initial size, upper and lower limits, and maximum change value of the design variables. After the optimization is completed, search for the keyword SUMMARY OF DESIGN CYCLEHISTORY in the generated F06 file to read the objective function value and the final value of the design variable after each step of optimization.

[0117] Compared with the CFD calculation results, the pressure, deformation and stress calculation results of the hypersonic vehicle obtained by the engineering algorithm are basically consistent, which proves the correctness and effectiveness of the aerodynamic calculation.

[0118] In terms of structural topology optimization, this embodiment significantly reduces the 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 dimensions of the beams and rib thicknesses, thereby reducing the structural weight and saving manufacturing costs while meeting the flight performance requirements.

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

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

Claims

1. A method for sequential aeroelastic optimization of aircraft structure topology and size, characterized in that: 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; Step S3: After the aeroelastic analysis is completed, the initial structural model is topologically optimized 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 the iteration is stopped, and the preliminary optimized structural data of the aircraft is output; 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 S2 specifically includes: Step S21: Calculating the 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 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 updating until the preset iterative condition is reached, then stop the iteration and output the current aerodynamic data.

3. The method for sequential aeroelastic optimization of aircraft structure topology and size according to claim 2, characterized in that: The step S22 specifically includes: The surface aerodynamic 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 data.

4. 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 compliance of the structure as the topology optimization objective function; Step S32: performing topology optimization on the initial structure 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 structure 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.

5. 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.

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

7. 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 best optimized structure of the aircraft is output.

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

Images