A method for combined design of thermal transport and load bearing structure of a hypersonic vehicle wing

Abstract

The application provides a combined design method for a heat transport and load-carrying structure of a hypersonic vehicle wing, and the combined design method is characterized in that: based on the topological optimization design of the load-carrying structure of the hypersonic vehicle under the condition of heat-force coupling and the topological design of the heat transport path of the wing based on a high-thermal-conductivity material, different reference schemes are provided on the load-carrying structure and the heat transport structure respectively, the two schemes are combined, and a combined design scheme of the load-carrying / heat transport structure is obtained. Compared with the single load-carrying structure and the single heat transport structure, the combined design scheme has higher implementation possibility, higher load-carrying / heat transport capacity and lower mass cost, and provides a more practically meaningful reference for the structural design of the hypersonic vehicle.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aircraft thermal protection system design, in particular to a method for combined design of thermal transport and load-carrying structure of a hypersonic aircraft wing. BACKGROUND

[0002] For a hypersonic aerospace vehicle flying at a certain angle of attack, when it is in a high Mach number return phase, the windward surface of the vehicle will bear a more severe aerodynamic heating, and the traditional thermal protection system uses heat protection / separation measures, which on the one hand further increases the structural mass of the vehicle, and on the other hand more and more difficult to meet the requirements of reusability. Taking the wing windward surface as an example, under the condition of 6Ma, 4° angle of attack flight, the calculated heat flow data shows that the maximum heat flux density of the leading edge of the windward surface is 22kW / m 2 , the maximum heat flux density of the leading edge of the vertical tail windward surface is 20kW / m 2 , which is smaller than 90kW / m 2 of the wing leading edge, if the traditional thermal protection system design is used, it will cause a large amount of heat protection / separation material, which is not conducive to the realization of the lightweight design goal of the structure, and the wing trailing edge and wing root are low-temperature regions, if the heat transport technology represented by high-thermal-conductivity materials or high-temperature heat pipes is used to transport heat from high-temperature regions to low-temperature regions, it will greatly alleviate the thermal load of the high-temperature regions, and at the same time, the joint design of high-thermal-conductivity solid medium or high-temperature heat pipes and load-carrying structure can be realized to a certain extent. SUMMARY

[0003] Technical problems to be solved

[0004] At present, when a hypersonic aerospace vehicle is in a high Mach number return phase, the windward surface of the vehicle bears a more severe aerodynamic heating, the traditional thermal protection system causes a large amount of heat protection / separation material, which is not conducive to the realization of the lightweight design goal of the structure, and the thermal protection efficiency is low, which cannot meet the actual working needs of the hypersonic vehicle in a large airspace and a wide speed range.

[0005] Technical scheme

[0006] In view of the above technical problems, the present application provides a method for combined design of thermal transport and load-carrying structure of a hypersonic aircraft wing, which is based on the lightweight design of load-carrying structure under the action of aerodynamic force and the topological optimization results of thermal transport path, and can achieve better load-carrying / thermal transport capacity and lower mass cost.

[0007] The technical scheme of the present application is as follows:

[0008] The method for combined design of thermal transport and load-carrying structure of a hypersonic aircraft wing comprises the following steps:

[0009] Step 1: Topology optimization design of wing load-bearing structure based on aerodynamic force:

[0010] For the wing load-bearing structure, first, the aerodynamic force distribution of the wing surface under different attack angles and Mach numbers is calculated using the CFD method, and the optimization design space is determined according to the wing geometry. Then, topology optimization is performed under extreme aerodynamic load conditions, where the volume fraction is used as the design variable, combined with displacement constraints and stiffness constraints, to create a topology optimization model with continuous changes in element relative density. The topology optimization force transmission path under aerodynamic load conditions is obtained after optimization.

[0011] Step 2: Topology design of wing heat transport path based on high thermal conductivity materials:

[0012] For the wing load-bearing structure, first, the heat flux density of the wing surface under severe conditions is obtained through wing aerodynamic heat calculation. Then, a topology optimization of the heat transport path is performed based on the heat transport path topology optimization model, with multiple cold sources and different material volume fractions as constraints, to obtain a better heat transport path. In the heat transport path topology optimization model, a finite element model of the wing is constructed based on the wing geometry, the properties of the high thermal conductivity materials used, and the multi-physical fields determined by the wing surface heat flux density, end face cold source, and temperature boundary. Then, a topology design of the three-dimensional wing heat transport path under non-uniform surface heat flux density conditions is carried out using the topology optimization design scheme based on the variable density method. The clarity of the heat transport path is considered during the optimization process, and the number and position of the cold ends are adjusted based on the optimization results to obtain the optimal wing heat transport path distribution.

[0013] Step 3: Joint design of wing load-bearing / heat transport structure:

[0014] First, similarity analysis is performed on the topology optimization force transmission path obtained in Step 1 and the wing heat transport path obtained in Step 2. If the similarity requirement is not met, return to Step 2 to adjust the thermal conductivity coefficient of the material. If the similarity requirement is met, use the graph skeleton extraction method in image processing to extract the skeleton of the path results, read the force transmission and heat transport paths, and design a load-bearing / heat transport structure with branched heat transport paths to complete the reconstruction and post-processing of the three-dimensional physical model of the wing.

[0015] Based on the load-bearing / heat transport structure with branched heat transport paths, the minimum overall structural mass of the wing is used as the objective function, the wing load-bearing beam cross-sectional size parameters and the skin and rib thickness are used as design variables, and strength and stiffness constraints are set to construct a size optimization model. Through optimization, the final lightweight design scheme of the wing load-bearing / heat transport structure is obtained.

[0016] Further, in step 1, the minimum flexibility of the wing bearing structure is taken as the optimization objective, the maximum displacement of the design domain node and the material allowable stress are taken as the optimization constraints, the design space volume fraction constraint is applied, and the optimization mathematical model is constructed as:

[0017]

[0018] s.t.K T T=P

[0019] KU=F=F m +F th

[0020]

[0021]

[0022] σ e ≤[σ]

[0023] 0<x min ≤x e ≤x max ≤1,e=1,2,…,N

[0024] In the above formula, C is the flexibility of the structure, K is the global stiffness matrix, and U is the global displacement matrix of the structure; K T is the heat conduction matrix, T is the temperature column, P is the heat load column, V(x) is the volume of the retained solid material of the wing after optimization, f is the volume fraction of the design domain, and V0 is the volume of the solid material unit filled with solid material; u i is the displacement of the structure due to the change of material distribution in the optimization process, u i * is the maximum displacement of the design domain; σ is the global stress generated in the optimization process of the design domain, and [σ] is the allowable stress of the material used by the structure.

[0025] Further, in step 1, a second-order algorithm method with global convergence is used to solve the topology optimization problem.

[0026] Further, in step 2, the geometric mean temperature of the wing is taken as the optimization objective, and the optimization model is constructed as:

[0027] find X=(x1,x2,…,x Ne )

[0028]

[0029]

[0030] P=kT

[0031] ∫ ΩkdΩ = const

[0032] T max ≤T lim

[0033] x min ≤x e ≤x max

[0034] In the above formula, V is the volume of the heat conduction domain obtained after optimization; n is the total number of units; v e is the volume of the e-th unit; q is the volume constraint factor; V0 is the initial volume of the design domain; x max , x min are upper and lower limits of the design variables.

[0035] Further, in step 3, the size optimization model constructed is:

[0036]

[0037]

[0038] Wherein, p is the material density, A0 is the surface area of the skin, T0 is the thickness of the skin, L j is the length of the load-bearing beam component, A k is the surface area of the k-th wing rib, H k is the thickness of the k-th wing rib, m0 is the total mass of the leading edge and trailing edge regions; g σ and g u are the strength and stiffness constraint conditions of the structure respectively, [sigma] and [u] are the allowable stress of the material and the maximum displacement generated by the allowable deformation of the structure respectively.

[0039] Advantages

[0040] The topological optimization design of the high-speed aircraft load-bearing structure under the condition of thermal-mechanical coupling and the topological design of the wing heat transport path based on high-thermal-conductivity materials provide different reference schemes on the load-bearing and heat transport structures respectively, and the present application further jointly designs the two schemes to obtain a load-bearing / heat transport structure joint design scheme. The scheme has higher implementability, load-bearing / heat transport capacity and lower mass cost than the single load-bearing structure and heat transport structure. The present application provides a joint method for realizing multiple targets under the condition of thermal-mechanical coupling, and provides a more practically meaningful reference for the structural design of the high-speed aircraft.

[0041] Additional aspects and advantages of the present application will be partially given in the following description, partially will become obvious from the following description, or will be understood by the practice of the present application. BRIEF DESCRIPTION OF DRAWINGS

[0042] The above and / or additional aspects and advantages of the present application will become apparent and more readily appreciated from the following description, taken in conjunction with the following drawings of which:

[0043] Figure 1 : wing load carrying / heat transport structure combined design procedure;

[0044] Figure 2 : wing load carrying structure topology optimization design procedure based on aerodynamic force;

[0045] Figure 3 : wing heat transport path topology optimization design procedure;

[0046] Figure 4 : wing airfoil region topology optimization results comparison;

[0047] Figure 5 : three-dimensional load carrying / heat transport structure model with branched heat transport paths;

[0048] Figure 6 : wing load carrying / heat transport structure combined design scheme (airfoil region heat transport material branches). DETAILED DESCRIPTION

[0049] The present application is directed to the actual working needs of high supersonic aircraft in large airspace and wide speed domain, based on the wing load carrying structure scheme and heat transport path topology optimization results under aerodynamic force, a tree branch-like heat transport branch structure model is constructed, heat transport characteristic analysis and mechanical characteristic analysis are carried out, and on this basis, wing load carrying / heat transport structure combined design is carried out, and a load carrying / heat transport structure combined design method is constructed.

[0050] As shown in Figure 1 , the wing load carrying / heat transport combined design procedure of the present application is divided into three steps.

[0051] The first step is the topology optimization design of the wing load carrying structure based on aerodynamic force, as shown in Figure 2 .

[0052] For the wing load carrying structure, first, the CFD method is used to calculate the aerodynamic force distribution of the wing surface under different attack angles and Mach numbers, and the optimization design space is determined according to the wing geometric configuration. Then, topology optimization is carried out under the condition of extreme aerodynamic load, wherein the volume fraction is taken as the design variable, combined with displacement constraint and stiffness constraint, a topology optimization model with continuous change of relative density of elements is constructed, and after optimization solution, the topology optimization transmission path under the condition of considering only aerodynamic load is obtained;

[0053] Specifically for wing load-bearing structure topology optimization design problem, the minimum flexibility of the wing load-bearing structure is taken as the optimization objective, and the material volume fraction of the wing load-bearing structure is taken as the constraint condition, and a topology optimization model with continuous change of element relative density is created:

[0054]

[0055]

[0056] KU=F

[0057] 0<x min ≤x e ≤x max ≤1,e=1,2,…,N

[0058] In the above formula, C is the flexibility of the wing load-bearing structure, and the physical meaning is the force caused by unit displacement; F is the load applied on the wing load-bearing structure, K is the global stiffness matrix of the wing load-bearing structure, U is the global displacement matrix of the wing load-bearing structure, k e is the element stiffness matrix, u e is the element displacement; N is the number of elements; V is the volume of the material after optimization, f is the volume fraction of the design domain, V0 is the volume of the solid model. x e is the element relative density, x min is the minimum relative density, x max is the maximum relative density, in order to avoid the singularity of the element stiffness matrix in the optimization process, the minimum relative density should be as small as possible.

[0059] The topology optimization process needs to consider the strength and stiffness changes of the structure, so further define the topology optimization problem suitable for the wing of a hypersonic aircraft under aerodynamic force, take the minimum structural flexibility as the objective function, consider the changes of structural stiffness and strength, i.e. take the maximum displacement of the design domain node and the allowable stress of the material as the optimization constraint, and apply the volume fraction constraint of the design space to construct the following optimization mathematical model:

[0060]

[0061] s.t.K T T=P

[0062] KU=F=F m +F th

[0063]

[0064]

[0065] σ e ≤[σ]

[0066] 0 < x min ≤ x e ≤ x max ≤ 1, e = 1, 2, …, N

[0067] In the above formula, C is the flexibility of the structure, K is the global stiffness matrix, and U is the global displacement matrix of the structure; K T is the heat conduction matrix, T is the temperature array, P is the thermal load array; V(x) is the volume of the retained solid material of the wing after optimization, f is the volume fraction of the design domain, and V0 is the volume filled with solid material elements. i is the displacement of the structure due to the change of material distribution during optimization, u i * is the maximum displacement of the design domain, which is represented as the maximum node displacement constraint of the wing surface and the vertical tail wing tip in this embodiment; σ is the global stress generated in the design domain during optimization, and [σ] is the allowable stress of the material used by the structure. To improve the convergence speed, the initial volume of the structure can be set as the upper limit of the volume fraction constraint, and in this embodiment, the volume fraction constraint upper limit is set to 30% in the process of topology optimization of the load-bearing structure.

[0068] Based on the above topology optimization model, the globally convergent second-order algorithm—GCMMA (Globally Convergent Method of Moving Asymptotes) method is used to solve the topology optimization problem in this embodiment, and the topology optimization result of the wing load-bearing structure is obtained.

[0069] The second step is the topology design of the wing heat transport path based on high thermal conductivity materials, as shown in Figure 3 .

[0070] For the wing load-bearing structure, the wing surface heat flux density under severe working conditions is first calculated through wing aerodynamic heat calculation, and then the wing geometric mean temperature is taken as the optimization target. Based on the heat transport path topology optimization model, topology optimization is carried out under the constraints of multiple cold sources and different material volume fractions, and a better heat transport path is obtained. In the heat transport path topology optimization model, according to the wing geometric configuration, the properties of the high thermal conductivity materials used, and the multiple physical fields determined according to the wing surface heat flux density, end face cold source and temperature boundary, a wing finite element model is constructed. The topology design of the three-dimensional wing heat transport path under the condition of non-uniform surface heat flux density is carried out by using the topology optimization design scheme based on the variable density method. The clarity of the heat transport path is considered in the optimization process, and the number and position of the cold ends are adjusted combined with the optimization result each time, so as to obtain the optimal distribution of the wing heat transport path.

[0071] In this embodiment, for the hypersonic vehicle wing, the design purpose of the heat transport system is to reduce the temperature of the wing severe aerodynamic heating part, that is, to reduce the geometric mean temperature of the wing as the heat conduction domain, while restricting the maximum temperature of the wing, ensuring that the wing components are less affected by the transport heat, obtaining the optimal heat conduction path of the wing, and obtaining an efficient and reliable heat transport system design scheme with a certain high thermal conductivity material mass cost. Based on the above analysis, the geometric mean temperature of the wing is taken as the optimization target in this embodiment, and the following mathematical model is established:

[0072] find X=(x1,x2,…,x Ne )

[0073]

[0074]

[0075] P=kT

[0076] ∫ Ω kdΩ=const

[0077] T max ≤T lim

[0078] x min ≤x e ≤x max

[0079] In the above formula, V is the volume of the heat conduction domain obtained after optimization; n is the total number of units; v e is the volume of the e-th unit; q is the volume constraint factor; V0 is the initial volume of the design domain; x max , x min are the upper and lower limits of the design variables. The design domain is discretized into a finite element grid, and in order to simplify the physical processing, it is assumed that the material distribution in each unit is uniform and has the same thermal conductivity.

[0080] Based on the above model, the engineering tool is used to obtain the heat transport path topology optimization result of the wing using high thermal conductivity material and ultra-high thermal conductivity coefficient under the condition of non-uniform heat flux density. The topology result shows that the transport heat flows from the hot end to the cold end, and the heat transport path presents a leaf vein-like shape.

[0081] The third step is the combined design of the wing bearing / heat transport structure.

[0082] The combined design method of bearing / heat transport structure is to reasonably distribute the bearing material which plays the bearing characteristics and the high thermal conductivity material which plays the heat transport function in the structure design to realize the combined design of the wing bearing / heat transport structure of the hypersonic aerospace vehicle, and meet the lightweight design requirement of the structure.

[0083] First, based on the topology optimization results of the load-bearing structure under aerodynamic-thermal-mechanical coupling obtained in the first step and the topology results of the wing's heat transport path obtained in the second step, a similarity analysis is conducted on the topology optimization results of the wing's force transmission path and heat transport path, such as... Figure 4 As shown, the force transmission and heat conduction paths formed by the two in the strake wing region are quite similar. The difference lies in the fact that the heat conduction path distribution shows more branch structures, ensuring an increase in heat dissipation area. From the overall distribution, the force transmission and heat conduction path distributions of the wing are highly similar, showing a high degree of overlap in the first-level branches. However, the second-level branch heat conduction paths are clearer and denser than the force transmission paths, which can provide a good reference for the joint design of the wing's load-bearing / heat transport structure.

[0084] Based on this, the image skeleton extraction method in image processing was used to extract the skeleton of the topology results for the two types of topology optimization structures, read the main force and heat transfer paths, and designed a load-bearing / heat transport structure with branched heat conduction paths to complete the reconstruction and post-processing of the wing's three-dimensional physical model.

[0085] The primary branch layout scheme for load-bearing / heat transport involves filling the load-bearing structure with a high thermal conductivity solid medium, thereby achieving a coordinated design between load-bearing and heat transport and obtaining a matching thickness and width for the heat transport channel.

[0086] To address the heat transport requirements, based on the heat transport topology optimization results from the second step, a secondary branch layout scheme with branched heat conduction paths was further designed under the primary branch structure. The heat transport at the connection points of the primary branch structure is more concentrated. The supporting structure uses TB-3 alloy with better thermal conductivity, while the remaining supporting structures use TC-4 alloy. The secondary branches use a single high thermal conductivity material to achieve higher heat transport efficiency. The resulting three-dimensional supporting / heat transport structure model with branched heat conduction paths is shown below. Figure 5 As shown in the diagram, this structural model uses a lightweight metal load-bearing structure designed considering aerodynamic effects as the load-bearing framework. The wing rib structure is eliminated. Each main load-bearing I-beam and secondary load-bearing I-beam is filled with ultra-high thermal conductivity material (first-level branches), and multiple heat transport channels (second-level branches) are formed by this ultra-high thermal conductivity material. Based on the wing's heat transfer topology optimization results, the heat transport channels composed of high thermal conductivity material are designed with numerous branched channel structures, which can further increase the contact area with the upper and lower wing panels, thus increasing the heat transfer area of ​​a large wing region. Simultaneously, more tree-like branched heat transport channels are designed at the wing leading edge to achieve rapid heat transfer from the high-temperature zone to the low-temperature zone, thereby reducing the temperature of the high-temperature zone.

[0087] The above method obtains the load / heat transport structure model of the wing, but cannot accurately obtain the specific structure size parameters, therefore, based on the three-dimensional load / heat transport structure model with branched heat conduction paths, the whole wing is subjected to load / heat transport structure joint design, through size optimization of the whole load / heat transport structure, detailed design of the load-carrying structure is completed and the purpose of lightweight design is achieved, and the load / heat transport joint design scheme of the whole wing is obtained, as shown in Figure 6

[0088] Generally, the objective or constraint functions used for size optimization include mass, volume, frequency, stress and strain, etc. Here, the structure size is defined to optimize the strength constraint and the stiffness constraint condition, the whole structure mass m w of the wing is taken as the objective function, the wing load beam section size parameters and the skin and rib thicknesses are taken as the design variables, a size optimization model is constructed, and then the size optimization method is further used to obtain the optimal design parameters of the wing load beam section under the action of aerodynamic force-heat, and the lightweight design scheme of the wing load / heat transport structure is obtained. The constructed size optimization model is:

[0089]

[0090]

[0091] Wherein, ρ is the material density, A0 is the surface area of the skin, T0 is the thickness of the skin, L j is the length of the load beam component, A k is the surface area of the kth rib, H k is the thickness of the kth rib, m0 is the total mass of the leading edge and trailing edge regions; g σ and g u are the strength and stiffness constraint conditions of the structure respectively, [σ] and [u] are the allowable stress of the material and the maximum displacement generated by the allowable deformation of the structure respectively, in this embodiment, [u] is set to 300mm.

[0092] Based on the above three steps, the wing load / heat transport joint design scheme which can be implemented in actual engineering is obtained, and the joint design method which meets the multi-objective conditions under the severe force-heat working conditions of the hypersonic vehicle in a wide speed range and a large airspace is realized.

[0093] Although the embodiments of the present application have been shown and described above, it should be understood that the above embodiments are exemplary and should not be construed as limiting the present application, and those skilled in the art can make changes, modifications, replacements and variations to the above embodiments without departing from the principles and purposes of the present application within the scope of the present application.​

Claims

1. A method for the combined design of heat transport and load-bearing structure of a hypersonic vehicle wing, characterized in that: Includes the following steps: Step 1: Topology optimization design of wing load-bearing structure based on aerodynamic forces: For the wing load-bearing structure, the aerodynamic force distribution on the wing surface under different angles of attack and Mach numbers is first calculated using the CFD method, and the optimization design space is determined according to the wing geometry. Then, topology optimization is performed under extreme aerodynamic load conditions, where volume fraction is used as the design variable. Combining displacement constraints and stiffness constraints, a topology optimization model with continuously changing relative density of the creation element is constructed. After optimization and solution, the topology-optimized force transmission path considering only aerodynamic load conditions is obtained. Step 2: Design of the airfoil heat transport path topology based on high thermal conductivity materials: For the wing load-bearing structure, the heat flux density on the wing surface under harsh conditions is first obtained through aerodynamic thermal calculations. Then, with the geometric mean temperature of the wing as the optimization target, topology optimization with multiple cold sources and different material volume fractions is performed based on the heat transport path topology optimization model to obtain a better heat transport path. In the heat transport path topology optimization model, a finite element model of the wing was constructed based on the wing geometry, the properties of the high thermal conductivity material used, and the multiphysics field determined by the wing surface heat flux density, end face cold source, and temperature boundary. Then, a topology optimization design scheme based on the variable density method was used to carry out the topology design of the three-dimensional wing heat transport path under the condition of non-uniform surface heat flux density. During the optimization process, the clarity of the heat transport path was considered, and the number and position of the cold ends were adjusted based on the results of each optimization to obtain the optimal wing heat transport path distribution. Step 3: Joint design of wing load-bearing / heat transport structure: First, based on the topology-optimized force transmission path obtained in step 1 and the wing heat transport path obtained in step 2, a similarity analysis is performed on the force transmission path and the heat transport path; if the similarity requirement is not met, return to step 2 to adjust the thermal conductivity of the material. If the similarity requirement is met, the graphic skeleton extraction method in image processing is used for the topology optimization force transmission path and the wing heat transport path. The skeleton of the path result is extracted, the force transmission and heat transfer paths are read, and the load-bearing / heat transport structure with branch heat conduction path is designed to complete the reconstruction and post-processing of the wing three-dimensional physical model. Based on the load-bearing / heat transport structure with branched heat conduction paths, the objective function is to minimize the overall structural mass of the wing. The design variables are the cross-sectional dimensions of the wing load-bearing beam and the thickness of the skin and ribs. Strength and stiffness constraints are set, and a size optimization model is constructed. Through optimization and solution, the final lightweight design scheme of the wing load-bearing / heat transport structure is obtained.

2. The method for jointly designing the heat transport and load-bearing structure of a hypersonic vehicle wing according to claim 1, characterized in that: In step 1, the optimization objective is to minimize the flexibility of the wing's load-bearing structure. The optimization constraints are the maximum displacement of the design domain nodes and the allowable material stress. A design space volume fraction constraint is applied, and the optimization mathematical model is constructed as follows: s.t.K T T=P KU=F=F m +F th s e ≤[σ] 0<x min ≤x e ≤x max ≤1,e=1,2,…,N,…, In the above formula, C is the structural flexibility, K is the global stiffness matrix, and U is the global displacement matrix of the structure; K T Here, T is the heat conduction matrix, P is the temperature array, and V(x) is the volume of solid material retained in the optimized wing, f is the volume fraction of the design domain, and V0 is the volume of the solid material element. i u represents the displacement of the structure caused by changes in material distribution during the optimization process. i * σ represents the maximum displacement occurring in the design domain; σ represents the global stress generated in the design domain during the optimization process; and [σ] represents the allowable stress of the material used in the structure.

3. The method for jointly designing the heat transport and load-bearing structure of a hypersonic vehicle wing according to claim 2, characterized in that: In step 1, a globally convergent second-order algorithm is used to solve the topology optimization problem.

4. The method for jointly designing the heat transport and load-bearing structure of a hypersonic vehicle wing according to claim 1, characterized in that: In step 2, the optimization model is constructed with the wing's geometric mean temperature as the optimization objective: find X=(x1,x2,…,x Ne ) P = kT ∫ Ω kdΩ=const T max ≤T lim x min ≤x e ≤x max In the above formula, V is the optimized thermal conductivity volume; n is the total number of elements; v e Let be the volume of the e-th element; q be the volume constraint factor; V0 be the initial volume of the design domain; x max x min Design upper and lower limits for variables.

5. The method for jointly designing the heat transport and load-bearing structure of a hypersonic vehicle wing according to claim 1, characterized in that: In step 3, the constructed size optimization model is as follows: Where ρ is the material density, A0 is the surface area of ​​the skin, T0 is the thickness of the skin, and L j Let A be the length of the load-bearing beam member. k Let H be the surface area of ​​the k-th rib. k Let m be the thickness of the k-th rib, m0 be the total mass of the leading and trailing edge regions; g σ and g u σ and u represent the strength and stiffness constraints of the structure, respectively, and [σ] and [u] represent the allowable stress of the material and the maximum displacement allowed by structural deformation, respectively.

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