A flat panel satellite structure considering heat dissipation and stiffness integration and a design method thereof

Abstract

The application provides a flat satellite structure considering heat dissipation and stiffness integration and a design method thereof, and the structure comprises a main extruded structure, a first side rib-shaped structure, a second side rib-shaped structure, a heat pipe and a force bearing seat; the main extruded structure is provided with an extruded profile with a rib-shaped section; a plurality of main extruded structures are spliced to form the flat satellite structure; the geometric shape of the rib-shaped section of the main extruded structure is determined according to a multi-target topology optimization method; the first side rib-shaped structure and the second side rib-shaped structure are respectively located at the first side and the second side of the main extruded structure; the heat pipe is prepared through an extrusion forming process and is fixedly connected with the main extruded structure; the heat pipe is used for uniform heat dissipation; and the force bearing seat is fixedly connected at the end of the flat satellite structure. The application uses an aluminum extruded profile as the main structure of the flat satellite, balances the stiffness and heat dissipation requirements of the structure through multi-target topology optimization, and realizes the integration of high stiffness, light weight and heat dissipation by welding the aluminum extruded profile with the heat pipe.

Description

Technical Field

[0001] This application relates to the technical field of equipment used in spaceflight, and in particular to a flat-panel satellite structure that integrates heat dissipation and rigidity, and its design method. Background Technology

[0002] In recent years, with the advancement of aerospace technology and the increasing demand for space satellite networks, countries have been vying to propose and deploy satellite constellation plans, requiring the launch of a large number of satellites to acquire resources in low Earth orbit. To maximize the use of the limited carrying capacity and space of launch vehicles and reduce satellite launch costs, multiple satellites are typically carried in a single launch, a "one rocket, multiple satellites" launch format. Taking stacked flat-panel satellites as an example, this structure is compact, small in size, and effectively utilizes the limited space within the launch vehicle fairing, making it the most suitable main structure for "one rocket, multiple satellites" launches. Traditional satellite flat-panel designs typically employ a skin-honeycomb sandwich composite structure, mainly composed of an upper skin, a lower skin, and a honeycomb core. To meet the heat dissipation and temperature control requirements of high-heat-flux components, heat pipes are usually embedded inside the satellite flat-panel. Furthermore, to meet other functional requirements, various embedded components are installed inside the satellite flat-panel using adhesive or bolt connections. Therefore, traditional satellite structures suffer from a large number of components, complex processes, and long manufacturing cycles. To meet the demand for launching a large number of satellites in a short time and to achieve mass production of satellite structures, [further solutions are needed].

[0003] The related technology (patent publication number: CN110856415A) invented an integrated satellite structural panel and its manufacturing method. It adopts a "double-layer skin + Kagome lattice structure" to replace the traditional skin honeycomb sandwich composite structure. In addition, the structure is manufactured by 3D printing. This satellite design and manufacturing method can reduce the use of structural adhesive and realize the integrated design and manufacturing of heat pipe and structural panel. To a certain extent, it simplifies the number of parts and production process of satellite flat panel. However, the 3D printing manufacturing process is not suitable for mass production.

[0004] The related technology (patent publication number: CN117429627A) invented a thermo-mechanical integrated satellite structure plate and satellite suitable for mass production. It uses thin metal sheets to replace the traditional skin honeycomb sandwich composite structure. It can be manufactured using die casting, sheet metal, stamping and other methods suitable for mass production. It also uses grooves to install heat pipes and improve the local rigidity of the structure. This method effectively simplifies the production process and improves production efficiency, reduces the number of structural parts, and takes into account the requirements of structural rigidity, strength and heat dissipation. However, compared with the traditional skin honeycomb sandwich composite structure, the simple groove design may not be able to meet the rigidity and strength requirements of the satellite structure under all working conditions.

[0005] The related technology (patent publication number: CN116329487A) uses an integral aluminum alloy casting process to manufacture flat satellite structures. This technology has high production efficiency and good process adaptability for complex structures. However, for non-mass manufacturing tasks, the mold development time is long and the preparation cost is too high.

[0006] In summary, the design and manufacture of flat-panel satellites, taking into account factors such as lightweight structure, thermodynamic performance, and manufacturing efficiency, still present significant challenges. Summary of the Invention

[0007] The purpose of this application is to provide a flat-panel satellite structure and its design method that integrates heat dissipation and stiffness. The structure uses aluminum extrusion profiles as the main structure of the flat-panel satellite and balances the stiffness and heat dissipation requirements of the structure through multi-objective topology optimization. Heat pipes are manufactured by extrusion molding and the aluminum extrusion profiles are welded to the heat pipes to achieve a high-stiffness, lightweight and heat dissipation integrated design for the flat-panel satellite structure.

[0008] To achieve the above objectives, as a first aspect of this application, this application provides a flat-panel satellite structure that integrates heat dissipation and rigidity. The structure includes: multiple main extruded structures, a first side rib structure, a second side rib structure, heat pipes, and a support base. The main extruded structures are extruded profiles with ribbed cross-sections. Multiple main extruded structures are spliced ​​together to form the flat-panel satellite structure. The geometry of the ribbed cross-sections of the main extruded structures is determined using a multi-objective topology optimization method. The first side rib structure and the second side rib structure are located on the first and second sides of the main extruded structures, respectively. The heat pipe is prepared by an extrusion molding process and is fixedly connected to the main extruded structures. The heat pipe is used for uniform heat dissipation. The support base is fixedly connected to the end of the flat-panel satellite structure.

[0009] As described above, in the integrated flat satellite structure considering both heat dissipation and rigidity, the first side rib structure is connected to the main extrusion structure via an integral extrusion method; the second side rib structure is formed by milling on the side of the main extrusion structure away from the first side rib structure.

[0010] As described above, the flat satellite structure integrates heat dissipation and rigidity, wherein the main extrusion structure includes an outer surface flat plate structure and an internal rib structure, and the internal rib structure is disposed inside the outer surface flat plate structure.

[0011] As described above, the flat-panel satellite structure considers both heat dissipation and stiffness, wherein the multi-objective topology optimization method takes the minimum of the total flexibility coupled with thermal flexibility and flexibility as the objective function.

[0012] As described above, in the planar satellite structure that integrates heat dissipation and rigidity, the cross-sectional shape of the first side rib structure is L-shaped, T-shaped, or other shapes.

[0013] As described above, the flat-plate satellite structure integrates heat dissipation and rigidity, wherein the second side rib structure is an orthogonal grid structure or other structures.

[0014] As described above, in the planar satellite structure that integrates heat dissipation and rigidity, the heat pipe is made of extruded aluminum and connected to the main extruded structure by friction stir welding.

[0015] As described above, the flat-panel satellite structure integrates heat dissipation and rigidity, wherein multiple main extrusion structures of the same length or multiple main extrusion structures of different lengths are welded together to form the flat-panel satellite structure.

[0016] As a second aspect of this application, this application provides a design method for a flat-panel satellite structure that integrates heat dissipation and stiffness, for obtaining the aforementioned flat-panel satellite structure that integrates heat dissipation and stiffness; the method includes:

[0017] Step S1: Perform finite element modeling on the flat-panel satellite structure to determine the design domain, loads, and boundary conditions;

[0018] Step S2: Construct a material interpolation model;

[0019] Step S3: Based on the BESO algorithm, establish a multi-objective topology optimization problem with the objective function of minimizing the total compliance coupled with structural compliance and thermal compliance.

[0020] Step S4: Perform finite element analysis on the heat transfer of the flat-panel satellite structure to obtain nodal temperature data;

[0021] Step S5: Perform thermo-mechanical coupling analysis on the flat-panel satellite structure to obtain node displacement data;

[0022] Step S6: Calculate the sensitivity of the objective function based on the node temperature data and node displacement data;

[0023] Step S7: Based on the sensitivity of the objective function, the design variables are updated and iterated using a gradient-based optimization algorithm;

[0024] Step S8: Determine whether the multi-objective topology optimization problem has converged. If it has not converged, return to step S2; if it has converged, the calculation is complete.

[0025] The above-described design method for a flat-panel satellite structure that considers both heat dissipation and stiffness integration involves establishing a multi-objective topology optimization problem, which is a thermo-mechanical coupled multi-objective topology optimization problem:

[0026] Find x = (x1, x2, ..., x Ne );

[0027]

[0028] Where Ne represents the total number of finite element elements; c represents the total compliance; x i Represents the topological variable (cell density) of the i-th cell; ω1 and ω2 both represent weighting coefficients; ω1 + ω2 = 1; c th For thermal flexibility, This indicates that the minimum thermal flexibility value is taken; This indicates taking the maximum thermal flexibility value; c m For softness; This indicates taking the minimum flexibility value; This indicates taking the maximum flexibility; st indicates satisfying the constraint conditions; V * For volume constraints, N represents the number of finite element elements; V i Q is the initial volume of the element; Q is the heat flux input column matrix; Q T and F T The T in the upper right corner indicates matrix transpose; T is the nodal temperature column matrix, U is the nodal displacement column matrix, and F is the nodal force load column matrix.

[0029] The beneficial effects achieved by this application are as follows:

[0030] (1) Compared with the traditional skin honeycomb sandwich composite structure, the flat satellite structure adopts the extrusion molding process, which can be extruded in one piece, reducing the use of structural adhesive and the weight of additional structural adhesive, which is conducive to the lightweighting of the structure. At the same time, the cross-sectional shape of the extruded profile can be optimized by topology design based on the stiffness requirements and heat dissipation requirements of the equipment to achieve good thermo-mechanical performance of the flat satellite.

[0031] (2) The first and second side rib structures of this application are designed flexibly, which can meet the functional requirements of different satellites and realize the lightweight load-bearing design of flat satellite structure, so as to better solve the contradiction between the lightweight design of flat satellite structure and insufficient mechanical load-bearing performance.

[0032] (3) The extrusion process of this application can realize the mechanical and thermal integration of the flat satellite structure. Compared with the traditional bonding method of upper and lower aluminum skin + aluminum honeycomb core, the process of this invention is simpler and more designable. At the same time, the aluminum extruded profile has excellent mechanical properties, and the extrusion molding technology is mature with low mold cost. Compared with the manufacturing method of 3D printing or the one-piece die casting process, the design and manufacturing cost of the flat satellite structure is reduced. Attached Figure Description

[0033] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments recorded in this application. For those skilled in the art, other drawings can be obtained based on these drawings.

[0034] Figure 1 This is a schematic diagram of an aluminum extruded profile flat satellite structure according to an embodiment of this application.

[0035] Figure 2 This is a schematic diagram of the aluminum extrusion profile structure according to an embodiment of this application.

[0036] Figure 3 These are schematic diagrams of two different types of flat-panel satellite main structures according to embodiments of this application.

[0037] Figure 4 This is a schematic diagram illustrating the connection between the heat pipe and the extruded profile, as well as the placement of the equipment, according to an embodiment of this application.

[0038] Figure 5 This is a schematic diagram of different rib structures milled according to different partitions in an embodiment of this application.

[0039] Figure 6 This is a flowchart illustrating a design method for a flat-panel satellite structure that integrates heat dissipation and rigidity, according to an embodiment of this application.

[0040] Reference numerals: 1-Main extrusion structure; 2-First side rib structure; 3-Second side rib structure; 4-Heat pipe; 5-Support seat; 6-Outer surface flat plate structure; 7-Internal rib structure; 8-First direction; 9-Second direction; 10-High heat consumption equipment; 11-Low heat consumption equipment. Detailed Implementation

[0041] The technical solutions of the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application.

[0042] Example 1

[0043] like Figure 1As shown, this application provides a flat-panel satellite structure that integrates heat dissipation and rigidity. The structure includes: multiple main extruded structures 1, a first side rib structure 2, a second side rib structure 3, a heat pipe 4, and a support base 5. The main extruded structure 1 is an extruded profile with a ribbed cross-section. The main extruded structure 1 uses aluminum extruded profiles as raw materials. The cross-sectional shape of the aluminum extruded profile is designed according to the load-bearing rigidity and heat dissipation requirements of the satellite structure to achieve good mechanical and thermal performance of the flat-panel satellite structure. Multiple main extruded structures 1 are spliced ​​together to form the flat-panel satellite structure. The geometry of the ribbed cross-section of the main extruded structure 1 is determined using a multi-objective topology optimization method. The first side rib structure 2 and the second side rib structure 3 are located on the first and second sides of the main extruded structure 1, respectively, and the first and second sides are opposite sides. The heat pipe 4 is fixedly connected to the main extruded structure 1 and is used for uniform heat dissipation. The support base 5 is fixedly connected to the end of the flat-panel satellite structure. The heat pipe 4 is made of extruded aluminum and connected to the main extruded structure 1 by friction stir welding to achieve high heat dissipation performance and structural reliability.

[0044] As a specific embodiment of the present invention, the main extrusion structure 1 is divided into regions according to functional differences based on the different requirements for equipment bearing, installation, and connection with other satellite stacks.

[0045] As a specific embodiment of the present invention, the first side rib structure 2 is connected to the main body extrusion structure 1 by an integral extrusion method; the second side rib structure 3 is formed by milling on the side of the main body extrusion structure 1 away from the first side rib structure 2.

[0046] like Figure 2 As shown, the main extrusion structure 1 includes an outer surface flat plate structure 6 and an inner rib structure 7, with the inner rib structure 7 located inside the outer surface flat plate structure 6.

[0047] As a specific embodiment of the present invention, the multi-objective topology optimization method is a thermo-mechanical coupled topology optimization method. The objective function is to minimize the total flexibility of the coupled thermal flexibility and the overall flexibility, in order to obtain an extruded profile cross-sectional shape with superior performance under certain operating conditions. A multi-objective coupled optimization model incorporating heat dissipation efficiency and structural stiffness is established, with the objective functions of minimizing the thermal resistance of the heat dissipation path and maximizing the structural stiffness. A topology optimization algorithm is then used for thermo-mechanical coupled finite element analysis. Specifically, thermo-mechanical coupled topology optimization is performed based on the heat dissipation and stiffness requirements of the flat-panel satellite structure. Thermal flexibility and flexibility are used to measure heat dissipation and stiffness respectively, and coupled into a total flexibility c. The optimization objective is to minimize the total flexibility c.

[0048] As a specific embodiment of the present invention, the first side rib structure 2 is designed according to the load distribution and size of the flat satellite structure, seeking to achieve lightweight and high specific stiffness of the flat satellite structure. Specifically, analysis is conducted through theoretical analysis or finite element numerical calculation to obtain a first side rib structure 2 with superior performance, so that the extruded profile maintains equivalent or higher bending / torsional stiffness while reducing weight. Different first side rib structures 2 are formed by extrusion, and the cross-sectional shape of the first side rib structure 2 includes, but is not limited to, L-shape, T-shape, etc. Alternatively, the first side rib structure 2 can also be other shapes, and the cross-sectional shape of the first side rib structure 2 is not limited here.

[0049] As a specific embodiment of the present invention, the first side rib structures 2 are spaced apart along the second direction 9. It should be noted that the embodiment of the present invention demonstrates one form of the rib structure 2 of the first side rib structure 2, but is not limited to this form. The spacing between adjacent first side rib structures 2 along the second direction 9, as well as the number of rib structures, can be designed according to the actual load-bearing requirements of the satellite. Furthermore, the first side rib structures 2 can also improve the bending / torsional stiffness of the main extrusion structure 1.

[0050] As a specific embodiment of the present invention, multiple main extrusion structures 1 are connected by welding. Since the aluminum extrusion profiles are all long strip profiles, friction stir welding can be considered for connection, but it is not limited to this welding method.

[0051] As a specific embodiment of the present invention, multiple main body extrusion structures 1 of the same length or multiple main body extrusion structures 1 of different lengths are spliced ​​together to form a flat satellite structure.

[0052] like Figure 3 As shown, two different types of flat-panel satellite main body structures are illustrated. One type is a flat-panel satellite main body structure formed by welding together main body extrusion structures 1 of the same length; the other type is a flat-panel satellite main body structure formed by welding together main body extrusion structures 1 of different lengths. However, the design is not limited to these two types, and the type of satellite flat-panel structure can be determined based on factors such as the actual design requirements of the satellite, the space available for the launch vehicle, and the equipment installation requirements.

[0053] As a specific embodiment of the present invention, the second side rib structure 3 divides the satellite flat panel structure into regions according to the load-bearing distribution and size of the flat satellite structure. The second side rib structure 3 is formed by milling to create different rib structures, which can be conventional geometric rib structures or rib structures obtained through topology optimization, etc. The layout and dimensions of the first side rib structure 2 and the second side rib structure 3 can be designed and laid out according to load-bearing requirements or functional requirements.

[0054] like Figure 2As shown, the first side rib structure 2 extends along the first direction 8 or along the profile extrusion direction of the main extrusion structure 1. The second side rib structure 3 can extend in different directions through milling. The second side rib structure 3 can be arranged along the first direction 8 or multiple directions, and the arrangement density and geometric dimensions are set according to the differences in functional areas. The structure is divided into functional zones, including but not limited to the load-bearing zone, the thermal control zone, and the installation zone. The second side rib structure of each zone can be designed or optimized independently.

[0055] Preferably, the milled second side rib structure 3 can extend along the first direction 8, or along a direction different from the first direction 8, or it can have multiple directions. Preferably, the milled second side rib structure 3 can be of one type, or it can be divided into multiple types according to different functional requirements.

[0056] In a specific embodiment of the present invention, the heat pipe 4 is formed by aluminum extrusion and connected to the main extrusion structure 1 by friction stir welding. This integrated design aims to meet the high heat dissipation requirements of high-heat-consuming equipment while avoiding the need for adhesive bonding of the heat pipe 4. Furthermore, the position, quantity, and size of the heat pipe 4 are arranged according to the installation location and heat flow path of the high-heat-consuming equipment 10, adapting to different heat dissipation needs.

[0057] As a specific embodiment of the present invention, for the high heat-consuming device 10, heat pipe 4 is used for heat dissipation. In this embodiment, the heat pipe 4 is manufactured using extrusion molding. The main extrusion structure 1 and the heat pipe 4 can be connected by friction stir welding. Compared to traditional satellite structures, this reduces the process of embedding the heat pipe 4 into the honeycomb core with adhesive, and also reduces the use of structural adhesive, which simplifies the manufacturing process of the satellite structure and reduces its weight. Simultaneously, the integrated design and manufacturing method of this embodiment allows the flat-panel satellite structure to simultaneously meet the requirements of lightweight load-bearing and heat dissipation. Specifically, for the high heat-consuming device 10, heat pipe 4 is used for heat dissipation, while for the low heat-consuming device 11, the main extrusion structure 1 itself, obtained through a thermo-mechanical coupling topology optimization method, provides heat dissipation.

[0058] like Figure 4 The diagram illustrates a connection arrangement between the heat pipe 4 and the main extrusion structure 1. It should be noted that the position and number of the heat pipe 4 can be designed and arranged according to the actual installation or heat dissipation requirements of the equipment, and are not limited to the arrangement shown in this embodiment.

[0059] As a specific embodiment of the present invention, the manufacturing process of the flat satellite structure includes extrusion molding, milling molding, and welding connection. The processing methods at each stage are flexible, the technology is mature, and the process complexity is low. Specifically, the cross-sectional shape of the aluminum extrusion profile is designed based on the results of thermo-mechanical coupling topology optimization, allowing for single continuous extrusion molding and obtaining multiple extrusion profiles through cutting. Based on the geometric dimensions of the flat satellite structure, the multiple extrusion profiles are welded together, and then the main body structure of the flat satellite is manufactured through milling. Finally, the load-bearing seat 5 is connected to the main body structure of the flat satellite through welding.

[0060] As a specific embodiment of the present invention, considering lightweight design, a second side rib structure 3 exists on the other side of the main extrusion structure, relative to the first side rib structure 2. Specifically, the second side rib structure 3 is formed on the main extrusion structure 1 by milling after the extruded profiles are welded together. Based on the load-bearing requirements and equipment installation requirements of the flat-panel satellite, the main extrusion structure 1 of the flat-panel satellite structure is divided into regions, and different second side rib structures 3 are designed to meet different load-bearing or functional requirements. Optionally, the second side rib structure 3 can be an orthogonal grid shape; or the second side rib structure 3 can be a shape obtained by thermo-mechanical coupling topology optimization; or the second side rib structure 3 can also be other shapes. Optionally, the second side rib structure 3 can extend along one direction; or the second side rib structure 3 can extend along multiple directions. Optionally, the second side rib structure 3 can be uniformly distributed; or the second side rib structure 3 can be non-uniformly distributed. Optionally, the second side rib structure 3 can be one type or multiple types. According to the actual needs of the satellite, the main extrusion structure 1 of the flat satellite structure can be divided into corresponding sections. According to the load-bearing or functional requirements of different areas, various second side rib structures 3 can be processed on the main extrusion structure 1 by milling.

[0061] like Figure 5 As shown, a partitioning pattern is illustrated on the main extrusion structure 1, including multiple partitions of different forms. The shapes or sizes of the second rib structures 3 corresponding to the different partitioning patterns are different. However, this application is not limited to this type of partitioning. Figure 5 The partitioning format shown can also be other partitioning formats.

[0062] The specific steps of the flat-panel satellite structure manufacturing method according to this invention are as follows:

[0063] Step T1: Based on the target satellite's carrying or functional requirements, determine the basic dimensions and geometry of the main extrusion structure and the first side rib structure; in other words, determine the cross-sectional dimensions and shape required for the extrusion molding of the main extrusion structure. The main extrusion structure and the first side rib structure are obtained by extrusion molding along the first direction.

[0064] Step T2, based on step T1, determines the type of the flat-panel satellite's main structure according to the target satellite's carrying or functional requirements, and determines the structural dimensions and quantity to be welded. Multiple aluminum extrusion profiles (including the main extrusion structure 1 and heat pipe 4) are obtained through cutting. These multiple aluminum extrusion profiles are then welded together using friction stir welding to obtain the satellite's flat-panel main structure. The location of heat pipe 4 can be determined based on the installation location of the equipment with high heat consumption.

[0065] Step T3, based on step T2, determines the basic dimensions and geometry of the second side rib structure according to the target satellite's load-bearing or functional requirements. The second side rib structure can be obtained through milling. Specifically, the main structure is divided into regions according to the functional requirements of the flat-panel satellite's main structure. Typically, equipment of different sizes and weights needs to be installed on the flat-panel satellite structure. Based on different installation requirements, the main structure of the flat-panel satellite can be divided into regions, and the dimensions, shape, and quantity of the second side rib structure in each region can be determined. Specifically, larger and more concentrated second side rib structures can be designed for the main load-bearing areas of the satellite's flat-panel structure, while smaller and sparser second side rib structures can be designed for secondary load-bearing or non-load-bearing areas.

[0066] Step T4 involves welding the load-bearing seats to the ends of the flat-panel satellite's main structure. The number and distribution of these load-bearing seats are determined based on the stacking method or installation requirements of the flat-panel satellite structure.

[0067] Example 2

[0068] like Figure 6 As shown, this application provides a design method for a flat-panel satellite structure that integrates heat dissipation and stiffness, used to obtain a flat-panel satellite structure that integrates heat dissipation and stiffness; the method includes:

[0069] Step S1: Perform finite element modeling of the flat-panel satellite structure.

[0070] Specifically, finite element modeling is performed on the flat-panel satellite structure to obtain the model's mesh information, loads and boundary conditions, and design and non-design domains.

[0071] Mesh information is the "computational carrier" for finite element analysis. A high-quality mesh can discretize a continuous flat-plate satellite structure into computable elements, ensuring accurate calculation results for mechanical parameters such as stress, strain, and displacement. A reasonable mesh density can balance computational accuracy and efficiency, avoiding distortion of results due to an excessively coarse mesh size, or waste of computational resources due to an excessively fine mesh size.

[0072] Loads and boundary conditions are the "input conditions" for simulating the actual working state of a flat-panel satellite structure. Loads (such as vibration loads during launch, temperature loads in orbit, and space radiation pressure) determine the external forces / environmental effects that the structure must withstand, and are a prerequisite for analyzing the structural strength and stiffness. Boundary conditions (such as connection constraints between the satellite and the launch vehicle, and assembly constraints between components) define the fixed / free state of the structure, ensuring that the calculation model is consistent with the actual constraints and avoiding calculation results that do not match engineering reality.

[0073] Design domain and non-design domain: These are the "division criteria" for clarifying the scope of structural optimization and retention. The design domain is the area where structural topology optimization and size adjustment can be performed, with the aim of reducing weight and cost while meeting performance requirements; the non-design domain is the key area where the original structural form must be retained (such as interfaces, sensor installation locations, etc.) to ensure that satellite functions are not affected by optimization.

[0074] Step S2: Construct a material interpolation model.

[0075] The material interpolation model includes:

[0076] Using a power function interpolation scheme, the thermal conductivity coefficient of each unit is defined as:

[0077] k i =x p k0;

[0078] Where, k i The value represents the thermal conductivity of the i-th element; x represents the element density; p represents the penalty coefficient; k0 represents the reference thermal conductivity.

[0079] The elastic modulus of each element is defined as follows:

[0080] E i =x p E0;

[0081] Among them, E i E represents the elastic modulus of the i-th element; E0 represents the reference elastic modulus.

[0082] Step S3: Based on the BESO algorithm (Bidirectional Asymptotic Structure Optimization Algorithm), a multi-objective topology optimization problem is established with the objective function of minimizing the total compliance coupled with structural compliance and thermal compliance.

[0083] The multi-objective topology optimization problem is as follows:

[0084] Find x = (x1, x2, ..., x Ne );

[0085]

[0086] Where Ne represents the total number of finite element elements; c represents the total compliance; x i Represents the topological variable of the i-th unit; ω1 and ω2 both represent weight coefficients; ω1 + ω2 = 1; c th For thermal flexibility, This indicates that the minimum thermal flexibility value is taken; This indicates taking the maximum thermal flexibility value; c m For softness; This indicates taking the minimum flexibility value; This indicates taking the maximum flexibility; st indicates satisfying the constraint conditions; V * For volume constraints, N represents the number of finite element elements; V i Q is the initial volume of the element; Q is the heat flux input array; Q T and F T The T in the upper right corner indicates matrix transpose; T is the nodal temperature column matrix, U is the nodal displacement column matrix, and F is the nodal force load column matrix.

[0087] Step S4: Perform finite element analysis on the heat transfer of the flat-panel satellite structure to obtain nodal temperature data.

[0088] The node temperature is calculated using the following formula:

[0089] Q = K th T

[0090] Where Q is the heat flux input column matrix; K th is the heat transfer stiffness matrix; T is the nodal temperature column matrix.

[0091] Step S5: Perform a thermo-mechanical coupling analysis on the flat-panel satellite structure to obtain node displacement data.

[0092] The nodal displacement data is calculated using the following formula:

[0093]

[0094] F m +F th =KU

[0095] Among them, F th For the equivalent thermal stress load, F m The mechanical load is represented by E; the elastic modulus is represented by α; and the coefficient of thermal expansion is represented by α. Let μ be the temperature difference, μ be Poisson's ratio, and ε be... th B represents thermal strain; B represents the strain-displacement matrix; B T The upper right corner T represents matrix transpose; D represents elastic constitutive matrix; K is stiffness matrix; U is nodal displacement column matrix.

[0096] Thermal expansion caused by temperature difference acts on the nodes in the form of nodal forces, forming an overall vector load together with mechanical loads. Its function is to realize an equivalent simulation of the actual stress state of the structure under "thermal-mechanical" coupled conditions, and to provide a unified load input that conforms to the actual working conditions for subsequent strength, stiffness and deformation analysis.

[0097] Step S6: Calculate the sensitivity of the objective function based on the node temperature data and node displacement data.

[0098] Step S6 includes:

[0099] Step S610: Calculate the sensitivity of thermal flexibility.

[0100] The formula for calculating the sensitivity of thermal flexibility is as follows:

[0101] S th =px p-1 T T K th T;

[0102] Among them, S th The sensitivity of thermal flexibility is represented by p; x is the penalty coefficient; and K is the element density. th is the heat transfer stiffness matrix; T is the nodal temperature column matrix.

[0103] Step S620: Calculate the sensitivity of the compliance.

[0104] The formula for calculating the sensitivity of flexibility is as follows:

[0105]

[0106] Among them, S m The sensitivity represents the compliance; U is the nodal displacement column matrix; x is the element density; F th is the equivalent thermal stress load; K is the stiffness matrix.

[0107] Step S630: Calculate the sensitivity of the objective function based on the sensitivity of thermal flexibility and the sensitivity of flexibility.

[0108] The formula for calculating the sensitivity of the objective function is as follows:

[0109]

[0110] Where S represents the sensitivity of the objective function; S th Sensitivity to thermal flexibility; S m Sensitivity indicating flexibility; This represents the maximum thermal flexibility. This represents the maximum flexibility. This represents the minimum thermal flexibility. ω1 and ω2 represent the weighting coefficients of thermal flexibility sensitivity and flexibility sensitivity sensitivity, respectively; ω1 + ω2 = 1.

[0111] This invention calculates the sensitivity of the objective function based on nodal temperature and displacement data and performs normalized weighting. Normalized weighting eliminates dimensional bias and ensures weighting fairness. Sensitivity reflects the contribution of changes in design variables (element density) to the overall structural flexibility and is the core basis for iteratively updating design variables during optimization. Thermo-mechanical coupled topology optimization can be understood as a multi-load optimization problem with both mechanical and thermal loads. Normalized weighting combines the sensitivity results under different loads through weighting coefficients to obtain a unified update criterion, ensuring that the influence of different loads on sensitivity is considered, and reflecting the importance or degree of influence of each load through weights.

[0112] Step S7: Based on the sensitivity of the objective function, the design variables are updated and iterated using a gradient-based optimization algorithm.

[0113] Specifically, based on the sensitivity of the objective function, a gradient-based optimization algorithm is used to update the design variables iteratively along the "optimal change direction of the objective function" indicated by the sensitivity, gradually adjusting the design variables to ultimately minimize the objective function while satisfying the constraints.

[0114] Step S7 includes:

[0115] Step S710: Initialize the iteration parameters.

[0116] Specifically, set initial values: assign initial feasible values ​​to design variables (such as initial values ​​based on experience or preliminary models), and set iteration termination conditions.

[0117] Step S720: Calculate the objective function value and sensitivity (gradient) for the current iteration.

[0118] Based on the current design variables, a finite element analysis is completed, outputting data such as nodal temperature and displacement, and then the objective function value and sensitivity under the current state are calculated.

[0119] Step S730: Determine the search direction for design variables based on gradients.

[0120] The gradient direction is the direction in which the objective function changes the fastest. Optimization algorithms will adjust design variables along the negative gradient direction (if the objective is to minimize, such as reducing mass or stress).

[0121] Step S740: Determine the update step size for the design variables.

[0122] Step S750: Update the design variables and verify the constraints.

[0123] Specifically, calculate the new design variable values ​​based on "search direction + optimal step size". Verify whether the new variable values ​​meet the constraints.

[0124] Step S760: Determine whether the iteration has terminated, and then loop or output the result.

[0125] Step S8: Determine whether the multi-objective topology optimization problem has converged. If it has not converged, return to step S2; if it has converged, the calculation is complete.

[0126] Specifically, check whether the current iteration meets the preset termination condition. If the termination condition is met, output the final optimal design variable (unit density) and obtain the optimal structural thickness distribution, i.e., the cross-sectional shape of the extruded profile, based on the unit density; if not, return to step S2 and repeat the loop of "calculate objective function and sensitivity → determine direction → select step size → update variables" until convergence.

[0127] After the calculation is completed, the shape of the internal rib structure 7 of the main extrusion structure 1 is obtained. Figure 2 The embodiments illustrate a shape optimized by thermo-mechanical coupling, which can meet a certain functional requirement or load-bearing requirement under a certain working condition. It should be noted that the cross-sectional shape of the internal rib structure 7 can be calculated by topology optimization under specific load-bearing and heat dissipation conditions, and is therefore not limited to the form shown in the embodiments, but can also be a cross-sectional shape that meets other working conditions.

[0128] The beneficial effects achieved by this application are as follows:

[0129] (1) Compared with the traditional skin honeycomb sandwich composite structure, the flat satellite structure adopts the extrusion molding process, which can be extruded in one piece, reducing the use of structural adhesive and the weight of additional structural adhesive, which is conducive to the lightweighting of the structure. At the same time, the cross-sectional shape of the extruded profile can be optimized by topology design based on the stiffness requirements and heat dissipation requirements of the equipment to achieve good thermo-mechanical performance of the flat satellite.

[0130] (2) The first and second side rib structures of this application are designed flexibly, which can meet the functional requirements of different satellites and realize the lightweight load-bearing design of flat satellite structures, so as to better solve the contradiction between the lightweight design of flat satellite structures and insufficient mechanical load-bearing performance.

[0131] (3) The extrusion process of this application can realize the mechanical and thermal integration of the flat satellite structure. Compared with the traditional bonding method of upper and lower aluminum skin + aluminum honeycomb core, the process of this invention is simpler and more designable. At the same time, the aluminum extruded profile has excellent mechanical properties, and the extrusion molding technology is mature with low mold cost. Compared with the manufacturing method of 3D printing or the one-piece die casting process, the design and manufacturing cost of the flat satellite structure is reduced.

[0132] In the description of this application, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of the stated features. In the description of this application, "multiple" means two or more, unless otherwise explicitly specified.

[0133] In the description of this application, the word "for example" is used to mean "used as an example, illustration, or description." Any embodiment described as "for example" in this application is not necessarily to be construed as being more preferred or advantageous than other embodiments. The following description is provided to enable any person skilled in the art to make and use the invention. Details are set forth in the following description for purposes of explanation. It should be understood that those skilled in the art will recognize that the invention can be made without using these specific details. In other instances, well-known structures and processes will not be described in detail to avoid obscuring the description of the invention with unnecessary detail. Therefore, the invention is not intended to be limited to the embodiments shown, but is consistent with the broadest scope of the principles and features disclosed in this application.

[0134] The above description is merely an embodiment of the present invention and is not intended to limit the invention. Various modifications and variations can be made to the present invention by those skilled in the art. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principle of the present invention should be included within the scope of the claims of the present invention.

Claims

1. A flat-panel satellite structure that integrates heat dissipation and rigidity, characterized in that, The structure includes: multiple main extrusion structures, a first side rib structure, a second side rib structure, heat pipes, and a load-bearing base; The main extruded structure is an aluminum extruded profile with a ribbed cross section; multiple main extruded structures are spliced ​​together to form a flat satellite structure; The geometry of the ribbed cross section of the main extrusion structure is determined by a multi-objective topology optimization method. The multi-objective topology optimization method takes the minimum of the total flexibility of coupled thermal flexibility and flexibility as the objective function, including: performing thermal-mechanical coupled topology optimization based on the heat dissipation and stiffness requirements of the flat satellite structure, using thermal flexibility and flexibility to measure heat dissipation and stiffness respectively, and coupling them into a total flexibility c, with the optimization objective being to minimize the total flexibility c, thereby obtaining an extrusion profile cross-sectional shape with better performance under certain working conditions. The first side rib structure and the second side rib structure are located on the first side and the second side of the main extrusion structure, respectively; The heat pipe is manufactured by an extrusion molding process and is fixedly connected to the main extrusion structure; the heat pipe is used for uniform heat dissipation. The load-bearing seat is fixedly connected to the end of the flat satellite structure; The heat pipe is made of aluminum extrusion and is connected to the main extrusion structure by friction stir welding. The method for obtaining the integrated flat-panel satellite structure that considers both heat dissipation and stiffness includes: Step S1: Perform finite element modeling on the flat-panel satellite structure to determine the design domain, loads, and boundary conditions; Step S2: Construct a material interpolation model; Step S3: Based on the BESO algorithm, establish a multi-objective topology optimization problem with the objective function of minimizing the total compliance coupled with structural compliance and thermal compliance. Step S4: Perform finite element analysis on the heat transfer of the flat-panel satellite structure to obtain nodal temperature data; Step S5: Perform thermo-mechanical coupling analysis on the flat-panel satellite structure to obtain node displacement data; Step S6: Calculate the sensitivity of the objective function based on the node temperature data and node displacement data; Step S7: Based on the sensitivity of the objective function, the design variables are updated and iterated using a gradient-based optimization algorithm; Step S8: Determine whether the multi-objective topology optimization problem has converged. If it has not converged, return to step S2; if it has converged, the calculation is complete. Among them, the multi-objective topology optimization problem is established as a thermo-mechanical coupled multi-objective topology optimization problem: ； ； ； in, Indicates the total number of finite element elements; Indicates overall compliance; Indicates the first Topological variables of each unit; and All represent weighting coefficients; ; For thermal flexibility, This indicates that the minimum thermal flexibility value is taken; This indicates the maximum value of the thermal flexibility; For softness; This indicates taking the minimum flexibility value; This indicates taking the maximum flexibility value; This indicates that the constraints are met; Due to volume constraints, Indicates the number of finite element elements; The initial volume of the element; The heat flux input column matrix; and Top right corner Indicates matrix transpose; This is a column matrix of node temperatures. This is a column matrix of nodal displacements. This is a column matrix of nodal force loads.

2. The flat-panel satellite structure considering both heat dissipation and rigidity as described in claim 1, characterized in that, The first side rib structure is connected to the main extrusion structure by an integral extrusion method; the second side rib structure is formed by milling on the side of the main extrusion structure away from the first side rib structure.

3. The flat-panel satellite structure considering both heat dissipation and rigidity as described in claim 1, characterized in that, The main extrusion structure includes an outer surface flat plate structure and an inner ribbed structure, with the inner ribbed structure disposed inside the outer surface flat plate structure.

4. The flat-panel satellite structure considering both heat dissipation and rigidity as described in claim 1, characterized in that, The cross-sectional shape of the first side rib structure is L-shaped or T-shaped.

5. The flat-panel satellite structure considering both heat dissipation and rigidity as described in claim 1, characterized in that, The second side rib structure is an orthogonal grid structure.

6. The flat-panel satellite structure considering both heat dissipation and rigidity as described in claim 1, characterized in that, Multiple main extrusion structures of the same length or multiple main extrusion structures of different lengths are spliced ​​together to form a flat satellite structure.

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