Satellite assembly model and finite element simulation method and system thereof

Abstract

The application provides a satellite assembly model and a finite element simulation method and system thereof, the simulation method comprising: building a satellite assembly model in three-dimensional software; the satellite assembly model is characterized by shell elements; importing the built satellite assembly model into a finite element pre-processing software, and performing feature pre-processing on the satellite assembly model; using a finite element analysis software to apply constraint conditions to the satellite assembly model after feature pre-processing and solving based on the shell elements; using a visualization software to view the solving result, and if the solving result does not meet the design requirements, rebuilding the satellite assembly model. The finite element simulation method provided by the application is suitable for analyzing the satellite assembly model, and the accuracy of the analysis result is high.

Description

TECHNICAL FIELD

[0001] The application belongs to the field of finite element simulation, and particularly relates to a satellite assembly model and a finite element simulation method and simulation system thereof. BACKGROUND

[0002] With the increasing demand for satellite in-orbit application, commercialization, low cost and fast deployment have become important directions of satellite scheme design. In recent years, the launch mode of one rocket with multiple satellites greatly improves the launch efficiency. The scheme of one rocket with multiple satellites adopts a satellite assembly in a stacking mode to replace the traditional combination scheme of a satellite and a satellite adapter, which can save the launch vehicle capacity and fully utilize the fairing space.

[0003] Due to the large size of the satellite assembly and the special material, the cost of the auxiliary equipment for ground testing is high, and therefore CAE (computer aided engineering) has become a powerful tool for current aerospace design and research. The present inventors have found in the research process that the finite element simulation method in CAE can have high precision in solving the natural frequency, statics and dynamics and the like, and the analysis using the finite element can accelerate the engineering development iteration and reduce the cost and increase the efficiency; however, the finite element analysis method in the prior art has problems of low applicability to the satellite assembly analysis and low precision of the analysis result. SUMMARY

[0004] To at least partially overcome the problems in the related art, the application provides a satellite assembly model and a finite element simulation method and simulation system thereof.

[0005] According to a first aspect of the embodiments of the application, the application provides a satellite assembly model, which adopts a three-dimensional structure model in a stacking mode, and includes 2n flat plate satellites, a satellite-rocket connecting plate and a compression release mechanism.

[0006] The 2n flat plate satellites are divided into two flat plate satellite stacks, the two flat plate satellite stacks are symmetrically arranged on the satellite-rocket connecting plate, and the two flat plate satellite stacks are positioned and connected by the compression release mechanism; and the satellite-rocket connecting plate is used to connect with a launch vehicle.

[0007] In the above satellite assembly model, each of the flat plate satellite stacks includes n flat plate satellites, the n flat plate satellites are stacked together along the height direction thereof, and each of the flat plate satellites has a consistent orientation.

[0008] Further, the flat plate satellite comprises an upper side plate, a lower side plate, a left side plate, a right side plate, a satellite bottom plate and an intermediate plate; wherein the upper side plate, the lower side plate, the left side plate, the right side plate and the satellite bottom plate form a semi-closed structure with an open top, the upper side plate and the lower side plate are parallel, the left side plate and the right side plate are symmetrically connected to two ends of the upper side plate and the lower side plate, and the intermediate plate is arranged in the semi-closed structure.

[0009] Further, the intermediate plate adopts an H-shaped structure, two ends of two vertical plates in the intermediate plate are connected to the upper side plate and the lower side plate respectively, and a horizontal plate in the intermediate plate is parallel to the upper side plate.

[0010] Further, a cylinder is arranged at an outer side of the upper side plate, at a connection between the left side plate and the lower side plate, and at a connection between the right side plate and the lower side plate; and the pressing and releasing mechanism comprises the cylinders arranged on the flat plate satellite.

[0011] According to a second aspect of the embodiment of the present application, the present application further provides a finite element simulation method of a satellite assembly model, which comprises the following steps:

[0012] building the satellite assembly model in any one of the above aspects in a three-dimensional software; the satellite assembly model is characterized by shell elements;

[0013] importing the built satellite assembly model into a finite element pre-processing software, and performing feature pre-processing on the satellite assembly model;

[0014] applying constraint conditions to the satellite assembly model after feature pre-processing by using a finite element analysis software, and solving based on shell elements;

[0015] viewing the solving result by using a visualization software, and rebuilding the satellite assembly model if the solving result does not meet the design requirements.

[0016] In the finite element simulation method of the satellite assembly model, the feature pre-processing on the satellite assembly model comprises:

[0017] deleting components that do not affect calculation, deleting round corners, LOGO and process gaps;

[0018] selecting entity element, shell element and beam element analysis methods.

[0019] Further, the applying of constraint conditions comprises assigning materials to the satellite assembly model after feature pre-processing, grouping, dividing grids, applying boundary conditions and setting expected output results;

[0020] The assigning of materials to the satellite assembly model after feature pre-processing comprises defining isotropy and anisotropy of materials;

[0021] The grouping of the satellite assembly model after feature pre-processing comprises: dividing according to the same material and the same boundary condition;

[0022] The grid dividing manner comprises: dividing according to hexahedron, tetrahedron of solid element or mixed dividing; or dividing according to quadrangle, triangle, right triangle of shell element or mixed dividing;

[0023] The boundary condition applying comprises: contact, fixed connection, fixed support, hinge, free vibration and external load;

[0024] The setting of the expected output result comprises: setting natural frequency, vibration mode, stress deformation and fatigue calculation.

[0025] In the finite element simulation method of the satellite assembly model, the software for building the satellite assembly model is Solidworks, the finite element pre-processing software is Hypermesh, the finite element analysis software is Nastran, and the visualization software is Hyperview.

[0026] According to a third aspect of the embodiments of the present application, the present application further provides a finite element simulation system of a satellite assembly model, which comprises a memory, a processor and a computer program stored in the memory and running on the processor, and the processor implements the finite element simulation method of the satellite assembly model according to any one of the above when processing the computer program.

[0027] According to the above specific embodiments of the present application, at least the following beneficial effects are obtained: the present application builds a satellite assembly model, and determines the boundary condition of finite element simulation according to the launch working condition of the satellite assembly, that is, determines the finite element simulation method according to the working characteristics of the stacked satellite co-compression cylinder, the layer-by-layer arrangement of the satellite and the separate release, so that the analysis result is more accurate, thereby improving the simulation accuracy.

[0028] According to the satellite body structure characteristics, the present application uses a shell element to represent the satellite assembly model, so that the number of grids in the finite element simulation is greatly reduced, which is beneficial to the calculation power distribution of large finite element simulation such as satellite-rocket coupling.

[0029] It should be understood that the above general description and the following specific embodiments are only exemplary and explanatory, and cannot limit the scope of the present application. BRIEF DESCRIPTION OF DRAWINGS

[0030] The following accompanying drawings are part of the specification of the present application, which show the embodiments of the present application, and together with the description of the specification, illustrate the principles of the present application.

[0031] Figure 1This is a schematic diagram of the overall structure of a satellite assembly model provided for a specific embodiment of this application.

[0032] Figure 2 This is a schematic diagram of the structure of a flat-panel satellite in a satellite assembly model provided for a specific embodiment of this application.

[0033] Figure 3 This is one of the flowcharts for a finite element simulation method of a satellite assembly model provided for a specific implementation of this application.

[0034] Figure 4 The second flowchart illustrates a finite element simulation method for a satellite assembly model provided for a specific implementation of this application.

[0035] Explanation of reference numerals in the attached figures:

[0036] 1. Flat panel satellite; 11. Top side panel; 12. Bottom side panel; 13. Left side panel; 14. Right side panel; 15. Satellite bottom panel; 16. Middle panel;

[0037] 2. Star-rocket connecting plate;

[0038] 3. Compression release mechanism; 31. Cylinder;

[0039] 4. Interstellar gaps. Detailed Implementation

[0040] To make the objectives, technical solutions, and advantages of the embodiments of this application clearer, the spirit of the content disclosed in this application will be clearly explained below with reference to the accompanying drawings and detailed description. After understanding the embodiments of this application, any person skilled in the art can make changes and modifications based on the technology taught in this application without departing from the spirit and scope of this application.

[0041] The illustrative embodiments and descriptions provided in this application are for explaining the application, but are not intended to limit the application. Furthermore, elements / components using the same or similar reference numerals in the drawings and embodiments are used to represent the same or similar parts.

[0042] The terms “first,” “second,” etc., used in this document are not intended to specifically refer to order or sequence, nor are they used to limit this application; they are merely used to distinguish elements or operations described using the same technical terms.

[0043] The directional terms used in this article, such as up, down, left, right, front, or back, are for reference only when referring to the accompanying drawings. Therefore, the use of directional terms is for illustrative purposes and not to limit this work.

[0044] The terms “include,” “including,” “have,” “contain,” etc., used in this article are all open-ended terms, meaning that they include but are not limited to.

[0045] As used herein in the specification and in the claims,“or” is used in its inclusive sense (and not in its exclusive sense) so that when used, for example, in a list of two or more items, the term“or” means that at least one of the items, but can also include more than one of the items. For example, when a list of two or more items is preceded by the term“one or more of the items” or“at least one of the items,” each of the items can be included in the list (i.e., some items can be included more than once).

[0046] As used herein in the specification and in the claims,“a” or“an” can mean one or more than one.

[0047] As used herein in the specification and in the claims, the phrase“about” and“substantially” are used to describe amounts or errors that can vary slightly from the recited amount or error. Generally, these phrases modify the preceding value or error by no more than 10%, preferably by no more than 5%, and more preferably by no more than 1%, and even more preferably by no more than 0.1% of the recited value or error. One of skill in the art will understand that the recited values and errors can be adjusted as necessary to achieve the indicated ranges.

[0048] Certain terminology can be used in the preceding description for the purpose of reference only, which will be discussed at various places in the specification and / or claims.

[0049] With the commercialization of satellite industry, the continuous promotion of low-cost and fast deployment, the stack satellite assembly is more suitable for the current launch task demand. Finite element modeling can divide the structure into unit grids to approximate the mathematical model of the actual structure. Finite element simulation analysis can have higher precision in solving problems such as modal analysis and structural deformation. In modal analysis, the natural frequencies of the reference structure are designed to avoid resonance phenomenon, and according to the mode shape, the structural strength optimization design can be performed in a certain direction.

[0050] As shown in Figure 1 The satellite assembly model provided by the present application is a three-dimensional structure model in a stacking manner, which comprises 2n flat plate satellites 1, a satellite-rocket connecting plate 2 and a pressing release mechanism 3. The 2n flat plate satellites 1 are divided into two flat plate satellite stacks, which are symmetrically arranged on the satellite-rocket connecting plate 2, and the two flat plate satellite stacks are positioned and connected through the pressing release mechanism 3. The satellite-rocket connecting plate 2 is used to connect with the carrier rocket.

[0051] Specifically, each flat plate satellite stack comprises n flat plate satellites 1, the n flat plate satellites 1 are stacked together along the height direction thereof, and each flat plate satellite 1 faces the same direction.

[0052] As shown in Figure 2As shown in the drawings, each flat panel satellite 1 comprises an upper side plate 11, a lower side plate 12, a left side plate 13, a right side plate 14, a satellite bottom plate 15 and an intermediate plate 16. The upper side plate 11, the lower side plate 12, the left side plate 13, the right side plate 14 and the satellite bottom plate 15 form a semi-closed structure with a top opening, the upper side plate 11 and the lower side plate 12 are parallel, and the left side plate 13 and the right side plate 14 are symmetrically arranged at the two ends of the upper side plate 11 and the lower side plate 12. The left side plate 13 and the right side plate 14 can adopt a stepped structure.

[0053] The intermediate plate 16 is arranged in the semi-closed structure. The intermediate plate 16 adopts an H-shaped structure, the two vertical plates in the intermediate plate 16 are respectively connected with the upper side plate 11 and the lower side plate 12 at the two ends, and the horizontal plate in the intermediate plate 16 is parallel to the upper side plate 11. The intermediate plate 16 is used for isolating part of the in-orbit single machine fixed on the satellite bottom plate 15, and also used for providing a fixable position for part of the in-orbit single machine and electrical and thermal structures. At the same time, the design of the intermediate plate 16 can improve the structural rigidity of the whole flat panel satellite 1.

[0054] A cylinder 31 is arranged at the outer side of the upper side plate 11, at the connection between the left side plate 13 and the lower side plate 12, and at the connection between the right side plate 14 and the lower side plate 12. The flat panel satellites 1 in the same flat panel satellite stack are positioned and connected through the cylinder 31 arranged at the outer side of the upper side plate 11, and the flat panel satellites 1 in different flat panel satellite stacks are positioned and connected through the cylinders 31 arranged at the connection between the left side plate 13 and the lower side plate 12 and at the connection between the right side plate 14 and the lower side plate 12. The pressing and releasing mechanism 3 comprises the cylinders 31 arranged on the flat panel satellites 1, and acts on the whole satellite assembly. Each flat panel satellite 1 is connected and pressed through the cylinder 31, and when reaching the specified working condition, the pre-tightening force is unloaded, and the flat panel satellites 1 generate a separation gap through a pre-set separation energy, and the separation process is completed.

[0055] Specifically, n can be set to 16 as needed, so that there are 32 flat panel satellites 1 in the satellite assembly, and the basic structures of the flat panel satellites 1 are the same. In order to make full use of the envelope, the height of the cylinder 31 is set to half of the sum of the heights of the side plates and the inter-orbit gap 4, so that the two flat panel satellite stacks are connected in opposite directions, which can ensure that the structural size is symmetrical. From the side of the satellite assembly, the two flat panel satellite stacks are arranged in a staggered manner, which also increases the structural rigidity of the satellite assembly.

[0056] As shown in the drawings, Figure 3 Based on the satellite assembly model provided in the present application, the present application further provides a finite element simulation method of the satellite assembly model, which comprises the following steps:

[0057] S1, building the above satellite assembly model in three-dimensional software.

[0058] Specifically, the satellite assembly model can be built in the 3D software Solidworks.

[0059] S2. Import the constructed satellite assembly model into the finite element preprocessing software and perform feature preprocessing on the satellite assembly model.

[0060] The feature preprocessing for the satellite assembly model includes:

[0061] Remove components that do not affect the calculation, as well as rounded corners, logos, and process gaps.

[0062] Choose analysis methods such as solid elements, shell elements, and beam elements.

[0063] If shell elements are selected, the mid-face is extracted, and its cross-sectional shape is defined. The connection method between shell element faces is determined based on the actual working conditions.

[0064] In this application, HyperMesh software can be used as the finite element preprocessing software. HyperMesh software can be used to trim, mesh, and apply boundary conditions to the imported satellite assembly model, and can export intermediate format files for use by finite element solvers such as Nastran and Ansys.

[0065] Specifically, in HyperMesh software, components, fillets, logos, and process gaps that do not affect the calculation can be deleted from the satellite assembly model.

[0066] S3. Apply constraints to the three-dimensional structural model after feature preprocessing using finite element analysis software and solve the problem based on shell elements.

[0067] Specifically, Nastran can be used as the finite element analysis software.

[0068] Among them, such as Figure 4 As shown, applying constraints includes assigning materials and grouping the satellite composite model after feature preprocessing, meshing, applying boundary conditions, and setting the desired output results.

[0069] Specifically, the materials assigned to the satellite composite model after feature preprocessing include:

[0070] Define the isotropic and anisotropic properties of materials;

[0071] If a material is defined as an isotropic material, then its Young's modulus, Poisson's ratio, density, coefficient of thermal expansion, etc., are defined.

[0072] Specifically, when grouping satellite composite models after feature preprocessing, they can be divided based on factors such as the same material and the same boundary conditions.

[0073] If a shell element is used, grouping is performed according to the same middle surface cross-sectional shape.

[0074] Specifically, the grid division method includes:

[0075] Division is performed according to hexahedron, tetrahedron of solid elements, or mixed division, etc.

[0076] Division is performed according to quadrilateral, triangle, right triangle of shell elements, or mixed division, etc.

[0077] Specifically, boundary conditions are applied to the grouped three-dimensional structure model, including:

[0078] Contact, solid connection, fixed support, hinge, free vibration, and external load, etc.

[0079] Specifically, the desired output result includes:

[0080] Natural frequency, mode shape, stress deformation, and fatigue calculation, etc.

[0081] In the prior art, satellite structure analysis is mostly performed by using solid elements for finite element modeling, because the number of satellites launched each time is mostly no more than 6. With the rise of stack satellite assemblies at the present stage, the number of satellites carried by each rocket is multiplied, and the grid division time and the number of grids will be greatly increased if each satellite in the assembly is divided by using solid elements, which will increase the running time and difficulty of subsequent analysis. Therefore, in the finite element analysis of the satellite assembly, shell elements are selected for analysis in the present application.

[0082] Since the thickness value of the satellite bottom plate 15 and the thickness value of each side plate are much smaller than the length value and the width value of the flat plate satellite 1, the satellite bottom plate 15 and each side plate can be characterized by using shell elements. The shell element has only one layer of grid in the thickness direction, and compared with the solid element, the shell element can reduce the total number of grids, thereby greatly reducing the difficulty of subsequent grid checking and improving the calculation speed.

[0083] At the same time, the design of the satellite assembly also needs to consider the coupling analysis with the launch vehicle, and the volume and mass of the satellite assembly structure are much smaller than the volume and mass of the launch vehicle. Therefore, the satellite assembly is characterized by using shell elements, which is more reasonable to characterize the finite element model of the entire satellite-rocket structure under the satellite-rocket coupling condition.

[0084] In the three-dimensional model of the satellite assembly of the present application, the shape and size of each flat plate satellite 1 are completely the same, the thickness of each side plate of each flat plate satellite 1 and the thickness of each cylinder 31 are the same, and the thickness of each side plate and the thickness of each cylinder 31 are represented by t1; the thickness of each satellite bottom plate 15 is different from the thickness t1 of each side plate, and the thickness of the satellite bottom plate 15 is represented by t2.

[0085] Divide mid-surface by using midmesh function in HyperMesh software; first, divide mid-surface of the component with thickness t1 and modify the name; then, divide mid-surface of the component with thickness t2. Define material properties in Materials.

[0086] In Properties option, define Card Image as PSHELL, i.e. shell element property; fill thickness t1 and t2 in T to define cross-section properties of each group of shell elements; select defined material properties in Materials, so that each surface represents a body with thickness t1 or t2.

[0087] In quick edit, through surface edge color, it can be judged that each surface edge belongs to free edge, shared edge, suppressed edge or T-type edge. According to the actual working condition of the satellite assembly, the connection mode between the surfaces of the shell elements is defined. The contact edge of the cylinder 31 provided on the different plate satellite 1 is a shared edge, the connection between the plate satellite 1 and the cylinder 31 is a T-type edge, the basic three surfaces of the plate satellite 1 are connected, the contact edge between the pressing and releasing mechanism 3 and the plate satellite 1 is a T-type edge, and the connection between the cylinder 31 and the satellite-rocket connecting plate 2 is a T-type edge. According to the setting of boundary conditions, the shell elements which are in contact but not fixedly connected and in contact and fixedly connected are divided, so that the analysis result is more accurate.

[0088] Mesh division is carried out by using automesh, and mesh division mode and mesh size are selected. The quality of the mesh is automatically checked by using elemcleanup, and the unqualified mesh is modified one by one by using qualityindex after checking, and the unqualified mesh under different requirement levels is displayed by controlling displaythresholds.

[0089] In this application, constraint modal analysis of basic structure design is taken as an example, and load and load step are set. A new loadcol1 is created in LoadCollectors. The calculation method of constraint modal is defined in loadcol1, and the upper and lower limits of modal value calculation result are defined; a new loadcol2 is created, and the constraint mode of constraint modal is defined in loadcol2.

[0090] In this application, the satellite-rocket connecting plate 2 is fixedly connected with the carrier rocket, i.e. fixedly immovable. The six degrees of freedom of the satellite-rocket connecting plate 2 are constrained by using constraints to fix the satellite-rocket connecting plate 2 with the carrier rocket.

[0091] New Load Steps, loadstep1, select analysis type as Normal modes, modal analysis. SPC, loadcol2, define constraint method, method(struct), loadcol1, define calculation method.

[0092] Select control card type in control cards, used to control solving strategy and output type. In this application, add SOL, PARAM, GLOBAL_OUTPUT_REQUEST cards. In SOL, define Anslysis as Normal Modes, modal analysis. Then define PARAM, GLOBAL_OUTPUT_REQUEST cards according to expected results. In PARAM, set POST_V1 as 0, then the result file after solving by finite element analysis software Nastran is.op2 format.

[0093] Export the finite element model set in HyperMesh as.bdf format file, open and run solving in finite element analysis software Nastran, generate.op2 file.

[0094] Open the.op2 file and the generated.bdf file in visualization software Hyperview, then you can view the solving results, view each order modal and mode shape. According to the simulation results, design iteration.

[0095] S4, use visualization software to view the solving results, if the solving results do not meet the design requirements, return to step S1 according to the calculation results, re-iterate and optimize the satellite assembly model.

[0096] Specifically, the visualization software can use Hyperview.

[0097] The satellite assembly structure is large in size and complex in application conditions. It has great acceleration and vibration in the rocket launch stage, and no gravity in the on-orbit stage, which are difficult to simulate experiment on the ground. However, by selecting CAE simulation design, the test conditions can be quickly configured, the experimental objects can be quickly iterated, and high-precision analysis results can be obtained at low cost. The analysis results can predict the stress state and displacement deformation, and the analysis process decomposes the structure into several simple polygon shapes, and combines the stiffness matrix of each polygon, which can represent the stiffness matrix of the entire structure. By solving the stiffness matrix of each simple polygon under boundary conditions, the displacement is obtained, and the stress and displacement state between polygons are obtained, and then the stress analysis results and deformation results of the entire structure are obtained.

[0098] The satellite assembly finite element simulation method for stack launching provided in the application is based on the HyperMesh and Nastran software commonly used in aerospace systems; the boundary conditions are determined according to the satellite assembly launching conditions, that is, the simulation method is defined according to the working characteristics of the stack satellite common compression cylinder, the satellite is arranged layer by layer, and is released respectively, which can make the analysis result more accurate. According to the satellite body structure characteristics, the satellite substrate is represented by a shell element, which greatly reduces the grid number of the finite element simulation, and is beneficial to the calculation power distribution of large finite element simulation such as satellite-rocket coupling.

[0099] In the example embodiment, based on the satellite assembly finite element simulation method provided in the embodiment of the application, the embodiment of the application further provides a satellite assembly finite element simulation system, which comprises a memory and a processor coupled to the memory, and the processor is configured to execute the satellite assembly finite element simulation method in any one of the embodiments of the application based on the instructions stored in the memory. The memory can be a system memory or a fixed non-volatile storage medium, etc., and the system memory can store operating systems, application programs, boot loaders, databases and other programs, etc.

[0100] It should be noted that the satellite assembly finite element simulation system and the satellite assembly finite element simulation method provided in the above embodiments belong to the same concept, and the specific implementation process is detailed in the method embodiments, which will not be repeated here.

[0101] In the example embodiment, the embodiment of the application further provides a computer storage medium, which is a computer readable storage medium, for example, a memory including a computer program, and the above computer program can be executed by a processor to complete the satellite assembly finite element simulation method in any one of the embodiments of the application.

[0102] The above embodiments of the application can be implemented in various hardware, software coding or combination of both. For example, the embodiments of the application can also represent program codes for executing the above methods in a data signal processor. The application can also relate to various functions performed by computer processors, digital signal processors, microprocessors or field programmable gate arrays. The above processors can be configured according to the application to perform specific tasks, which are completed by executing machine readable software codes or firmware codes defining specific methods disclosed in the application. Software codes or firmware codes can be developed to represent different program languages and different formats or forms. Different target platforms can also be compiled with software codes. However, the software codes for performing tasks according to the application and other types of configuration codes are different in code style, type and language, which do not deviate from the spirit and scope of the application.

[0103] The above description is only the specific implementation of the present application, and any equivalent changes and modifications made by any person skilled in the art without departing from the concept and principles of the present application shall fall within the scope of protection of the present application.

Claims

1. A satellite assembly model, characterized in that, The three-dimensional structural model adopts a stacked approach, which includes 2n flat-panel satellites, a star-rocket connection plate, and a clamping and releasing mechanism; Each of the 2n flat-panel satellites is divided into two flat-panel satellite stacks, which are symmetrically arranged on the satellite-rocket connection plate. The two flat-panel satellite stacks are positioned and connected by the clamping and releasing mechanism. The satellite-rocket connection plate is used to connect with the launch vehicle. The flat-panel satellite includes an upper side plate, a lower side plate, a left side plate, a right side plate, a satellite base plate, and a middle plate; wherein the upper side plate, lower side plate, left side plate, right side plate, and satellite base plate form a semi-enclosed structure with an open top, the upper side plate and lower side plate are parallel, the left side plate and right side plate are symmetrically connected to the two ends of the upper side plate and lower side plate, and the middle plate is disposed in the semi-enclosed structure; A cylinder is provided on the outer side of the upper side plate, at the connection between the left side plate and the lower side plate, and at the connection between the right side plate and the lower side plate; the clamping release mechanism includes the cylinders provided on each of the flat satellites; Each flat satellite in the same flat satellite stack is positioned and connected by a cylinder set on the outer side of the upper side plate. Each flat satellite in different flat satellite stacks is positioned and connected by a cylinder set at the connection between the left side plate and the lower side plate, and a cylinder set at the connection between the right side plate and the lower side plate.

2. The satellite assembly model according to claim 1, characterized in that, Each of the aforementioned flat-panel satellite stacks includes n flat-panel satellites, which are stacked together along their height direction, and all the flat-panel satellites face the same direction.

3. The satellite assembly model according to claim 1, characterized in that, The middle plate adopts an H-shaped structure, and the two ends of the two vertical plates in the middle plate are respectively connected to the upper side plate and the lower side plate. The horizontal plate in the middle plate is parallel to the upper side plate.

4. A finite element simulation method for a satellite assembly model, characterized in that, Includes the following steps: A satellite assembly model as described in any one of claims 1 to 3 is constructed in 3D software; the satellite assembly model is characterized using shell elements; The constructed satellite assembly model is imported into the finite element preprocessing software to perform feature preprocessing on the satellite assembly model; Constraints were applied to the satellite assembly model after feature preprocessing using finite element analysis software, and the solution was obtained based on shell elements. Use visualization software to view the solution results. If the solution results do not meet the design requirements, rebuild the satellite assembly model.

5. The finite element simulation method for the satellite assembly model according to claim 4, characterized in that, The feature preprocessing of the satellite assembly model includes: Remove components that do not affect the calculation, and delete fillets, logos, and process gaps; Choose from solid element, shell element, and beam element analysis methods.

6. The finite element simulation method for the satellite assembly model according to claim 5, characterized in that, The constraint conditions include assigning materials and grouping the satellite composite model after feature preprocessing, dividing the mesh, applying boundary conditions, and setting the desired output results; Assigning materials to the satellite composite model after feature preprocessing includes: defining the isotropic and anisotropic properties of the materials; The grouping of the satellite assembly model after feature preprocessing includes: dividing it according to the same material and the same boundary conditions; The meshing methods include: dividing according to the hexahedrons, tetrahedrons, or mixed division of solid elements; dividing according to the quadrilaterals, triangles, right triangles, or mixed division of shell elements; the applied boundary conditions include: contact, fixed connection, fixed support, hinge, free vibration, and external load; The desired output results include: setting the natural frequency, mode shape, stress deformation, and fatigue calculation.

7. The finite element simulation method for the satellite assembly model according to claim 4, characterized in that, The software used to build the satellite assembly model is Solidworks, the finite element preprocessing software is Hypermesh, the finite element analysis software is Nastran, and the visualization software is Hyperview.

8. A finite element simulation system for a satellite assembly model, characterized in that, The system includes a memory, a processor, and a computer program stored in the memory and running on the processor, wherein the processor, when processing the computer program, implements a finite element simulation method for the satellite assembly model as described in any one of claims 4 to 7.

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