A modular design method for mechanical and thermal integration of spacecraft in extreme environments

Abstract

The present invention discloses a modular design method for mechanical and thermal integration of spacecraft in extreme environments. The method obtains the initial structure in advance through multi-objective optimization, and then comprehensively considers the actual working conditions of mechanical and thermal integration in extreme environments, various factors such as heat treatment protection and structural micro-deformation, and repeatedly iterates and calculates through CAE simulation to determine the final size parameters and structural scheme. At the same time, it also takes into account various types of spacecraft payloads to meet the needs of diversified spacecraft payloads. The present invention realizes the reconfigurable combination of structures, considers the mechanical and thermal integration design in extreme environments, and considers the influence of the surface treatment of materials on thermal radiation. The spacecraft platform designed using this design method can be used for ground verification tests, as well as for the modular base of future space-based platforms, and can be equipped with multiple types of payloads, such as cameras, antennas, special plane payloads, etc.

Description

Technical Field

[0001] The present application relates to the field of spacecraft platform technology, and in particular to a modular design method for mechanical and thermal integration of spacecraft for use in extreme environments. Background Art

[0002] In recent years, spacecraft missions have become increasingly complex and require increasingly diverse payloads. In order to meet and adapt to this trend, ground-based experimental platforms need to be designed with the goals of extreme, modular, and thermal-mechanical coupling. Currently, experimental platforms for spacecraft have not fully considered the types and materials of external payloads contained in satellites, the overall structure of satellites, and the simulation of the vacuum and coldness of outer space in similar satellite operating environments.

[0003] At the same time, with the development of strong external heat flux technology, the demonstration of equipment damage effects under strong external heat flux irradiation and the development of experimental prototypes have also received more and more attention. In the early stage, domestic and foreign scholars' research on strong external heat flux test technology mainly focused on heat flux power, irradiation area, incident angle, etc., and did not consider the different damage effects of different irradiated targets.

[0004] In summary, there is currently no research on platforms and design methods for simulating various effect loads of spacecraft in a vacuum and dark cold environment and verifying irradiation damage conditions under strong external heat flux irradiation. Summary of the invention

[0005] The purpose of this application example is to provide a modular design method for spacecraft mechanical and thermal integration in extreme environments, so as to solve the technical problem that the related technology fails to fully consider the spacecraft operating conditions and the different damage effects on different target loads under strong external heat flux.

[0006] The technical solution adopted by the present invention is as follows:

[0007] A modular design method for mechanical and thermal integration of spacecraft in extreme environments, the method is a mechanical and thermal integration platform design method that comprehensively considers extreme environments, thermal treatment protection and structural micro-deformation, and includes the following steps:

[0008] Step 1: Based on the load requirements and the target parameters of the working conditions to be met, the initial structure of the mechanical and thermal integration platform is designed under constraints;

[0009] Step 2: constructing a three-dimensional solid model based on the initial structure and performing pre-processing on it;

[0010] Step 3: A CAE simulation model is established based on the initial structural 3D solid model after pre-processing, and a thermal-mechanical coupling simulation calculation is performed on the model. In this process, not only the relevant changes in thermal properties are considered, but also the fixed constraints and mechanical effects on the structure are considered;

[0011] Step 4: Obtain the overall deformation of the initial structure, the overall stress parameter value of the initial structure, and the strength value of the key connection parts through simulation, and analyze and verify whether the design scheme of the initial structure is feasible;

[0012] Step 5: If the design scheme is feasible, the design scheme is directly output; otherwise, the design scheme is adjusted according to the characterization parameters of the overall structural deformation, the overall structural stress parameter value, and the strength value of the key connection parts, and the initial structural design is carried out again in step 1.

[0013] In the above technical solution, further, in step 1, the specific steps of designing the initial structure are as follows:

[0014] Step 1.1, determine the load of the initial structure and the target parameters of the working condition that need to be met;

[0015] The initial structure includes the following loads: one or more antenna loads; one or more carbon fiber loads; one or more thermal control multi-layer loads; one or more camera loads; one or more ground information processing platforms; one or more batteries;

[0016] The working condition target parameters include: the storage ratio of the initial structure as a whole, the center of gravity coordinates of the initial structure as a whole, the center of gravity coordinates of the antenna load, the center of gravity coordinates of the carbon fiber load, the center of gravity coordinates of the thermal control multi-layer load, the center of gravity coordinates of the camera load, the center of gravity coordinates of the ground information processing platform, and the center of gravity coordinates of the battery;

[0017] Step 1.2, determine the constraint parameters that the initial structure needs to satisfy;

[0018] The constraint parameters include: the deformation of the initial structure, the outer envelope size of the initial structure, and the weight of the initial structure;

[0019] Step 1.3, determine the design variables of the initial structure;

[0020] The design variable parameters include: connection strength between parts inside the initial structure, connection hole positions inside the initial structure;

[0021] Step 1.4, taking the design variable parameters as input, setting multiple groups of different working condition parameters as targets, constructing a multi-objective optimization solution model under constraints, and using a multi-objective solution algorithm to optimize multiple targets until the structure converges to obtain the optimal design variable solution set;

[0022] Step 1.5: extract the optimal design variable solution set, design the initial structure, and obtain the initial structure of the platform to be verified.

[0023] Furthermore, the storage ratio refers to the percentage of the volume in the stored and combined state to the volume in the working unfolded state; the center of gravity coordinates are the coordinate values ​​of the center of gravity in three directions of XYZ.

[0024] Furthermore, the following constraints are introduced simultaneously when determining the design variables of the initial structure in step 1.3: the connection strength value is not greater than the allowable stress of the material used; the center coordinates of the connection hole position do not exceed the outer envelope of the initial structure; the diameter value of the connection hole matches the size of the selected bolt connection.

[0025] Furthermore, in step 1.4, the optimization objective function expression constructed is:

[0026] minR(Δ)=min{r1=-k s (Δ),r2=-k z (Δ)}

[0027] Among them, R (Δ) is the objective function of the initial structure storage ratio and the initial structure overall center of gravity coordinates, r1 is the initial structure storage ratio, r2 is the initial structure overall center of gravity coordinates, k s represents the optimization function considering the minimum storage ratio condition, k z represents the optimization function considering the lowest coordinates of the center of gravity, and Δ is the function of the optimized design variables;

[0028] The expression of the initial structural design variable is:

[0029] Δ=(F(d),N(d))

[0030] Among them, Δ is the function value of the optimized design variable, F(d) is the preload strength provided by the connecting bolts at the load connection, N(d) is the number of connecting holes at the load connection, and d is the numerical value of the hole diameter of the connecting hole;

[0031] The constraint conditions are the deformation of the initial structure, the outer envelope size of the initial structure and the total weight of the initial structure, and the expression is:

[0032]

[0033] If p1>0, then k1=0

[0034] If p1≤0, then k1=1

[0035] Among them, p1 is the initial structural deformation coefficient, λ deflection is the target value of initial structural deformation, λ deflecti on0 is the value after optimization iteration of the initial structural deformation, and k1 is the control coefficient of the penalty term of the initial structural deformation;

[0036]

[0037] If p2>0, then k2=0

[0038] If p2≤0, then k2=1

[0039] Among them, p2 is the initial structure outer envelope size coefficient, λ outsize is the target value of the initial structure outer envelope size, λ outsize_0 is the value after iterative optimization of the initial structure outer envelope size, k2 is the penalty term control coefficient of the initial structure outer envelope size;

[0040]

[0041] If p3>0, then k3=0

[0042] If p3≤0, then k3=1

[0043] Among them, p3 is the initial structure weight limit coefficient, λ weight is the target value of the initial structure weight, λ weight_0 is the value after iterative optimization of the initial structure weight, and k3 is the control coefficient of the initial structure weight penalty term.

[0044] Furthermore, the three-dimensional solid model of the initial structure is pre-processed, including the following steps:

[0045] Step 2.1, use modeling software to perform geometric processing on the three-dimensional solid model of the initial structure to meet the calculation accuracy, delete small features in the structure, and avoid mesh distortion;

[0046] Step 2.2, use the finite element meshing software to mesh the initial structure three-dimensional solid model, the meshing level is fine, and the meshing type is tetrahedral mesh.

[0047] Furthermore, a CAE simulation model is established based on the initial structural three-dimensional solid model after pre-processing, and a thermal-mechanical coupling simulation calculation is performed on the model, specifically including the following:

[0048] The external temperature of the initial structure can be adjusted according to actual needs. The extreme environmental temperature includes three different types: extremely high temperature, extremely low temperature and extreme high and low temperature sudden change. CAE simulation model is established and simulation calculation is performed for each type of working condition. The extremely high temperature means that the external environmental temperature is in the range of [0℃, 500℃], the extremely low temperature means that the external environmental temperature is in the range of [-196℃, 0℃], and the extreme high and low temperature sudden change means that the external environmental temperature changes between the extremely high temperature range and the extremely low temperature range.

[0049] Furthermore, the establishment and simulation of the CAE simulation model under extreme low temperature conditions include the following:

[0050] Determination of external boundary conditions:

[0051] S3.1.1. Specify the initial temperature of the external environment at the initial time point;

[0052] S3.1.2. Specify the target temperature value of the external environment at the target time point;

[0053] S3.1.3. Specify the duration of the initial structural external extreme low temperature;

[0054] Determination of internal boundary conditions:

[0055] S3.1.4. Specify that the thermal protection measure under extremely low temperature is electronically controlled active heating and temperature regulation;

[0056] S3.1.5. Specify the temperature of each load under the electronically controlled active heating protection;

[0057] S3.1.6. Specify the initial structure temperature under the electronically controlled active heating protection;

[0058] S3.1.7. Clarify the mechanical boundary conditions of the initial structure, including the initial structure fixed constraints and gravity boundary conditions;

[0059] S3.1.8. Simulate and calculate and output the results.

[0060] Furthermore, the establishment and simulation of the CAE simulation model under extreme high temperature conditions include the following:

[0061] Determination of external boundary conditions:

[0062] S3.2.1. Specify the initial temperature of the external environment at the initial time point;

[0063] S3.2.2. Specify the target temperature value of the external environment at the target time point;

[0064] S3.2.3. Specify the duration of the initial structural external extreme heat;

[0065] Determination of internal boundary conditions:

[0066] S3.2.4. Specify that the thermal control protection measures under extreme high temperatures are thermal insulation protection and sandblasting reflection protection;

[0067] S3.2.5. Specify the temperature of each load under thermal insulation protection;

[0068] S3.2.6. Specify the initial structural temperature under sandblasting reflection;

[0069] S3.2.7. Clarify the mechanical boundary conditions of the initial structure, including the initial structure fixed constraints and gravity boundary conditions;

[0070] S3.2.8. Simulate and calculate and output the results.

[0071] Furthermore, the establishment and simulation of the CAE simulation model under extreme high and low temperature sudden change conditions include the following:

[0072] Determination of external boundary conditions:

[0073] S3.3.1 Specify the minimum external ambient temperature;

[0074] S3.3.2 Specify the maximum temperature of the external environment;

[0075] S3.3.3 Specify the change curve and frequency of sudden high and low temperature changes;

[0076] Determination of internal boundary conditions:

[0077] S3.3.4 Low temperature section, the temperature of each load after starting the electronic control active heating;

[0078] S3.3.5 Low temperature section, initial structure temperature after starting electronically controlled active heating;

[0079] S3.3.6 High temperature section, temperature of each load under thermal insulation protection;

[0080] S3.3.7 High temperature section, initial structural temperature under sandblasting reflection;

[0081] S3.3.8 Define the mechanical boundary conditions of the initial structure, including the initial structure fixed constraints and gravity boundary conditions;

[0082] S3.3.9 Perform simulation calculations and output the results.

[0083] Compared with the prior art, the present invention has the following beneficial effects:

[0084] The present application provides a modular design method for mechanical and thermal integration of spacecraft in extreme environments. On the one hand, the design method realizes the optimization design under multiple objectives and multiple constraints. By setting the working condition target parameters, constraint parameters and design variable parameters, the design parameter solution model is constructed, and multiple optimization objectives are optimized to obtain a solution set with multiple target optimal values, thereby designing and forming the initial structure. On the other hand, the design method comprehensively considers the actual working conditions of mechanical and thermal integration under extreme environments, considers various factors such as heat treatment protection and structural micro-deformation, and determines the final size parameters and structural scheme through CAE simulation calculation. On the third hand, the design method also takes into account various types of spacecraft payloads to meet the diversified needs of spacecraft payloads. The design method realizes the reconfigurable combination of structures, considers the mechanical and thermal integration design under extreme environments, and considers the influence of the surface treatment of materials on thermal radiation. The spacecraft platform designed using this design method can be used for ground verification tests, as well as for the modular base of future space-based platforms, and can be equipped with multiple types of payloads, cameras, antennas, special plane payloads, etc.

[0085] It is to be understood that both the foregoing general description and the following detailed description are exemplary and explanatory only and are not restrictive of the invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0086] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the present application and, together with the description, serve to explain the principles of the present application.

[0087] The drawings are only for the purpose of illustrating particular embodiments and are not to be construed as limiting the invention.

[0088] Figure 1 It is a schematic diagram of the whole process steps of the modular design method of mechanical and thermal integration of spacecraft for extreme environments;

[0089] Figure 2 It is a flow chart of the method for obtaining the initial structure in the design method;

[0090] Figure 3 It is a schematic diagram of the simulation process for the initial structure in extremely low temperature conditions;

[0091] Figure 4 It is a schematic diagram of the simulation process for the initial structure under extremely high temperature conditions;

[0092] Figure 5 It is a schematic diagram of the simulation process for the initial structure under extreme high and low temperature sudden change conditions;

[0093] Figure 6 It is a comparison chart of the efficiency of the design method of the present invention and the traditional design method in achieving the design goal. DETAILED DESCRIPTION

[0094] Exemplary embodiments will be described in detail herein, examples of which are shown in the accompanying drawings. When the following description refers to the drawings, the same numbers in different drawings represent the same or similar elements unless otherwise indicated. The embodiments described in the following exemplary embodiments do not represent all embodiments consistent with the present application. Instead, they are merely examples of devices and methods consistent with some aspects of the present application as detailed in the appended claims.

[0095] The terms used in this application are for the purpose of describing specific embodiments only and are not intended to limit this application. The singular forms of "a", "said" and "the" used in this application and the appended claims are also intended to include plural forms unless the context clearly indicates other meanings. It should also be understood that the term "and / or" used herein refers to and includes any or all possible combinations of one or more associated listed items.

[0096] It should be understood that although the terms first, second, third, etc. may be used in the present application to describe various information, these information should not be limited to these terms. These terms are only used to distinguish the same type of information from each other. For example, without departing from the scope of the present application, the first information may also be referred to as the second information, and similarly, the second information may also be referred to as the first information. Depending on the context, the word "if" as used herein may be interpreted as "at the time of" or "when" or "in response to determining".

[0097] A specific embodiment of the present invention, as Figure 1 As shown, a modular design method for mechanical and thermal integration of spacecraft in extreme environments is disclosed. The method is a mechanical and thermal integration design method that takes into account extreme environments, thermal treatment protection, and structural micro-deformation. In order to illustrate the effectiveness of the method proposed by the present invention, the above technical solution of the present invention is described in detail through a specific embodiment as follows. Figure 1 As shown, the specific implementation steps of the mechanical-thermal integration design method of the present invention are as follows:

[0098] Step 1: Design of initial structure; the process is as follows Figure 2 As shown;

[0099] The specific steps for designing the initial structure are as follows:

[0100] Step 1.1, determine the working condition target parameters that the initial structure needs to meet;

[0101] Preferably, the initial structure includes the following loads: one or more antenna loads; one or more carbon fiber loads; one or more thermal control multi-layer loads; one or more camera loads; one or more ground information processing platforms; and one or more batteries.

[0102] Preferably, the target parameters in step 1.1 include: the storage ratio of the initial structure as a whole, the center of gravity coordinates of the initial structure as a whole; the center of gravity coordinates of the antenna load; the center of gravity coordinates of the carbon fiber load; the center of gravity coordinates of the thermal control multi-layer load; the center of gravity coordinates of the camera load; the center of gravity coordinates of the ground information processing platform; and the center of gravity coordinates of the battery.

[0103] The storage ratio of the initial structure is the percentage of the volume of the initial structure in the stored and combined state to the volume of the initial structure in the working and unfolded state;

[0104] The center of gravity coordinates are the coordinate values ​​of the center of gravity in three directions XYZ;

[0105] The expression of the storage ratio of the initial structure is:

[0106]

[0107] Among them, L s is the length of the initial structure in the storage and merging state, W s H is the width of the initial structure in the storage and merging state, s is the height of the initial structure in the storage and merging state, L z is the length of the initial structure in the working state, W z is the width of the initial structure in the working state, H z It is the height of the initial structure in the working state.

[0108] The expression of the barycentric coordinates is:

[0109]

[0110] Where X represents the X coordinate axis, m i represents the mass center of a load i in the initial structure, x i represents the X-axis coordinate of a load i in the initial structure, X m Represents the horizontal coordinate value of the center of mass of the initial structure as a whole in the coordinate system.

[0111]

[0112] Among them, Y represents the Y coordinate axis, m i represents the mass center of a load i in the initial structure, y i represents the Y-axis of a load i in the initial structure, Y mRepresents the vertical coordinate value of the center of mass of the initial structure as a whole in the coordinate system.

[0113]

[0114] Where Z represents the Z coordinate axis, m i represents the mass center of a load i in the initial structure, z i It represents the Z-direction abscissa of a load i in the initial structure, Z m Represents the ordinate value of the center of mass of the initial structure as a whole in the coordinate system.

[0115] Step 1.2, determine the constraint parameters that the initial structure needs to satisfy;

[0116] Preferably, the constraint parameters in step 1.2 include: the deformation of the initial structure, the outer envelope size of the initial structure, and the weight of the initial structure;

[0117] Step 1.3, determine the design variables of the initial structure;

[0118] Preferably, the design variable parameters described in step 1.3 include: connection strength between parts inside the initial structure, connection hole positions inside the initial structure;

[0119] Furthermore, the connection strength value is generally not greater than the allowable stress of the material used; the center coordinates of the connection hole do not exceed the outer envelope of the initial structure; and the diameter value of the connection hole matches the size of the selected bolt connection.

[0120] Step 1.4, the design variable parameters are used as input, multiple groups of different working condition parameters are set as targets, and a solution model is constructed. The solution model uses a multi-objective genetic algorithm and a multi-objective linear programming algorithm to optimize multiple targets until the structure converges and the optimal design variable solution set is obtained;

[0121] Step 1.5, extract the optimal design variable solution set to design the initial structure, and finally obtain the initial structure that meets the requirements;

[0122] Preferably, the optimization objective function expression is:

[0123] minR(Δ)=min{r1=-k s (Δ),r2=-k z (Δ)}

[0124] Among them, R (Δ) is the objective function of the initial structure storage ratio and the initial structure overall center of gravity coordinates, r1 is the initial structure storage ratio, r2 is the initial structure overall center of gravity coordinates, k s represents the optimization function considering the minimum storage ratio condition, k zrepresents the optimization function considering the lowest coordinates of the center of gravity, and Δ is the function of optimizing the design variables.

[0125] Preferably, the expression of the initial structural design variables is:

[0126] Δ=(F(d),N(d))

[0127] Among them, Δ is the function value of the optimized design variable, F(d) is the preload strength provided by the connecting bolts at the load connection, N(d) is the number of connecting holes at the load connection, and d is the numerical size of the hole diameter of the connecting hole.

[0128] Preferably, the constraint conditions are the deformation of the initial structure, the outer envelope size of the initial structure and the total weight of the initial structure, and the expression is:

[0129]

[0130] If p1>0, then k1=0

[0131] If p1≤0, then k1=1

[0132] Among them, p1 is the initial structural deformation coefficient, λ deflection is the target value of initial structural deformation, λ deflection_0 is the value after optimization iteration of the initial structural deformation, k1 is the control coefficient of the penalty term of the initial structural deformation;

[0133]

[0134] If p2>0, then k2=0

[0135] If p2≤0, then k2=1

[0136] Among them, p2 is the initial structure outer envelope size coefficient, λ outsize is the target value of the initial structure outer envelope size, λ outsize_0 is the value after iterative optimization of the initial structure outer envelope size, k2 is the penalty term control coefficient of the initial structure outer envelope size;

[0137]

[0138] If p3>0, then k3=0

[0139] If p3≤0, then k3=1

[0140] Among them, p3 is the initial structure weight limit coefficient, λ weight is the target value of the initial structure weight, λ weight_0 is the value after iterative optimization of the initial structure weight, and k3 is the control coefficient of the initial structure weight penalty term.

[0141] Step 2: Pre-processing the three-dimensional solid model of the initial structure;

[0142] The pre-processing of the 3D solid model of the initial structure includes the following steps:

[0143] Step 2.1, use modeling software to perform geometric processing on the three-dimensional solid model of the initial structure to meet the calculation accuracy, delete small features in the structure, avoid mesh distortion, and improve calculation efficiency;

[0144] Step 2.2, using finite element meshing software to mesh the initial structure three-dimensional solid model, the meshing level is fine, and the meshing type is tetrahedral mesh;

[0145] Step 3: Perform thermal-mechanical coupling simulation calculation on the initial structure based on the CAE simulation model established by the pre-processed initial structure three-dimensional solid model;

[0146] The thermal-mechanical coupling simulation here means that in the simulation process, not only the relevant changes in thermals are considered, but also the mechanical effects such as fixed constraints on the structure and gravity acceleration are considered. The final simulation result is the result after comprehensive consideration of the thermal and mechanical coupling effects.

[0147] The extreme environment described in this embodiment means that the external temperature of the initial structure can be adjusted according to actual needs, and the temperature adjustment area is very large. The extreme environment temperature includes three different types: extremely high temperature, extremely low temperature and extreme high and low temperature sudden change.

[0148] The simulation process includes:

[0149] Step 3.1, simulation under extreme low temperature conditions, the initial structure is in the external environment temperature range of [-196℃, 0℃]. In this example, the extreme low temperature can simulate the low temperature of the space environment. The gas in the space is extremely rare, and non-heating objects cannot actually obtain energy through collisions between gas molecules. They can only rely on absorbing radiation to maintain the temperature level. The radiation energy on the background of the space is extremely small, with a radiation energy of only 10 -5 W / m 2 , which is equivalent to 3K absolute black body radiation. And the geometric size of the spacecraft is negligibly small compared to its distance from the planet or star. From the point of view of heat exchange, the reflection of the satellite radiation by the planet or star can be completely ignored. It can be considered that the spacecraft's own radiation all enters the universe, that is, the space is a black body to the satellite. Usually, many materials in extreme low temperature conditions will become cold and brittle when they are in a low temperature environment for a long time, and the strength will decrease and fall off.

[0150] The simulation under extreme low temperature conditions in step 3.1 includes:

[0151] Determination of external boundary conditions:

[0152] S3.1.1. Specify the initial temperature of the external environment at the initial time point;

[0153] S3.1.2. Specify the target temperature value of the external environment at the target time point;

[0154] S3.1.3. Specify the duration of the initial structural external extreme low temperature;

[0155] Determination of internal boundary conditions:

[0156] S3.1.4. Clarify the thermal control protection measures under extremely low temperatures;

[0157] S3.1.5. Specify the temperature of each load under the electronically controlled active heating protection;

[0158] S3.1.6. Specify the initial structure temperature under the electronically controlled active heating protection;

[0159] S3.1.7. Clarify the mechanical boundary conditions of the initial structure;

[0160] Simulation calculation output results:

[0161] S3.1.8. Simulate and calculate and output the results.

[0162] In this example, the thermal control protection measure under extreme low temperature described in step S3.1.4 is electrically controlled active heating and temperature control. This device converts electrical energy into thermal energy based on the principle of Joule effect, so that the controlled object is maintained within a predetermined temperature range. It is one of the thermal control methods widely used in spacecraft. This device must rely on the spacecraft power supply to provide electrical energy. The electrically controlled active heating and temperature control device generally includes three parts: an electric heater, a controller and a temperature sensor, which constitute a closed-loop control circuit. The electric heater provides heat, and the controller compares the obtained temperature signal with the set value and controls the working state of the heater after making a judgment. The temperature sensor is used to obtain the temperature change signal of the controlled object. On spacecraft, the electrically controlled active heating and temperature control device is mostly used in the following occasions: (1) It can be used to heat the temperature-controlled components on the spacecraft to prevent circuit signal failure, pipeline blockage and other faults caused by the extremely cold external environment, and ensure that the temperature of the components is appropriate before cold start; (2) Keep the temperature of instruments and equipment stable and uniform, and reduce deformation caused by uneven thermal stress; (3) Keep certain components that strictly require constant temperature in a constant temperature environment.

[0163] In this example, the mechanical boundary conditions of the initial structure in step S3.1.7 refer to the determination of the initial structure fixed constraints and gravity boundary conditions; since the initial structure is connected to the turntable in the experimental cabin by bolts, it is necessary to reproduce the bolt-fixed constraint relationship in the simulation environment. And because the experiment is conducted in the earth environment, all objects are affected by gravity acceleration, so it is also necessary to apply the gravity acceleration effect to the initial structure.

[0164] Specifically include:

[0165] (1) Set the bolts between the initial structure and the test cabin turntable;

[0166] (2) Setting bolts to fix the components inside the initial structure;

[0167] (3) Setting the gravitational acceleration of the initial structure;

[0168] In this example, the simulation calculation and output of the results in step S3.1.8 refers to the thermal-mechanical coupling simulation calculation of the initial structure through the thermal-mechanical coupling simulation model, external boundary condition determination and internal boundary condition determination. The thermal-mechanical coupling simulation here means that in the simulation process, not only the relevant changes in thermals are considered, but also the mechanical effects such as the fixed constraints and gravity acceleration of the structure are considered. The final simulation result is the result after comprehensive consideration of the thermal and mechanical coupling effects.

[0169] Step 3.2, simulation under extreme high temperature conditions, the external ambient temperature range of the initial structure is in the range of [0℃, 500℃].

[0170] The simulation under extreme high temperature conditions in step 3.2 includes:

[0171] Determination of external boundary conditions:

[0172] S3.2.1. Specify the initial temperature of the external environment at the initial time point;

[0173] S3.2.2. Specify the target temperature value of the external environment at the target time point;

[0174] S3.2.3. Specify the duration of the initial structural external extreme heat;

[0175] Determination of internal boundary conditions:

[0176] S3.2.4. Clarify the thermal control protection measures under extreme high temperatures;

[0177] S3.2.5. Specify the temperature of each load under thermal insulation protection;

[0178] S3.2.6. Specify the initial structural temperature under sandblasting reflection;

[0179] S3.2.7. Clarify the mechanical boundary conditions of the initial structure;

[0180] Simulation calculation output results:

[0181] S3.2.8. Simulate and calculate and output the results.

[0182] In this example, the thermal control protection measures under extremely high temperatures described in step S3.2.4 are thermal insulation protection and sandblasting reflection protection.

[0183] Thermal insulation protection is that different types of loads are connected to the mounting panel in a way that there is a thermal insulation spacer in the middle. Various types of loads (one or more antenna loads; one or more carbon fiber loads; one or more thermal control multi-layer loads; one or more camera loads; one or more ground information processing platforms; one or more batteries) will be installed on the mounting panel of the initial structure. These loads are made of different materials, and have different heat absorption and reflectivity, and ultimately the temperature changes of different loads are also different. Thermal insulation spacers mostly use loose fiber spacing materials, such as glass fiber in this embodiment. After using this spacing material, the contact heat transfer between the load and the mounting panel needs to be achieved through heat transfer between the panel and the fiber, the fiber itself, the fiber and the fiber, and the fiber and the load, so the contact thermal resistance is large and can effectively insulate.

[0184] Sandblasting reflection protection means that due to the need for protection, the installation panel of the initial structure needs to be heat treated. When the external heat flux is used as a load and loaded into the installation panel, part of the heat will cause the temperature of the structure to rise, and part of it will be reflected through diffuse reflection. This treatment method avoids the initial structure from rising too fast or rising too high. Directly using the surface state of the metal substrate to adjust the thermal radiation properties of the spacecraft surface is the most practical and cost-saving thermal control coating method. For the same metal, their thermal radiation properties may be very different due to different surface states. In this embodiment, the thermal control treatment method of sandblasting the metal surface is used. The metal surface after sandblasting is relatively rough, and usually has a high absorptivity and emissivity. The metal surface after sandblasting is relatively rough and the influence of surface oxidation, mechanical and other damage on the surface roughness is relatively small, so its stability is excellent.

[0185] In this example, the clarification of the mechanical boundary conditions of the initial structure in step S3.2.7 refers to the determination of the initial structure fixed constraints and gravity boundary conditions; since the initial structure is connected to the turntable in the experimental cabin by bolts, it is necessary to reproduce the bolt-fixed constraint relationship in the simulation environment. And because the experiment is conducted in the earth environment, all objects are affected by the acceleration of gravity, so the influence of gravity acceleration also needs to be applied to the initial structure.

[0186] Specifically include:

[0187] (1) Set the bolts between the initial structure and the test cabin turntable;

[0188] (2) Setting bolts to fix the components inside the initial structure;

[0189] (3) Setting the gravitational acceleration of the initial structure;

[0190] In this example, the simulation calculation and output of the results in step S3.2.8 refers to the thermal-mechanical coupling simulation calculation of the initial structure through the thermal-mechanical coupling simulation model, external boundary condition determination and internal boundary condition determination. The thermal-mechanical coupling simulation here means that in the simulation process, not only the relevant changes in thermals are considered, but also the mechanical effects such as the fixed constraints and gravity acceleration of the structure are considered. The final simulation result is the result after comprehensive consideration of the thermal and mechanical coupling effects.

[0191] Step 3.3, extreme high and low temperature sudden change working condition simulation, the external ambient temperature range of the initial structure varies between two different ranges: [-196℃, 0℃] and [0℃, 500℃].

[0192] The simulation under extreme high and low temperature sudden change conditions in step 3.3 includes:

[0193] Determination of external boundary conditions:

[0194] S3.3.1 Specify the minimum external ambient temperature;

[0195] S3.3.2 Specify the maximum temperature of the external environment;

[0196] S3.3.3 Specify the change curve and frequency of sudden high and low temperature changes;

[0197] Determination of internal boundary conditions:

[0198] S3.3.4 Low temperature section, the temperature of each load after starting the electronic control active heating;

[0199] S3.3.5 Low temperature section, initial structure temperature after starting electronically controlled active heating;

[0200] S3.3.6 High temperature section, temperature of each load under thermal insulation protection;

[0201] S3.3.7 High temperature section, initial structural temperature under sandblasting reflection;

[0202] S3.3.8 Specify the mechanical boundary conditions of the initial structure;

[0203] Simulation calculation output results:

[0204] S3.3.9 Perform simulation calculations and output the results.

[0205] In this example, the low temperature protection measures adopted in step S3.3.4 and step S3.3.5 are electronically controlled active heating and temperature adjustment, which are the same as the measures and principles described in step 3.1.4.

[0206] In this example, the high temperature protection measure adopted in step S3.3.6 is thermal insulation protection, which is the same as the thermal insulation protection measure and principle described in step S3.2.4.

[0207] In this example, the high temperature protection measure adopted in step S3.3.7 is sandblasting reflection, which is the same as the sandblasting reflection measure and principle described in step S3.2.4.

[0208] In this example, the clarification of the mechanical boundary conditions of the initial structure in step S3.3.8 refers to the determination of the initial structure fixed constraints and gravity boundary conditions; since the initial structure is connected to the turntable in the experimental cabin by bolts, it is necessary to reproduce the bolt-fixed constraint relationship in the simulation environment. And because the experiment is conducted in the earth environment, all objects are affected by the acceleration of gravity, so it is also necessary to impose the influence of gravity acceleration on the initial structure.

[0209] Specifically include:

[0210] (1) Set the bolts between the initial structure and the test cabin turntable;

[0211] (2) Set the bolts between the components inside the initial structure

[0212] (3) Set the gravitational acceleration of the initial structure.

[0213] In this example, the simulation calculation and output of the results in step S3.3.9 refers to the thermal-mechanical coupling simulation calculation of the initial structure through the thermal-mechanical coupling simulation model, external boundary condition determination and internal boundary condition determination. The thermal-mechanical coupling simulation here means that in the simulation process, not only the relevant changes in thermals are considered, but also the mechanical effects such as the fixed constraints and gravity acceleration of the structure are considered. The final simulation result is the result after comprehensive consideration of the thermal and mechanical coupling effects.

[0214] Step 4: obtain the overall deformation of the initial structure, the overall stress parameter value of the initial structure, and the strength value of the key connection parts through simulation, and analyze and verify whether the design scheme of the initial structure is feasible;

[0215] Step 5: If the design is feasible, the design is directly output; otherwise, the design is adjusted according to the characterization parameters of the overall deformation of the structure, the overall stress parameter value of the structure, and the strength value of the key connection parts, and the initial structure is returned to repeat steps 1 to 4.

[0216] Preferably, the characterization parameter evaluation criteria of the overall structural deformation of the initial structure, the overall structural stress parameter value, and the strength value of the key connection parts include:

[0217] (1) The temperature difference of the overall temperature field of the initial structure should be as small as possible, and the maximum temperature value of the temperature field should be as low as possible;

[0218] (2) The smaller the deformation difference of the initial structure and the overall structural field, the better, and the lower the maximum deformation value of the deformation field, the better;

[0219] (3) The smaller the stress difference of the overall stress field of the initial structure, the better, and the lower the maximum stress value of the stress field, the better;

[0220] (4) The stress intensity value of the key connection parts inside the initial structure should be as small as possible, and the maximum stress intensity value shall not exceed the allowable stress value of the material.

[0221] Preferably, the CAE simulation software for platform design and analysis may be UG NX, HYPERMESH, ANSYS_WORKBENCH, or ABAQUS.

[0222] Figure 6 This is a comparison chart of the efficiency of achieving the design goal using the design method of the present invention and the traditional design method. It can be seen that the design method of the present invention also has a great advantage in efficiency.

[0223] Those skilled in the art will readily appreciate other embodiments of the present application after considering the description and practicing the contents disclosed herein. The present application is intended to cover any modification, use or adaptation of the present application, which follows the general principles of the present application and includes common knowledge or customary techniques in the art that are not disclosed in the present application. The description and examples are intended to be exemplary only, and the true scope and spirit of the present application are indicated by the claims.

[0224] It should be understood that the present application is not limited to the precise structures that have been described above and shown in the drawings, and that various modifications and changes may be made without departing from the scope thereof. The scope of the present application is limited only by the appended claims.

Claims

1. A modular design method for spacecraft mechanical and thermal integration in extreme environments, characterized in that: The method is a modular design method for mechanical and thermal integration that comprehensively considers extreme environments, heat treatment protection, and structural micro-deformation, and includes the following steps: Step 1: Based on the load requirements and the target parameters of the working conditions to be met, the initial structure of the mechanical and thermal integration platform is designed under constraints; Step 2: constructing a three-dimensional solid model based on the initial structure and performing pre-processing on it; Step 3: A CAE simulation model is established based on the initial structural 3D solid model after pre-processing, and a thermal-mechanical coupling simulation calculation is performed on the model. In this process, not only the relevant changes in thermal properties are considered, but also the fixed constraints and mechanical effects on the structure are considered; Step 4: Obtain the overall deformation of the initial structure, the overall stress parameter value of the initial structure, and the strength value of the key connection parts through simulation, and analyze and verify whether the design scheme of the initial structure is feasible; Step 5: If the design is feasible, the design is directly output; otherwise, the design is adjusted according to the characterization parameters of the overall deformation of the structure, the overall stress parameter value of the structure, and the strength value of the key connection parts, and the initial structure design is carried out again in step 1; In step 1, the specific steps of designing the initial structure are as follows: Step 1.1, determine the load of the initial structure and the target parameters of the working condition that need to be met; The initial structure includes the following loads: one or more antenna loads; one or more carbon fiber loads; one or more thermal control multi-layer loads; one or more camera loads; one or more ground information processing platforms; one or more batteries; The working condition target parameters include: the storage ratio of the initial structure as a whole, the center of gravity coordinates of the initial structure as a whole, the center of gravity coordinates of the antenna load, the center of gravity coordinates of the carbon fiber load, the center of gravity coordinates of the thermal control multi-layer load, the center of gravity coordinates of the camera load, the center of gravity coordinates of the ground information processing platform, and the center of gravity coordinates of the battery; Step 1.2, determine the constraint parameters that the initial structure needs to satisfy; The constraint parameters include: the deformation of the initial structure, the outer envelope size of the initial structure, and the weight of the initial structure; Step 1.3, determine the design variables of the initial structure; The design variables include: connection strength between parts inside the initial structure, connection hole positions inside the initial structure; Step 1.4, taking the design variables as input, setting multiple groups of different working condition parameters as targets, constructing a multi-objective optimization solution model under constraints, and using a multi-objective solution algorithm to optimize multiple targets until the structure converges to obtain the optimal design variable solution set; Step 1.5: extract the optimal design variable solution set, design the initial structure, and obtain the initial structure of the platform to be verified.

2. The modular design method for spacecraft mechanical and thermal integration in extreme environments according to claim 1 is characterized in that: The storage ratio refers to the percentage of the volume in the stored and combined state to the volume in the working unfolded state; the center of gravity coordinates are the coordinate values ​​of the center of gravity in three directions of XYZ.

3. The modular design method for mechanical and thermal integration of spacecraft in extreme environments according to claim 1 is characterized in that: When determining the design variables of the initial structure in step 1.3, the following constraints are simultaneously introduced: the connection strength value is not greater than the allowable stress of the material used; the center coordinates of the connection hole position do not exceed the outer envelope of the initial structure; the diameter value of the connection hole matches the size of the selected bolt connection.

4. The modular design method for spacecraft mechanical and thermal integration in extreme environments according to claim 1 is characterized in that: In step 1.4, the optimization objective function expression constructed is: minR(Δ)=min{r1=-k s (Δ),r2=-k z (Δ)} Among them, R (Δ) is the objective function of the initial structure storage ratio and the initial structure overall center of gravity coordinates, r1 is the initial structure storage ratio, r2 is the initial structure overall center of gravity coordinates, k s represents the optimization function considering the minimum storage ratio condition, k z represents the optimization function considering the lowest coordinates of the center of gravity, and Δ is the function of the optimized design variables; The expression of the initial structural design variable is: Δ=(F(d),N(d)) Among them, Δ is the function value of the optimized design variable, F(d) is the preload strength provided by the connecting bolts at the load connection, N(d) is the number of connecting holes at the load connection, and d is the numerical value of the hole diameter of the connecting hole; The constraints are the deformation of the initial structure, the outer envelope size of the initial structure, and the total weight of the initial structure, expressed as: If p1>0, then k1=0 If p1≤0, then k1=1 Among them, p1 is the initial structural deformation coefficient, λ deflection is the target value of initial structural deformation, λ deflection_0 is the value after optimization iteration of the initial structural deformation, k1 is the control coefficient of the penalty term of the initial structural deformation; If p2>0, then k2=0 If p2≤0, then k2=1 Among them, p2 is the initial structure outer envelope size coefficient, λ outsize is the target value of the initial structure outer envelope size, λ outsize_0 is the value after iterative optimization of the initial structure outer envelope size, k2 is the penalty term control coefficient of the initial structure outer envelope size; If p3>0, then k3=0 If p3≤0, then k3=1 Among them, p3 is the initial structure weight limit coefficient, λ weight is the target value of initial structure weight, λ weight_0 is the value after iterative optimization of the initial structure weight, and k3 is the control coefficient of the initial structure weight penalty term.

5. The modular design method for mechanical and thermal integration of spacecraft in extreme environments according to claim 1 is characterized in that: The pre-processing of the 3D solid model of the initial structure includes the following steps: Step 2.1, use modeling software to perform geometric processing on the three-dimensional solid model of the initial structure to meet the calculation accuracy, delete small features in the structure, and avoid mesh distortion; Step 2.2, use the finite element meshing software to mesh the initial structure three-dimensional solid model, the meshing level is fine, and the meshing type is tetrahedral mesh.

6. The modular design method for mechanical and thermal integration of spacecraft in extreme environments according to claim 1 is characterized in that: A CAE simulation model is established based on the initial structural three-dimensional solid model after pre-processing, and the model is subjected to thermal-mechanical coupling simulation calculation, including the following: The external temperature of the initial structure can be adjusted according to actual needs. The extreme environmental temperature includes three different types: extremely high temperature, extremely low temperature and extreme high and low temperature sudden change. CAE simulation model is established and simulation calculation is performed for each type of working condition. The extremely high temperature means that the external environmental temperature is in the range of [0℃, 500℃], the extremely low temperature means that the external environmental temperature is in the range of [-196℃, 0℃], and the extreme high and low temperature sudden change means that the external environmental temperature changes between the extremely high temperature range and the extremely low temperature range.

7. The modular design method for mechanical and thermal integration of spacecraft in extreme environments according to claim 6 is characterized in that: The establishment and simulation of CAE simulation model under extreme low temperature conditions include the following: Determination of external boundary conditions: S3.1.

1. Specify the initial temperature of the external environment at the initial time point; S3.1.

2. Specify the target temperature value of the external environment at the target time point; S3.1.

3. Specify the duration of the initial structural external extreme low temperature; Determination of internal boundary conditions: S3.1.

4. Specify that the thermal protection measure under extremely low temperature is electronically controlled active heating and temperature regulation; S3.1.

5. Specify the temperature of each load under the electronically controlled active heating protection; S3.1.

6. Specify the initial structure temperature under the electronically controlled active heating protection; S3.1.

7. Clarify the mechanical boundary conditions of the initial structure, including the initial structure fixed constraints and gravity boundary conditions; S3.1.

8. Simulate and calculate and output the results.

8. The modular design method for mechanical and thermal integration of spacecraft in extreme environments according to claim 6 is characterized in that: The establishment and simulation of CAE simulation model under extreme high temperature conditions include the following: Determination of external boundary conditions: S3.2.

1. Specify the initial temperature of the external environment at the initial time point; S3.2.

2. Specify the target temperature value of the external environment at the target time point; S3.2.

3. Specify the duration of the initial structural external extreme heat; Determination of internal boundary conditions: S3.2.

4. Specify that the thermal control protection measures under extreme high temperatures are thermal insulation protection and sandblasting reflection protection; S3.2.

5. Specify the temperature of each load under thermal insulation protection; S3.2.

6. Specify the initial structural temperature under sandblasting reflection; S3.2.

7. Clarify the mechanical boundary conditions of the initial structure, including the initial structure fixed constraints and gravity boundary conditions; S3.2.

8. Simulate and calculate and output the results.

9. The modular design method for mechanical and thermal integration of spacecraft in extreme environments according to claim 6, characterized in that: The establishment and simulation of CAE simulation model under extreme high and low temperature sudden change conditions include the following: Determination of external boundary conditions: S3.3.1 Specify the minimum external ambient temperature; S3.3.2 Specify the maximum temperature of the external environment; S3.3.3 Specify the change curve and frequency of sudden high and low temperature changes; Determination of internal boundary conditions: S3.3.4 Low temperature section, the temperature of each load after starting the electronic control active heating; S3.3.5 Low temperature section, initial structure temperature after starting electronically controlled active heating; S3.3.6 High temperature section, temperature of each load under thermal insulation protection; S3.3.7 High temperature section, initial structural temperature under sandblasting reflection; S3.3.8 Define the mechanical boundary conditions of the initial structure, including the initial structure fixed constraints and gravity boundary conditions; S3.3.9 Perform simulation calculations and output the results.

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