A method for designing the layout of modular units of reconfigurable spacecraft

Abstract

The present invention discloses a method for designing the configuration layout of a reconfigurable spacecraft module unit, including: step 1, determining the system index requirements for the reconfigurable spacecraft module unit; step 2, designing the structural type specifications of the reconfigurable spacecraft module unit according to the determined system index requirements; step 3, designing the type specifications of the standard docking interface according to the determined system index requirements; step 4, designing the installation position of the internal load on the module structure body according to the determined system index requirements; step 5, checking whether the design results of steps 2 to 4 meet the system index requirements and task requirements, if not, repeating steps 2 to 4, and iteratively optimizing the designs of steps 2 to 4. The present invention can quickly complete the configuration layout design of the module unit according to the index requirements of the reconfigurable spacecraft module unit, including the module structure type specifications, standard docking interface type specifications, module unit layout mode and other contents.

Description

Technical Field

[0001] The invention belongs to the field of aerospace technology, and in particular relates to a method for designing the layout of a reconfigurable spacecraft module unit configuration. Background Art

[0002] With the expansion of human demand for space utilization and the increase in demand for satellite development, the one-time launch mode of traditional satellites can no longer meet the rapid development needs of space technology. Reconfigurable spacecraft has gradually become an important direction for the development of future space systems.

[0003] Reconfigurable spacecraft adopts modular design and consists of multiple serialized module units. The module units are connected through standard docking interfaces to facilitate on-orbit maintenance and upgrades. The design of module units needs to comprehensively consider multiple constraints such as system functional requirements and performance indicators. The quality of their design status directly affects the functional performance of the entire spacecraft.

[0004] At present, modular reconfigurable spacecraft projects in various countries around the world have not formed a systematic module unit configuration layout design method, nor have they fully covered the mechanical, electrical, information, and gas-liquid connection functions between spacecraft modules, and cannot meet the design requirements of reconfigurable spacecraft module units. Summary of the invention

[0005] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art and provide a method for designing the configuration layout of a reconfigurable spacecraft module unit, which can quickly complete the configuration layout design of the module unit according to the index requirements of the reconfigurable spacecraft module unit, including module structure type specifications, standard docking interface type specifications, module unit layout methods and other contents.

[0006] In order to solve the above technical problems, the present invention discloses a method for designing a configuration layout of a reconfigurable spacecraft module unit, comprising:

[0007] Step 1, determining system indicator requirements for a reconfigurable spacecraft module unit; wherein the reconfigurable spacecraft module unit includes: a module structure body, a standard docking interface installed on the module structure body, and an internal load arranged inside the module structure body;

[0008] Step 2: Design the structural type and specifications of the reconfigurable spacecraft module unit according to the determined system indicator requirements;

[0009] Step 3: Design the type and specifications of the standard docking interface according to the determined system indicator requirements;

[0010] Step 4: Design the installation position of the internal load on the module structure body according to the determined system indicator requirements;

[0011] Step 5, check whether the design results of steps 2 to 4 meet the system indicator requirements and task requirements. If not, repeat steps 2 to 4 and iteratively optimize the designs of steps 2 to 4 until the design results of steps 2 to 4 meet the system indicator requirements and task requirements.

[0012] In the above-mentioned reconfigurable spacecraft module unit configuration layout design method, the system indicator requirements include: the connection relationship requirements between reconfigurable spacecraft module units, the connection force requirements F0 between reconfigurable spacecraft module units, the power requirements P0 and voltage requirements V0 of the electric energy transmission between reconfigurable spacecraft module units, the transmission rate requirements S0 between reconfigurable spacecraft module units, the gas-liquid transmission requirements between reconfigurable spacecraft module units, the internal load mass requirements m of the reconfigurable spacecraft module units and the installation space requirements L required for the internal load zx ×L zy ×L zz , Maximum dimensions required for reconfigurable spacecraft module unit L wx ×L wy ×L wz And the mechanical environment adaptability requirements of reconfigurable spacecraft module units.

[0013] In the above-mentioned reconfigurable spacecraft module unit configuration layout design method, the structural type specification of the reconfigurable spacecraft module unit is designed according to the determined system indicator requirements, including:

[0014] According to the connection relationship requirements between the reconfigurable spacecraft module units, determine the installation location, quantity, position and interface status of the standard docking interface on the module structure body;

[0015] According to the connection force requirement F0 between reconfigurable spacecraft module units and the installation space requirement L required for the internal load of the reconfigurable spacecraft module unit zx ×L zy ×L zz and the maximum dimension requirement L of the reconfigurable spacecraft module unit wx ×L wy ×L wz , determine the actual dimensions of the module structure body;

[0016] According to the internal load mass requirement m of the reconfigurable spacecraft module unit and the mechanical environment adaptability requirement of the reconfigurable spacecraft module unit, the specific type of the module structure body is determined through digital model design and simulation analysis: connection form, structural plate specifications.

[0017] In the above-mentioned reconfigurable spacecraft module unit configuration layout design method, the module structure body is a hexahedral structure, including: +X surface, -X surface, +Y surface, -Y surface, +Z surface and -Z surface; wherein, the interface states of the standard docking interfaces on the +X surface, -X surface, +Y surface, -Y surface, +Z surface and -Z surface are recorded as: e +X 、e -X 、e +Y 、e -Y 、e +Z and e -Z , then according to the connection relationship requirements between the reconfigurable spacecraft module units, we have:

[0018] If the +X surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e +X =1; otherwise, e +X =0;

[0019] If the -X surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e -X =1; otherwise, e -X =0;

[0020] If the +Y surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e +Y =1; otherwise, e +Y =0;

[0021] If the -Y surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e -Y =1; otherwise, e -Y =0;

[0022] If the +Z surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e +Z =1; otherwise, e +Z =0;

[0023] If the -Z surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e -Z =1; otherwise, e -Z =0;

[0024] Furthermore, the total number of standard docking interfaces on the module structure of the current reconfigurable spacecraft module unit is e a :

[0025] e a =e +X +e -X +e+Y +e -Y +e +Z +e -Z 。

[0026] In the above method for designing the configuration layout of the reconfigurable spacecraft module unit, according to the connection force requirement F0 between the reconfigurable spacecraft module units, the installation space requirement L zx ×L zy ×L zz required for the internal payload of the reconfigurable spacecraft module unit, and the maximum outer dimension requirement L wx ×L wy ×L wz , determine the actual outer dimensions of the module structure body, including:

[0027] According to F0, determine the thickness L of the standard docking interface on the module structure body k ;

[0028] According to the installation space requirement L zx ×L zy ×L zz required for the internal payload of the reconfigurable spacecraft module unit, combined with L k , e +X , e -X , e +Y , e -Y , e +Z and e -Z , perform iterative calculations until the actual outer dimensions L wx ×L wy ×L wz of the module structure body that meet the maximum outer dimension requirement L x ×L y ×L z are obtained.

[0029] In the above method for designing the configuration layout of the reconfigurable spacecraft module unit, if 0N < F0 ≤ 200N, then L k = 30mm; if 200N < F0 ≤ 1000N, then L k = 60mm; if 1000N < F0 ≤ 5000N, then L k = 120mm.

[0030] In the above method for designing the configuration layout of the reconfigurable spacecraft module unit, according to the installation space requirement L zx ×L zy ×L zz , combined with L k , e +X , e -X , e+Y 、e -Y 、e +Z and e -Z , perform iterative calculations until the maximum dimension requirement L of the reconfigurable spacecraft module unit is obtained. wx ×L wy ×L wz The actual dimensions of the module structure body are L x ×L y ×L z ,include:

[0031] Determine L x ×L y ×L z The solution formula is:

[0032] L x =L zx +L k ·(e +X +e -X )+L b

[0033] L y =L zy +L k ·(e +Y +e -Y )+L b

[0034] L z =L zz +L k ·(e +Z +e -Z )+L b

[0035] Among them, L b It is a value related to the thickness of the structural plate. For preliminary calculation, L is taken b =50mm, and then make iterative adjustments according to the situation; if L x ≤L wx And L y ≤L wy And L z ≤L wz , then for the actual size L of the module structure x ×L y ×L z If the design is established, continue with the subsequent design steps; otherwise, for the actual size L of the module structure body x ×L y ×L z If the design is not feasible, communicate with the user to determine whether to modify the requirements.

[0036] In the above-mentioned reconfigurable spacecraft module unit configuration layout design method, the standard docking interface includes at least one of the following components: a mechanical connection component, an electric energy transmission component, an information transmission component, and a gas-liquid transmission component.

[0037] In the above-mentioned reconfigurable spacecraft module unit configuration layout design method, the type and specifications of the standard docking interface are designed according to the determined system indicator requirements, including:

[0038] Determine the specification parameters of mechanical connection components according to the connection relationship requirements between reconfigurable spacecraft module units;

[0039] Determine the specification parameters of the power transmission components according to the power requirement P0 and voltage requirement V0 of the power transmission between the reconfigurable spacecraft module units;

[0040] Determine the specification parameters of the information transmission component according to the transmission rate requirement S0 between the reconfigurable spacecraft module units;

[0041] Determine the specification parameters of gas-liquid transmission components according to the gas-liquid transmission requirements between reconfigurable spacecraft module units;

[0042] Determine the layout of the internal components of the standard docking interface, combine the above-mentioned mechanical connection components, power transmission components, information transmission components and gas-liquid transmission components, form a variety of feasible internal type combinations and layout methods of the standard docking interface, and select the design state with the smallest volume to carry out subsequent work.

[0043] In the above-mentioned reconfigurable spacecraft module unit configuration layout design method, the installation position of the internal load on the module structure body is designed according to the determined system indicator requirements, including: determining the installation position of each internal load on the module structure body according to the mass requirements and the required installation space requirements of each internal load of the reconfigurable spacecraft module unit; wherein, the internal load with a mass not greater than the preset mass threshold is installed on the top plate of the module structure body, and the internal load with a mass greater than the preset mass threshold is installed on the bottom plate of the module structure body.

[0044] The present invention has the following advantages:

[0045] (1) The present invention discloses a method for designing the configuration layout of a reconfigurable spacecraft module unit. The method can be used to quickly determine the type and specifications of the module structure body, the type and specifications of the standard docking interface, the layout of the reconfigurable spacecraft module unit, and other design conditions.

[0046] (2) The present invention discloses a method for designing the layout of a modular unit configuration of a reconfigurable spacecraft, which completely covers the mechanical, electrical, information, and gas-liquid connection functional designs between the modular units of the spacecraft, and can effectively support the functional realization of a modular reconfigurable spacecraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0047] Figure 1 It is a flowchart of the steps of a method for designing a configuration layout of a module unit of a reconfigurable spacecraft in an embodiment of the present invention;

[0048] Figure 2 is a configuration example diagram of a reconfigurable spacecraft module unit in an embodiment of the present invention;

[0049] Figure 3 It is a schematic diagram of the combined assembly of multiple reconfigurable spacecraft module units in an embodiment of the present invention;

[0050] Figure 4 This is an example diagram of a standard docking interface internal layout mode 1 in an embodiment of the present invention;

[0051] Figure 5 This is an example diagram of a standard docking interface internal layout mode 2 in an embodiment of the present invention;

[0052] Figure 6 This is an example diagram of a standard docking interface internal layout method 3 in an embodiment of the present invention. DETAILED DESCRIPTION

[0053] In order to make the objectives, technical solutions and advantages of the present invention more clear, the embodiments disclosed in the present invention will be further described in detail below with reference to the accompanying drawings.

[0054] In response to the design problem of reconfigurable spacecraft module units, the present invention discloses a method for designing the configuration layout of reconfigurable spacecraft module units. According to the index requirements of the reconfigurable spacecraft module units, the design states such as the type and specifications of the module structure body, the type and specifications of the standard docking interface, and the layout of the module units are determined, and optimized through multiple rounds of iterations.

[0055] like Figure 1 In this embodiment, the reconfigurable spacecraft module unit configuration layout design method includes:

[0056] Step 1: Determine the system indicator requirements for reconfigurable spacecraft module units.

[0057] In this embodiment, if Figures 2-3 The reconfigurable spacecraft module unit may specifically include: a module structure body 100, a standard docking interface installed on the module structure body, and an internal load arranged inside the module structure body. There may be multiple standard docking interfaces, such as Figure 2 The standard docking interface A201, standard docking interface B202, etc. are shown; the internal load can also be multiple, such as Figure 2The internal load A301, internal load B302, etc. are shown. It should be noted that, according to different mission requirements, the types and quantities of standard docking interfaces and the types and quantities of internal loads in different reconfigurable spacecraft module units are different and can be determined according to actual conditions.

[0058] Preferably, the system indicator requirements mainly include:

[0059] 1) The connection relationship requirements between reconfigurable spacecraft module units, that is, to clarify which surface of a reconfigurable spacecraft module unit needs to be connected to other reconfigurable spacecraft module units;

[0060] 2) Connection force requirement F0 between reconfigurable spacecraft module units;

[0061] 3) Power transmission demand P0 and voltage demand V0 between reconfigurable spacecraft module units;

[0062] 4) Transmission rate requirement S0 between reconfigurable spacecraft module units;

[0063] 5) Requirements for gas and liquid transmission between reconfigurable spacecraft modules;

[0064] 6) The mass requirement m of the internal load of the reconfigurable spacecraft module unit and the installation space requirement L of the internal load zx ×L zy ×L zz , that is, X-direction dimension × Y-direction dimension × Z-direction dimension;

[0065] 7) Maximum dimensions of reconfigurable spacecraft module unit L wx ×L wy ×L wz , that is, the maximum size in X direction × the maximum size in Y direction × the maximum size in Z direction;

[0066] 8) Mechanical environment adaptability requirements of reconfigurable spacecraft module units.

[0067] Step 2: Design the structural type and specifications of the reconfigurable spacecraft module unit according to the determined system indicator requirements.

[0068] In this embodiment, the specific implementation of step 2 can be as follows:

[0069] (21) According to the connection relationship requirements between the reconfigurable spacecraft module units, determine the installation location, quantity, position and interface status of the standard docking interface on the module structure body.

[0070] In this embodiment, if Figure 2As shown, the module structure body is generally a hexahedral structure, specifically including six faces: +X face, -X face, +Y face, -Y face, +Z face and -Z face. Among them, the interface states of the standard docking interfaces on the +X face, -X face, +Y face, -Y face, +Z face and -Z face are recorded as follows: e +X 、e -X 、e +Y 、e -Y 、e +Z and e -Z , then according to the connection relationship requirements between reconfigurable spacecraft module units, we have: 1) If the +X surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e +X =1; otherwise, e +X = 0; 2) If the -X surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e -X =1; otherwise, e -X = 0; 3) If the +Y surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e +Y =1; otherwise, e +Y =0; 4) If the -Y surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e -Y =1; otherwise, e -Y = 0; 5) If the +Z surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e +Z =1; otherwise, e +Z = 0; 6) If the -Z surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e -Z =1; otherwise, e -Z = 0. Furthermore, the total number of standard docking interfaces on the module structure body of the current reconfigurable spacecraft module unit is e a :e a =e +X +e -X +e +Y +e -Y +e +Z +e -Z .

[0071] (22) According to the connection force requirement F0 between reconfigurable spacecraft module units and the installation space requirement L for the internal load of the reconfigurable spacecraft module unit zx ×L zy ×L zz and the maximum dimension requirement L of the reconfigurable spacecraft module unit wx ×Lwy ×L wz , determine the actual external dimensions of the module structure body.

[0072] In this embodiment, considering the thickness L of the standard docking interface k is related to the connection force requirement F0 between the reconfigurable spacecraft module units, the thickness L of the standard docking interface on the module structure body can be determined according to F0 k . For example, in a typical example: if 0N < F0 ≤ 200N, then L k can be taken as 30mm; if 200N < F0 ≤ 1000N, then L k can be taken as 60mm; if 1000N < F0 ≤ 5000N, then L k can be taken as 120mm.

[0073] Furthermore, according to the installation space requirement L zx ×L zy ×L zz , combined with L k , e +X , e -X , e +Y , e -Y , e +Z and e -Z , perform iterative calculations until the actual external dimensions L wx ×L wy ×L wz of the module structure body that meet the maximum external dimension requirement L x ×L y ×L z are obtained.

[0074] The solution formula for L x ×L y ×L z is as follows:

[0075] L x = L zx + L k ·(e +X + e -X ) + L b

[0076] L y = L zy + L k ·(e +Y + e -Y ) + L b

[0077] L z = L zz + Lk ·(e +Z +e -Z )+L b

[0078] Among them, L b It is a value related to the thickness of the structural plate. For preliminary calculation, L is taken b =50mm, and then make iterative adjustments according to the situation; if L x ≤L wx And L y ≤L wy And L z ≤L wz , then for the actual size L of the module structure x ×L y ×L z If the design is established, continue with the subsequent design steps; otherwise, for the actual size L of the module structure body x ×L y ×L z If the design is not feasible, communicate with the user to determine whether to modify the requirements.

[0079] (23) Based on the internal load mass requirement m of the reconfigurable spacecraft module unit and the mechanical environment adaptability requirement of the reconfigurable spacecraft module unit, the specific type of the module structure body (such as connection form, structural plate specifications, etc.) is determined through digital model design and simulation analysis.

[0080] Step 3: Design the type and specifications of the standard docking interface based on the determined system indicator requirements.

[0081] In this embodiment, the standard docking interface includes at least one of the following components: a mechanical connection component, an electric energy transmission component, an information transmission component, and a gas-liquid transmission component. During the design, the specification parameters of the mechanical connection component can be determined according to the connection relationship requirements between the reconfigurable spacecraft module units; the specification parameters of the electric energy transmission component can be determined according to the power requirements P0 and voltage requirements V0 of the electric energy transmission between the reconfigurable spacecraft module units; the specification parameters of the information transmission component can be determined according to the transmission rate requirements S0 between the reconfigurable spacecraft module units; the specification parameters of the gas-liquid transmission component can be determined according to the gas-liquid transmission requirements between the reconfigurable spacecraft module units; then, the layout of the components inside the standard docking interface is determined, and the above-mentioned mechanical connection components, electric energy transmission components, information transmission components, and gas-liquid transmission components are combined to form a variety of feasible standard docking interface internal type combinations and layout methods, and the design state with the smallest volume is selected to carry out subsequent work.

[0082] Preferably, the specification parameters of a typical optional mechanical connection component include but are not limited to the following:

[0083] Specification shape In-plane dimensions thickness Connectivity JH-200 Ring Φ60mm×Φ45mm 30mm (0～200]N JZ-200 Cylindrical Φ50mm 30mm (0～200]N JH-1000 Ring Φ120mm×Φ90mm 60mm (0～1000]N JZ-1000 Cylindrical Φ100mm 60mm (0～1000]N JH-5000 Ring Φ240mm×Φ180mm 120mm (0～5000]N JZ-5000 Cylindrical Φ200mm 120mm (0～5000]N

[0084] Table 1, Example of mechanical connection parts specification parameters

[0085] Preferably, the specification parameters of a typical optional power transmission component include but are not limited to the following:

[0086] Specification shape In-plane dimensions thickness power Voltage DZ-30 Cylindrical Φ30mm 40mm 800W 100V DZ-50 Cylindrical Φ50mm 40mm 1600W 100V

[0087] Table 2. Example of specification parameters of power transmission components

[0088] Preferably, the specification parameters of a typical optional information transmission component include but are not limited to the following:

[0089] Specification shape In-plane dimensions thickness Information transmission rate XZ-40 Cylindrical Φ40mm 50mm 500Mbps XZ-80 Cylindrical Φ80mm 50mm 1Gbps

[0090] Table 3, Example of specification parameters of information transmission components

[0091] Preferably, the specification parameters of a typical optional gas-liquid transmission component include but are not limited to the following:

[0092] Specification shape In-plane dimensions thickness Working fluid GQ-40 Cylindrical Φ50mm 60mm Gas, liquid GY-40 Cylindrical Φ40mm 50mm liquid

[0093] Table 4, Example of Specifications of Gas and Liquid Transmission Components

[0094] Step 4: Design the installation position of the internal load on the module structure body according to the determined system indicator requirements.

[0095] In this embodiment, the installation position of the internal load on the module structure body can be designed according to the determined system indicator requirements, including: determining the installation position of each internal load on the module structure body according to the mass requirements and the required installation space requirements of each internal load of the reconfigurable spacecraft module unit. That is, the installation position of each internal load on the module structure body is jointly determined by the "mass requirement" and the "installation space requirement". For example, an internal load with a smaller mass (not greater than a preset mass threshold) can be preferentially installed on the top plate of the module structure body, and an internal load with a larger mass (greater than a preset mass threshold) can be preferentially installed on the bottom plate of the module structure body. Of course, when the "mass requirement" conflicts with the "installation space requirement", the internal load with a smaller mass can also be installed on the bottom plate of the module structure body, and the internal load with a larger mass can also be installed on the top plate of the module structure body, and this embodiment does not limit this.

[0096] Step 5, check whether the design results of steps 2 to 4 meet the system indicator requirements and task requirements. If not, repeat steps 2 to 4 and iteratively optimize the designs of steps 2 to 4 until the design results of steps 2 to 4 meet the system indicator requirements and task requirements.

[0097] Based on the above embodiment, a specific example is described below.

[0098] 1. Determine system indicator requirements.

[0099] In a certain task, there are the following system indicator requirements:

[0100] 1) The +X surface, -X surface, and +Y surface of the reconfigurable spacecraft module unit need to be connected to the other module units;

[0101] 2) F0 is 1000N;

[0102] 3) P0 is 6000W, V0 is 100V;

[0103] 4) S0 is 1 Gbps;

[0104] 5) The gas-liquid transmission requirement is 2-way liquid working fluid;

[0105] 6) m is 30 kg, L zx ×L zy ×L zz 200mm×200mm×200mm;

[0106] 7) L wx ×L wy ×L wz 600mm×600mm×600mm;

[0107] 8) Mechanical environment adaptability requirements of reconfigurable spacecraft module units.

[0108] 2. Determine the structural type and specifications of the reconfigurable spacecraft module units.

[0109] As mentioned above, the +X surface, -X surface, and +Y surface of the reconfigurable spacecraft module unit need to be connected with the other module units, so: +X =1, e -X =1, e +Y =1, e -Y =0, e +Z =0, e -Z =0,e a =3.

[0110] Furthermore, if F0 is 1000N, then L k =60mm. Substituting into the formula, we can calculate:

[0111] L x =L zx +L k ·(e +X +e -X )+Lb =370mm

[0112] L y =L zy +L k ·(e +Y +e -Y )+L b =310mm

[0113] L z =L zz +L k ·(e +Z +e -Z )+L b =250mm

[0114] Among them, L b The initial value of is 50mm. wx ×L wy ×L wz is 600mm×600mm×600mm, so “L x ≤L wx And L y ≤L wy And L z ≤L wz "If the condition is met, the design is established and the subsequent design steps can be continued.

[0115] Furthermore, for the outer dimension L x ×L y ×L z =360mm×300mm×240mm module structure body, through digital model design, simulation analysis and other means, it is determined that the module structure body connection form is box panel installation form, and the structural plate specification is 25mm thick aluminum honeycomb panel.

[0116] 3. Determine the type and specifications of the standard docking interface.

[0117] For the connection force requirement F0 of 1000N, the models and quantities of the mechanical connection components initially selected are as follows:

[0118] Specification size quantity JH-1000 Φ120mm×Φ90mm×60mm 1 JZ-1000 Φ100mm×60mm 1

[0119] Table 5, Mechanical connection parts model and quantity example table

[0120] For the requirements of P0 of 6000W and V0 of 100V, the models and quantities of the initially selected power transmission components are as follows:

[0121] Specification size quantity DZ-30 Φ30mm×40mm 8 DZ-50 Φ50mm×40mm 4

[0122] Table 6. Example of models and quantities of power transmission components

[0123] For the requirement of S0 of 1Gbps, the models and quantities of the information transmission components initially selected are as follows:

[0124] Specification size quantity XZ-40 Φ40mm×50mm 2 XZ-80 Φ80mm×50mm 1

[0125] Table 7, Information transmission component model and quantity example table

[0126] For the demand of 2-way liquid working fluid, the models and quantities of gas-liquid transmission components initially selected are as follows:

[0127] Specification size quantity GY-40 Φ40mm×50mm 2

[0128] Table 8, Example of gas and liquid transmission component models and quantities

[0129] Further, determine the layout of the components inside the standard docking interface, combine the above components, form a variety of feasible standard docking interface internal component type combination and layout methods, and select the design state with the smallest volume to carry out subsequent work. For the tasks mentioned above, the optional combination methods are as follows: Table 9, Figures 4 to 6 As shown, method 3 can be preferred to carry out subsequent work.

[0130] part Method 1 Method 2 Method 3 Mechanical connection parts JH-1000 JZ-1000 JZ-1000 Power transmission components DZ-30 DZ-50 DZ-30 Information transmission components XZ-80 XZ-40 XZ-40 Gas and liquid transmission components GY-40 GY-40 GY-40 Overall outer diameter Φ220mm Φ220mm Φ210mm

[0131] Table 9, examples of optional combinations of internal components of standard docking interfaces

[0132] 4. Determine the module unit layout and optimize the overall iteration.

[0133] According to the mass characteristics of each internal load of the module unit and the installation space requirements, determine the installation position of each internal load on the module unit, for example, install the internal load with smaller mass on the top plate of the module unit, and install the load with larger mass on the bottom plate of the module unit. Further, check whether the design status can meet the task requirements. If not, repeat the above steps and perform iterative optimization by adjusting the size, changing the specifications of each component, etc. until the task requirements are met.

[0134] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

[0135] The contents not described in detail in the specification of the present invention belong to the common knowledge of the professionals in this field.

Claims

1. A method for designing the layout of a module unit of a reconfigurable spacecraft, characterized in that: include: Step 1, determining system indicator requirements for a reconfigurable spacecraft module unit; wherein the reconfigurable spacecraft module unit includes: a module structure body, a standard docking interface installed on the module structure body, and an internal load arranged inside the module structure body; Step 2: Design the structural type and specifications of the reconfigurable spacecraft module unit according to the determined system indicator requirements; Step 3: Design the type and specifications of the standard docking interface according to the determined system indicator requirements; Step 4: Design the installation position of the internal load on the module structure body according to the determined system indicator requirements; Step 5, check whether the design results of steps 2 to 4 meet the system indicator requirements and task requirements. If not, repeat steps 2 to 4 and iteratively optimize the designs of steps 2 to 4 until the design results of steps 2 to 4 meet the system indicator requirements and task requirements; System indicator requirements, including: connection relationship requirements between reconfigurable spacecraft module units, connection force requirements F0 between reconfigurable spacecraft module units, power transmission power requirements P0 and voltage requirements V0 between reconfigurable spacecraft module units, transmission rate requirements S0 between reconfigurable spacecraft module units, gas-liquid transmission requirements between reconfigurable spacecraft module units, internal load mass requirements m of reconfigurable spacecraft module units and installation space requirements L for internal loads zx ×L zy ×L zz , Maximum dimensions required for reconfigurable spacecraft module unit L wx ×L wy ×L wz and the mechanical environment adaptability requirements of reconfigurable spacecraft module units; According to the determined system index requirements, the structural type specifications of the reconfigurable spacecraft module unit are designed, including: According to the connection relationship requirements between the reconfigurable spacecraft module units, determine the installation location, quantity, position and interface status of the standard docking interface on the module structure body; According to the connection force requirement F0 between reconfigurable spacecraft module units and the installation space requirement L required for the internal load of the reconfigurable spacecraft module unit zx ×L zy ×L zz and the maximum dimension requirement L of the reconfigurable spacecraft module unit wx ×L wy ×L wz , determine the actual dimensions of the module structure body; According to the internal load mass requirement m of the reconfigurable spacecraft module unit and the adaptability requirement of the mechanical environment of the reconfigurable spacecraft module unit, the specific type of the module structure body is determined through digital model design and simulation analysis: connection form, structural plate specifications; The module structure body is a hexahedral structure, including: +X face, -X face, +Y face, -Y face, +Z face and -Z face; among them, the interface states of the standard docking interfaces on the +X face, -X face, +Y face, -Y face, +Z face and -Z face are recorded as follows: e +X 、e -X 、e +Y 、e -Y 、e +Z and e -Z , then according to the connection relationship requirements between the reconfigurable spacecraft module units, we have: If the +X surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e +X =1; otherwise, e +X =0; If the -X surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e -X =1; otherwise, e -X =0; If the +Y surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e +Y =1; otherwise, e +Y =0; If the -Y surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e -Y =1; otherwise, e -Y =0; If the +Z surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e +Z =1; otherwise, e +Z =0; If the -Z surface of the module structure body of the current reconfigurable spacecraft module unit is connected to other reconfigurable spacecraft module units, then e -Z =1; otherwise, e -Z =0; Furthermore, the total number of standard docking interfaces on the module structure of the current reconfigurable spacecraft module unit is e a : And a =and +X +e -X +e +Y +e -Y +e +Z +e -Z ; According to the connection force requirement F0 between reconfigurable spacecraft module units and the installation space requirement L required for the internal load of the reconfigurable spacecraft module unit zx ×L zy ×L zz and the maximum dimension requirement L of the reconfigurable spacecraft module unit wx ×L wy ×L wz , determine the actual dimensions of the module structure, including: According to F0, determine the thickness L of the standard docking interface on the module structure body k ; According to the installation space requirement L of the internal load of the reconfigurable spacecraft module unit zx ×L zy ×L zz , combined with L k 、e +X 、e -X 、e +Y 、e -Y 、e +Z and e -Z , perform iterative calculations until the maximum dimension requirement L of the reconfigurable spacecraft module unit is obtained. wx ×L wy ×L wz The actual dimensions of the module structure body are L x ×L y ×L z ; According to the installation space requirement L of the internal load of the reconfigurable spacecraft module unit zx ×L zy ×L zz , combined with L k 、e +X 、e -X 、e +Y 、e -Y 、e +Z and e -Z , perform iterative calculations until the maximum dimension requirement L of the reconfigurable spacecraft module unit is obtained. wx ×L wy ×L wz The actual dimensions of the module structure body are L x ×L y ×L z ,include: Determine L x ×L y ×L z The solution formula is: THE x =L zx +L k ·(And +X +e -X )+L b THE y =L zy +L k ·(And +Y +e -Y )+L b THE z =L zz +L k ·(And +Z +e -Z )+L b Among them, L b It is a value related to the thickness of the structural plate. For preliminary calculation, L is taken b =50mm, and then make iterative adjustments according to the situation; if L x ≤L wx And L y ≤L wy And L z ≤L wz , then for the actual size L of the module structure x ×L y ×L z If the design is established, continue with the subsequent design steps; otherwise, for the actual size L of the module structure body x ×L y ×L z If the design is not feasible, communicate with the user to determine whether to modify the requirements.

2. The method for designing the layout of a module unit of a reconfigurable spacecraft according to claim 1, characterized in that: If 0N < F0 ≤ 200N, then L k = 30mm; If 200N < F0 ≤ 1000N, then L k = 60mm; if 1000N < F0 ≤ 5000N, then L k = 120mm.

3. The method for designing the configuration layout of a reconfigurable spacecraft module unit according to claim 1, characterized in that: The standard docking interface includes at least one of the following components: a mechanical connection component, an electrical energy transmission component, an information transmission component, and a gas-liquid transmission component.

4. The method for designing the configuration layout of a reconfigurable spacecraft module unit according to claim 3, characterized in that: According to the determined system indicator requirements, the type and specifications of the standard docking interface are designed, including: Determine the specification parameters of mechanical connection components according to the connection relationship requirements between reconfigurable spacecraft module units; Determine the specification parameters of the power transmission components according to the power requirement P0 and voltage requirement V0 of the power transmission between the reconfigurable spacecraft module units; Determine the specification parameters of the information transmission component according to the transmission rate requirement S0 between the reconfigurable spacecraft module units; Determine the specification parameters of gas-liquid transmission components according to the gas-liquid transmission requirements between reconfigurable spacecraft module units; Determine the layout of the internal components of the standard docking interface, combine the above-mentioned mechanical connection components, power transmission components, information transmission components and gas-liquid transmission components, form a variety of feasible internal type combinations and layout methods of the standard docking interface, and select the design state with the smallest volume to carry out subsequent work.

5. The method for designing the layout of a module unit of a reconfigurable spacecraft according to claim 1, characterized in that: The installation positions of the internal loads on the module structure body are designed according to the determined system indicator requirements, including: determining the installation positions of the internal loads on the module structure body according to the mass requirements and the required installation space requirements of the internal loads of the reconfigurable spacecraft module unit; wherein, the internal loads with a mass not greater than a preset mass threshold are installed on the top plate of the module structure body, and the internal loads with a mass greater than the preset mass threshold are installed on the bottom plate of the module structure body.

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