Microsatellite structure based on VPX standard cabin section

Abstract

The invention discloses a microsatellite structure based on a VPX standard cabin section, relates to the technical field of aerospace engineering and satellite structures, and solves the problems that an existing microsatellite structure is poor in universality, low in assembly efficiency, complex in thermal design, difficult to consider both light weight and rigidity and the like. The satellite main body structure adopts the structural form of the VPX standard equipment cabin and the non-standard equipment cabin, and the VPX standard equipment cabin, the non-standard equipment cabin and the solar wing all adopt modular design and are easy to disassemble, assemble and redistribute according to operation requirements; the solar wing can be a solar wing for top side installation or a solar wing for rear side installation and is installed above a VPX standard equipment compartment or behind a non-standard equipment compartment according to task requirements, the non-standard equipment compartment is in lap joint according to the task requirements, an embedded connection structure is adopted in the whole, the design of an external support is reduced, and therefore assembling is simplified, the mass of the whole satellite is reduced, and the overall performance of the satellite is improved. A separating mechanism is installed at the bottom of the non-standard equipment cabin and used for being matched with a carrier rocket connector.

Description

Technical Field

[0001] This invention relates to the technical field of aerospace engineering and satellite structure, and in particular to a microsatellite structure based on the VPX standard module. Background Technology

[0002] In recent years, with the widespread application of near-Earth orbit satellites in Earth observation, communication networking, and space science experiments, microsatellites have become an important development direction for space missions due to their advantages of short development cycles, low costs, and high flexibility. To meet the needs of rapid mission response and mass production, microsatellite platforms are gradually evolving towards standardization, modularization, and scalability. However, existing small satellite structural designs mostly adopt customized solutions for single missions, with significant differences in structural forms, interface standards, and assembly methods between different models. This results in long design cycles, a large amount of repetitive verification work, and difficulty in achieving efficient resource reuse and constellation deployment.

[0003] In traditional small satellite structural design, the main structure is often designed independently based on specific mission requirements, lacking unified standards for material selection, mechanical configuration, and thermal control measures. The absence of a standardized module interface system necessitates repeated adjustments to the integration layout of internal equipment, impacting system design and assembly efficiency. Furthermore, traditional thermal design often employs multi-layered thermal insulation components to control internal temperature distribution. While this approach can mitigate the impact of external heat flow fluctuations on the internal thermal environment to some extent, it increases structural complexity, assembly difficulty, and weight burden, hindering satellite lightweighting and heat homogenization.

[0004] Existing satellite design technologies also suffer from a disconnect between structural and thermal design, lacking systematic and coordinated optimization methods. Structural design focuses on load-bearing and stiffness requirements, while thermal design independently considers heat absorption, dissipation, and insulation functions, resulting in an inability to achieve a reasonable balance between strength, mass, and thermal equilibrium in the overall performance of the satellite. As small satellite missions increasingly demand rapid assembly, on-orbit upgrades, and high integration, traditional non-standardized design methods are becoming increasingly inadequate to meet the needs of engineering and mass production.

[0005] Therefore, existing small satellite structural design technologies still suffer from problems such as insufficient versatility, low integration efficiency, difficulty in balancing lightweight and rigidity, and disconnect between thermal control systems and structural design, making it difficult to support future space application needs that require multiple missions, short cycles, scalability, and constellation-based architecture. Summary of the Invention

[0006] In view of the problems mentioned above, such as poor versatility, low assembly efficiency, complex thermal design, and difficulty in balancing lightweight and rigidity in existing microsatellite structures, the purpose of this invention is to provide a microsatellite structure based on the VPX standard module.

[0007] To achieve the above objectives, the technical solution adopted by the present invention is as follows:

[0008] A microsatellite structure based on a VPX standard module includes: a main satellite structure and solar panels. The main satellite structure includes: a VPX standard equipment module 1 and a non-standard equipment module 2. The VPX standard equipment module 1 is located on top of the non-standard equipment module 2, and the VPX standard equipment module 1 and the non-standard equipment module 2 are detachably connected.

[0009] The solar array is either a top-mounted solar array 3 or a rear-mounted solar array 4, which is detachably mounted on the main structure of the satellite.

[0010] When the number of external equipment is small, a top-mounted solar panel 3 can be used; when the number of external equipment is large, a rear-mounted solar panel 4 can be used to install the external equipment on the top of the VPX standard equipment bay 1.

[0011] The aforementioned microsatellite structure based on the VPX standard module includes a non-standard equipment module 2 comprising: a front end plate 21, a rear end plate 22, a left side plate 23, a right side plate 24, a top side plate 25, and a bottom side plate 26. The top side plate 25 and the bottom side plate 26 are parallel, with the top side plate 25 located above the bottom side plate 26. The front end faces of the top side plate 25 and the bottom side plate 26 are connected by the front end plate 21, the rear end faces of the top side plate 25 and the bottom side plate 26 are connected by the rear end plate 22, the left end faces of the top side plate 25 and the bottom side plate 26 are connected by the left side plate 23, and the right end faces of the top side plate 25 and the bottom side plate 26 are connected by the right side plate 24.

[0012] The microsatellite structure based on the VPX standard module described above includes a non-standard equipment module 2, which further includes a limit switch 27 and a limit switch bracket 28. Two limit switch brackets 28 are installed on the bottom surface of the bottom side plate 26, and one limit switch 27 is installed on each limit switch bracket 28.

[0013] The aforementioned microsatellite structure based on the VPX standard module includes a non-standard equipment module 2, which further includes a flange 29, with a flange 29 installed at each corner of the bottom surface of the bottom side plate 26.

[0014] The aforementioned microsatellite structure based on the VPX standard module further includes a separation mechanism 5, which comprises a base 51, a vibration damper 52, a shape memory alloy puller 53, and a pin connector 54. Each corner of the top of the base 51 is connected to a flange 29, and each corner of the bottom of the base 51 is equipped with a vibration damper 52. A shape memory alloy puller 53 and two pin connectors 54 are mounted on the base 51.

[0015] The aforementioned microsatellite structure based on the VPX standard module includes a VPX standard equipment module 1 comprising: an upper cover plate 11, a lower cover plate 12, a left cover plate 13, a right cover plate 14, and a back plate 15. The lower cover plate 12 is mounted on the upper surface of the top side plate 25, and the upper cover plate 11 is located above the upper cover plate 12. The left end face of the upper cover plate 11 and the left end face of the lower cover plate 12 are connected by the left cover plate 13, and the right end face of the upper cover plate 11 and the right end face of the lower cover plate 12 are connected by the right cover plate 14. The rear end faces of the upper cover plate 11, the lower cover plate 12, the left cover plate 13, and the right cover plate 14 are all connected to the back plate 15.

[0016] The microsatellite structure based on the VPX standard module described above, wherein when the solar panel is a top-mounted solar panel 3, the VPX standard equipment module 1 also includes: a rear cover plate 16, which is installed on the rear side of the back plate 15; and a rear end plate 22 extending upward to be on the same plane as the upper surface of the upper cover plate 11.

[0017] The top-side mounted solar wing 3 includes: a solar wing body mounting plate 31, a solar wing deployment plate 32, a solar wing deployment plate 33, a solar wing connector 34, and a solar wing hinge 35. The solar wing body mounting plate 31 is located on the top side of the satellite's main structure. The top side end faces of the upper cover plate 11 and the rear end plate 22 are connected to the solar wing body mounting plate 31. The two solar wing deployment plates 32 are respectively located on the left and right sides of the satellite's main structure. A solar wing deployment plate 33 is provided on the outer side of each solar wing deployment plate 32. The left and right side end faces of the solar wing body mounting plate 31 are respectively hinged to the top side end face of a solar wing deployment plate 32 through the solar wing hinge 35. The bottom side end face of each solar wing deployment plate 32 is hinged to the bottom side end face of the solar wing deployment plate 33 located on its outer side through the solar wing hinge 35.

[0018] At least one solar wing connector 34 is installed on the outer side of both the left side plate 23 and the right side plate 24. The solar wing deployment plate 32 and the solar wing deployment plate 33 on the left side are closably connected to the solar wing connector 34 on the left side plate 23, and the solar wing deployment plate 32 and the solar wing deployment plate 33 on the right side are closably connected to the solar wing connector 34 on the right side plate 24, thereby realizing the folding and unfolding of the top-mounted solar wing 3.

[0019] The microsatellite structure based on the VPX standard module described above, when the solar panel is the rear-mounted solar panel 4, also includes: an equipment compartment connector 6, an equipment compartment connector 6 is detachably connected between the left side plate 23 and the left cover plate 13, and another equipment compartment connector 6 is detachably connected between the right side plate 24 and the right end cover, and the two equipment compartment connectors 6 are symmetrically arranged on the left and right sides;

[0020] The rear-mounted solar wing 4 includes: solar wing body mounting plate 2 41, solar wing deployment plate 3 42, solar wing connector 2 43, and solar wing hinge 2 44. The solar wing body mounting plate 2 41 is located on the rear side of the satellite main structure and is connected to the rear end plate 22. The two solar wing deployment plates 3 42 are respectively located on the left and right sides of the satellite main structure. The left and right end faces of the solar wing body mounting plate 2 41 are respectively hinged to the rear end face of one solar wing deployment plate 3 42 through the solar wing hinge 2 44.

[0021] Each equipment compartment connector 6 is equipped with a solar panel connector 2 43, and each solar panel connector 2 43 is closably connected to a solar panel deployment plate 3 42, thereby enabling the folding and unfolding of the rear-mounted solar panel 4.

[0022] The aforementioned microsatellite structure based on the VPX standard module includes multiple standard interfaces, lightweight weight reduction grooves, and local reinforcing ribs on the main satellite structure.

[0023] The microsatellite structure based on the VPX standard module described above, wherein the VPX standard equipment module 1 and the non-standard equipment module 2 are made of aluminum alloy or magnesium alloy, the inner surface of the VPX standard equipment module 1 is treated with black anodizing or blackening, and the outer surface of the VPX standard equipment module 1 is treated with natural conductive oxidation.

[0024] The present invention, by employing the above-mentioned technology, has the following positive effects compared with the prior art:

[0025] (1) The main structure of the satellite of the present invention adopts the structural form of VPX standard equipment compartment and non-standard equipment compartment. The VPX standard equipment compartment, non-standard equipment compartment and solar array are all modularly designed, which is easy to disassemble and redistribute according to the operational requirements. The VPX standard equipment compartment, non-standard equipment compartment and solar array are all formed by overlapping structural plates. The structural plates are connected by an embedded structure, which reduces the design of external support, thereby simplifying the assembly and reducing the overall satellite mass.

[0026] (2) In this invention, the solar array can be a top-mounted solar array or a rear-mounted solar array, which is installed above the VPX standard equipment compartment or behind the non-standard equipment compartment according to mission requirements. The non-standard equipment compartment is overlapped according to mission requirements. The whole adopts an embedded connection structure, reducing the design of external support, thereby simplifying assembly and reducing the overall satellite mass. A separation mechanism is installed at the bottom of the non-standard equipment compartment. The separation mechanism is used to adapt to the launch vehicle interface.

[0027] (3) In this invention, the inner surface of the cabin is treated with black anodizing or blackening, while the outer surface of the cabin is treated with natural conductive oxidation. Through this partitioned surface treatment design, the satellite can maintain a relatively stable temperature distribution under different operating conditions. At the same time, for areas with insufficient local heat conduction, measures such as thermal pads can be used for compensation. The overall satellite thermal control system relies on the structural body to achieve coordinated regulation of heat conduction and radiation, thereby simplifying the design of the thermal control system, reducing mass and improving on-orbit reliability. Each structural module forms an overall load-bearing system through standardized interfaces and bolted nodes. During the launch phase, it bears mechanical loads and transmits them to the separation mechanism. During the on-orbit operation phase, each component maintains the thermal balance of the satellite through thermally conductive connections and surface radiation. Attached Figure Description

[0028] Figure 1 This is a schematic diagram of a microsatellite structure based on a VPX standard module, using a top-side mounted solar array. Figure 2 yes Figure 1 The main view. Figure 3 yes Figure 1 Rear view. Figure 4 yes Figure 1 A schematic diagram of the main structure of the satellite. Figure 5 yes Figure 4 A schematic diagram of the bottom side structure. Figure 6 yes Figure 1 A schematic diagram of the structure of the solar panel installed on the top side. Figure 7 yes Figure 6 The main view. Figure 8 yes Figure 7 A side sectional view. Figure 9 yes Figure 1 A structural schematic diagram of the VPX standard equipment compartment. Figure 10 yes Figure 1 A schematic diagram of the structure of the Sino-African standard equipment compartment. Figure 11 yes Figure 10 Exploded view. Figure 12 yes Figure 1 Assembly diagram of the separation mechanism. Figure 13 This is a schematic diagram of a microsatellite structure based on a VPX standard module, using a rear-mounted solar array. Figure 14 yes Figure 13 The main view. Figure 15 yes Figure 13 A schematic diagram of the main structure of the satellite. Figure 16 yes Figure 15 Exploded view. Figure 17 yes Figure 13 A schematic diagram of the structure for mounting solar panels on the middle and rear sides. Figure 18 yes Figure 17 The main view. Figure 19 yes Figure 18A side sectional view. Figure 20 yes Figure 13 A schematic diagram of the structure of the Sino-African standard equipment compartment. Figure 21 yes Figure 20 Exploded view. Figure 22 yes Figure 13 Assembly diagram of the separation mechanism. Figure 23 This is a schematic diagram of an embodiment of the VPX standard equipment bay. Figure 24 This is a schematic diagram of the VPX standard equipment bay. Figure 25 yes Figure 23 A schematic diagram of the locking bar structure.

[0029] In the attached diagram: 1. VPX standard equipment bay; 2. Non-standard equipment bay; 3. Top-mounted solar panel; 4. Rear-mounted solar panel; 5. Separation mechanism; 6. Equipment bay connector; 11. Top cover; 12. Bottom cover; 13. Left cover; 14. Right cover; 15. Back panel; 16. Rear cover; 21. Front panel; 22. Rear panel; 23. Left side panel; 24. Right side panel; 25. Top side panel; 26. Bottom side panel; 27. Travel length. 28. Switch; 29. ​​Limit switch bracket; 30. Flange; 31. Solar wing mounting plate one; 32. Solar wing deployment plate one; 33. Solar wing deployment plate two; 34. Solar wing connector one; 35. Solar wing hinge one; 41. Solar wing mounting plate two; 42. Solar wing deployment plate three; 43. Solar wing connector two; 44. Solar wing hinge two; 51. Base; 52. Vibration damper; 53. Shape memory alloy puller; 54. Pin connector. Detailed Implementation

[0030] The present invention will be further described below with reference to the accompanying drawings and specific embodiments, but this is not intended to limit the scope of the invention.

[0031] Please refer to Figures 1 to 25 As shown, a microsatellite structure based on a VPX standard module is illustrated. The main satellite structure consists of a VPX standard equipment module 1 and a non-standard equipment module 2. The VPX standard equipment module 1 is shown below. Figure 9 As shown, the lower cover plate 12 has multiple mounting ears on its left and right sides, each mounting ear having a screw mounting hole. The lower cover plate 12 and the top side plate 25 of the non-standard equipment compartment 2 are connected by multiple screws. The bottom of the non-standard equipment compartment 2 is equipped with a separation mechanism 5, which is used to adapt to the launch vehicle interface. The solar array can be either a top-mounted solar array 3 or a rear-mounted solar array 4, which can be installed above the VPX standard equipment compartment 1 or behind the non-standard equipment compartment 2 according to mission requirements. The non-standard equipment compartment 2 is overlapped according to mission requirements. The whole adopts an embedded connection structure, reducing the design of external supports, thereby simplifying assembly and reducing the overall satellite mass.

[0032] Furthermore, in a preferred embodiment, the VPX standard equipment compartment 1, as the core load-bearing unit, is composed of an upper cover plate 11, a lower cover plate 12, a left cover plate 13, a right cover plate 14, a back plate 15, and a rear cover plate 16, as shown below. Figure 9 As shown, the left and right end faces of the upper cover plate 11 and the lower cover plate 12 are provided with multiple wedge-shaped grooves, and the upper and lower end faces of the left cover plate 13 and the right cover plate 14 are provided with multiple insert ears. Each insert ear is engaged in a wedge-shaped groove, and each insert ear has a screw hole. The upper cover plate 11, the lower cover plate 12, the left cover plate 13, and the right cover plate 14 are assembled and connected by multiple screws. The inner and outer sides of the left cover plate 13 and the right cover plate 14 are provided with grid-like local reinforcing ribs. The upper cover plate 11 and the lower cover plate 12... The inner side is provided with multiple strip-shaped protruding local reinforcing ribs, which are used to separate and install circuit boards and battery modules, etc. Each strip-shaped protruding local reinforcing rib has multiple lightweight reduction grooves. Multiple strip-shaped local reinforcing ribs are provided between any two adjacent strip-shaped protruding local reinforcing ribs. At least one standard interface is provided on both the back plate 15 and the rear cover plate 16. The surface of the upper cover plate 11 has multiple arrayed screw holes for assembling the solar panels.

[0033] Furthermore, in a preferred embodiment, the non-standard equipment compartment 2 adopts a box-type configuration, consisting of a front panel 21, a rear panel 22, a left side panel 23, a right side panel 24, a top side panel 25, and a bottom side panel 26. The front panel 21, rear panel 22, left side panel 23, right side panel 24, top side panel 25, and bottom side panel 26 can be replaced according to actual application requirements. Local reinforcing ribs and lightweight weight-reducing grooves are arranged according to load requirements, and the number of standard interfaces is increased or decreased, and the configuration of standard interfaces is selected.

[0034] Furthermore, in a preferred embodiment, the VPX standard equipment compartment 1, the non-standard equipment compartment 2, and the solar array are all modularly designed, making them easy to disassemble, assemble, and redistribute according to operational needs; the VPX standard equipment compartment 1, the non-standard equipment compartment 2, and the solar array are all formed by overlapping structural panels, with embedded connection structures between the structural panels, reducing the design of external supports, thereby simplifying assembly and reducing the overall satellite mass.

[0035] Furthermore, in a preferred embodiment, both the VPX standard equipment compartment 1 and the non-standard equipment compartment 2 are made of aluminum or magnesium alloy, which have excellent thermal conductivity. The compartments are connected by mounting lugs, connecting screws, and locking strips, forming a continuous heat conduction path that allows for rapid heat conduction and diffusion within the satellite, preventing localized overheating. The inner surface of the compartment is treated with black anodizing or blackening to improve infrared absorption and emissivity, effectively dissipating heat generated by the internal equipment. The outer surface of the compartment uses a natural-colored conductive anodizing process, maintaining electromagnetic compatibility while controlling solar radiation absorption and reducing external heat load. Through this zoned surface treatment design, the satellite can maintain a relatively stable temperature distribution under different operating conditions.

[0036] Furthermore, in a preferred embodiment, when the solar array is a top-mounted solar array 3, the upper cover plate 11 of the VPX standard equipment compartment 1 and the rear end plate 22 of the non-standard equipment compartment 2 are both connected to the solar array mounting plate 31; the lower cover plate 12 of the VPX standard equipment compartment 1 and the top side plate 25 of the non-standard equipment compartment 2 are connected by multiple screws; two solar array connectors 34 are respectively installed on the left and right sides of the non-standard equipment compartment 2, and embedded parts are provided in the solar array deployment plate 32 and the solar array deployment plate 33, which cooperate with the solar array hinge 35 to realize the folding and unfolding of the top-mounted solar array 3; this structure realizes the modular connection of the VPX standard equipment compartment 1, the non-standard equipment compartment 2 and the top-mounted solar array 3; the optimized frame structure can withstand high dynamic loads while ensuring installation accuracy, and meet the near-Earth orbit launch conditions.

[0037] Furthermore, in a preferred embodiment, when the solar wing is a rear-mounted solar wing 4, the lower cover plate 12 of the VPX standard equipment compartment 1 and the top side plate 25 of the non-standard equipment compartment 2 are connected by multiple screws; the VPX standard equipment compartment 1 and the non-standard equipment compartment 2 are connected and assembled by two equipment compartment connectors 6 and multiple screws; the solar wing body mounting plate 41 is connected to the rear end plate 22 of the non-standard equipment compartment 2; the two solar wing connectors 43 are respectively installed on the two equipment compartment connectors 6, and embedded parts are provided in the two solar wing deployment plates 42, which cooperate with the solar wing hinge 44 to realize the folding and unfolding of the rear-mounted solar wing 4; this structure realizes the modular connection of the VPX standard equipment compartment 1, the non-standard equipment compartment 2 and the rear-mounted solar wing 4; the optimized frame structure can withstand high dynamic loads while ensuring installation accuracy, and meet the near-Earth orbit launch conditions.

[0038] The solar array is folded and deployed via embedded parts, which can be driven by spring / burst bolts, motors, or shape memory alloys (SMA). Existing technologies are used in this folding and deployment method.

[0039] Furthermore, in a preferred embodiment, the VPX standard equipment compartment 1, the non-standard equipment compartment 2, and the solar array are modularly connected, and the solar array 3 for top mounting and the solar array 4 for rear mounting are selected according to actual working needs.

[0040] The above are merely preferred embodiments of the present invention and are not intended to limit the implementation methods and protection scope of the present invention.

[0041] In addition to the above, the present invention also has the following embodiments:

[0042] In a further embodiment of the present invention, the satellite's main structure adopts a structure of "VPX standard equipment compartment 1" and "non-standard equipment compartment 2". The VPX standard equipment compartment 1 and the non-standard equipment compartment 2 are mechanically connected by 10 M6 hexagonal screws on the mounting lugs of the VPX standard compartment. A separation mechanism 5 is installed at the bottom of the non-standard equipment compartment 2 to realize the separation function of the satellite and the launch vehicle interface. The solar array is directly configured above the VPX standard equipment compartment 1 or behind the non-standard equipment compartment 2, thereby simplifying the connection structure, reducing the overall satellite structure mass, and improving the overall satellite rigidity. Embedded parts are provided inside the body plate and deployment plate of the solar array, and direct connection is achieved through a hinge structure, avoiding the need for a separate mounting bracket and further optimizing the mass. For the two different installation orientations of the solar array, see [reference needed]. Figure 1 and Figure 13 .

[0043] This invention, through standardized, modular, and integrated design concepts, achieves the universality, scalability, and comprehensive mechanical and thermal optimization of satellite structures, providing an efficient and reliable structural design approach for next-generation microsatellite platforms. Addressing the problems of poor universality, low assembly efficiency, complex thermal design, and difficulty in balancing lightweighting and stiffness in existing microsatellite structures, this invention proposes a modular and standardized satellite structural system. This system achieves platform standardization, lightweighting, and integrated mechanical and thermal optimization, encompassing satellite structural layout, structural design, material selection, thermal control strategies, and lightweight reinforcement design. It aims to improve the structural strength, thermal stability, and modular integration efficiency of microsatellite platforms. Compared with existing customized structural design methods for small satellites, this invention significantly improves and enhances structural universality, assembly efficiency, thermal control performance, and reliability, effectively supporting the rapid development and mass production of near-Earth orbit microsatellites.

[0044] First, this invention introduces a modular system for standard VPX modules, unifying satellite structural interface standards and module size specifications, enabling the satellite platform to flexibly combine different functional modules according to mission requirements. This design standardizes the mechanical, electrical, and thermal control interfaces between modules, significantly improving structural versatility and mission expandability. The platform reusability rate can be increased to over 50%, and the overall satellite assembly cycle can be shortened by approximately 30%.

[0045] Secondly, in terms of structural performance, this invention employs a lightweight weight-reduction groove design and a localized reinforcing rib structure to form a cabin configuration that balances rigidity and mass. Compared with traditional box structures, the overall cabin mass is reduced by an average of approximately 15% to 20%, while the first-order natural frequency is increased by approximately 25%, resulting in a more stable structural response under launch vibration environments. The optimized frame structure can withstand higher dynamic loads while ensuring installation accuracy, meeting the requirements for low Earth orbit launch.

[0046] In terms of thermal design, this invention simplifies and integrates the thermal control system through the synergistic design of structural thermal conduction and surface radiation. The rational combination of cabin structural materials and surface treatments controls the maximum low-temperature temperature difference inside the low-orbit conventional heat-consuming satellite to around 20°C, significantly improving equipment temperature stability. It can maintain the satellite's thermal balance without the need for large-area insulation layers or additional heating devices, thereby reducing the mass, heat consumption, and complexity of thermal control system implementation.

[0047] Furthermore, the structural design of this invention fully considers standardization and ease of operation in the manufacturing and assembly processes. A large number of structural panels utilize M3 screws, resulting in modular assembly and highly versatile tools, enabling rapid assembly and replacement of different satellite models. Compared to traditional structural solutions, the overall satellite assembly time can be reduced by approximately 40%, and the preparation cycle for structural vibration testing and thermal vacuum testing is simultaneously shortened.

[0048] In summary, this invention significantly improves the engineering level and on-orbit reliability of microsatellite structures through innovations in standard module design, lightweight configuration, integrated thermo-mechanical optimization, and modular assembly. This method can shorten design cycles, reduce costs, and improve mission response speed while ensuring performance. It is suitable for the mass production and constellation-based development of low-Earth orbit microsatellite platforms, achieving standardization, lightweighting, and thermo-mechanical optimization of low-Earth orbit microsatellite platforms.

[0049] In satellite system architecture, when using the VPX (VITA 46) standard equipment bay (hereinafter referred to as VPX board) as the core processing unit, its mechanical dimensions mainly follow two modular standard specifications: 3U and 6U. The 3U specification, i.e., 100mm (height) x 160mm (depth), features a compact physical form and low power consumption. It is suitable for satellite platforms with high functional integration and strict space and weight constraints. For even higher performance requirements, the 6U specification, i.e., 233.35mm (height) x 160mm (depth), can be used. This specification offers stronger processing power, higher I / O bandwidth, and greater expansion flexibility, suitable for integrating high-performance multi-core processors, large-capacity storage units, or complex reconfigurable logic (such as field-programmable gate arrays).

[0050] Taking the 3U VPX standard equipment bay 1 as an example of this invention, the structure of this bay mainly consists of four structural plates connected in pairs by 10 M3 hexagonal screws. The left and right structural plates are identical in structure, and the top of the upper structural plate has several M3 threaded holes for mounting solar panels, antennas, or other extra-satellite equipment and their support brackets. The front panel of the bay is used to insert VPX boards and has several M3 threaded holes for vibration damping installation. The back of the bay is connected to a PCB backplate by several M3 hexagonal screws. The VPX standard bay can be equipped with several 3U slots according to the actual needs of the satellite mission, and can accommodate 0.8in, 0.85in, and 1in VPX 3U standard boards. The lithium-ion battery pack can also be integrated and installed in the center of the VPX standard bay. The VPX boards are locked and heat-conducted to the entire VPX standard equipment bay 1 by locking strips and connected to the PCB backplate by two pins. In addition to the locking strip, the lithium-ion battery pack should also be connected to the compartment at the front panel using hexagonal screws to reduce relative vibration. A schematic diagram of the 3U VPX standard equipment compartment 1 is shown below. Figure 23 and Figure 24 As shown.

[0051] The envelope dimensions of the non-standard equipment compartment 2 should be determined based on the specific satellite subsystem equipment and payload configuration requirements, but should not be smaller than the standard compartment dimensions. The non-standard equipment compartment 2 is assembled from six structural panels connected by M3 hexagonal screws. The connection surface with the VPX standard equipment compartment 1 has pre-drilled openings for cable routing.

[0052] When the number of external equipment is small, a top-mounted solar panel 3 can be used; when the number of external equipment is large, a rear-mounted solar panel 4 can be used to install the external equipment on the top of the VPX standard equipment bay 1.

[0053] To achieve both lightweighting and enhanced rigidity, this invention incorporates weight-reducing groove structures in the non-load-bearing areas of the VPX compartment and reinforcing ribs or plates in critical structural regions. The weight-reducing grooves effectively reduce redundant mass, while the reinforcing ribs resist stress concentration caused by locking preload and launch loads, significantly improving the out-of-plane stiffness and bending resistance of the compartment and ensuring the precision and stability of the board connections. The non-standard equipment compartment 2 adopts a box-plate configuration, consisting of six structural plates connected by M3 hexagonal screws. Local thickening or reinforcement is applied to equipment interfaces and opening areas to form a continuous rigid frame, enhancing overall strength and vibration resistance. The reinforcing rib layout is optimized to construct a continuous and complete rigid structural frame, comprehensively covering the following key stiffness control areas:

[0054] 1. The four edges of the structural plate ensure a rigid connection with other main structures of the satellite and efficient load transfer;

[0055] 2. All mounting holes for equipment or brackets should be designed to prevent localized deformation under preload and dynamic load.

[0056] 3. Effectively suppresses stress concentration and structural weakening caused by all functional openings around them;

[0057] 4. Other vibration-sensitive or specially reinforced areas identified through analysis can be targeted to strengthen specific dynamic responses or areas with stiffness deficiencies.

[0058] In terms of thermal design, this invention proposes an integrated design approach that combines structure and thermal control. Instead of using traditional multi-layered insulation components, it achieves passive thermal balance for the celestial body through the thermal conductivity and surface treatment properties of the materials. Both the VPX standard equipment compartment 1 and the non-standard equipment compartment 2 are made of aluminum or magnesium alloys with excellent thermal conductivity. The compartments are connected by mounting lugs, connecting screws, and locking strips to form a continuous heat conduction path, allowing for rapid heat transfer and diffusion within the celestial body and preventing localized overheating.

[0059] The inner surface of the satellite's interior is treated with black anodizing or blackening to improve infrared absorption and emissivity, allowing heat generated by internal equipment to dissipate effectively. The outer surface, however, uses a natural-colored conductive oxidation process to control solar radiation absorption while maintaining electromagnetic compatibility and reducing external heat load. This zoned surface treatment design ensures the satellite maintains a relatively stable temperature distribution under various operating conditions.

[0060] Furthermore, this invention ensures high flatness of the contact surface on the equipment installation interface of the non-standard equipment compartment 2 to reduce contact thermal resistance; for areas with insufficient local thermal conduction conditions, measures such as thermally conductive pads can be used for compensation. The overall satellite thermal control system relies on the structural body to achieve coordinated regulation of thermal conduction and radiation, thereby simplifying the thermal control system design, reducing mass, and improving on-orbit reliability.

[0061] The working principle of this invention is as follows: each structural module forms an integrated load-bearing system through standardized interfaces and bolted joints, bearing mechanical loads during launch and transmitting them to the separation mechanism; during on-orbit operation, each component maintains the satellite's thermal balance through thermally conductive connections and surface radiation. This method achieves a unified approach to structural design, thermal design, and integrated assembly. The standardized module design improves the satellite platform's compatibility and scalability, significantly shortens the design cycle, and enhances reliability.

[0062] The key technologies of this invention are: proposing an innovative module structure system based on VPX standard modules, and providing structural dimension standards for VPX standard equipment modules in 3U / 6U sizes. A balance between structural strength and mass is achieved through standard interfaces, lightweight weight-reduction grooves, and local reinforcing ribs; a thermo-mechanical coupling design system is constructed through surface conductive oxidation and anodizing processes; and rapid satellite assembly and multi-mission expansion are achieved through a unified structural interface, providing a systematic engineering method for the constellation and mass production of small satellites.

[0063] The satellite configuration of this invention consists of a standard VPX equipment compartment 1 and a non-standard equipment compartment 2. The two compartments are mechanically connected by standard mounting lugs and screws, and a separation mechanism is provided at the bottom to adapt to the launch vehicle interface. The solar panels can be installed above the standard compartment or behind the non-standard compartment according to mission requirements, using an embedded connection structure to reduce the design of external supports, thereby simplifying assembly and reducing the overall satellite weight.

[0064] The VPX standard equipment compartment 1, serving as the core load-bearing unit, adopts a 3U or 6U modular design. The compartment consists of structural panels, a front panel, a back panel, and locking strips, and can accommodate circuit boards and battery modules, featuring high integration and excellent thermal conductivity. The non-standard equipment compartment 2 employs a box-plate configuration with flexible internal layout. Interfaces can be pre-installed and localized reinforcement designs can be incorporated according to load requirements, ensuring a balance between the satellite's center of mass and structural rigidity.

[0065] Locking strip Figure 25 As shown, its structure adopts an alternating arrangement of isosceles trapezoids with screws inserted. By rotating the screws at the ends, the inner part of the frame can be made to protrude outward, thereby locking the VPX standard equipment compartment 1 and its internal structural components.

[0066] To achieve both lightweight and high rigidity, this invention incorporates weight-reduction grooves in the non-load-bearing areas of the module and reinforcing ribs at critical connection points. The weight-reduction grooves effectively reduce redundant mass, while the reinforcing ribs enhance local rigidity and vibration resistance, ensuring structural stability under launch loads.

[0067] In terms of thermal design, this invention adopts an integrated design approach for structure and thermal control. Utilizing the high thermal conductivity of aluminum and magnesium alloys, and the heat conduction paths formed by structural connectors, rapid heat transfer and diffusion within the spacecraft are achieved. The inner surface of the cabin is treated with blackening or anodizing to enhance heat dissipation, while the outer surface uses natural-colored conductive anodizing to reduce heat absorption and maintain electromagnetic compatibility, thereby achieving stable passive thermal equilibrium during on-orbit operation. Through the coordinated control of structural thermal conduction and surface radiation, traditional multi-layer thermal insulation components are eliminated, reducing the weight of the thermal control system and improving on-orbit reliability.

[0068] This invention achieves the unification of structural design, assembly, and thermal control systems through standardized module interfaces, modular design, and thermo-mechanical synergistic optimization, significantly improving the compatibility, scalability, and engineering efficiency of the satellite platform. It is suitable for the rapid mass production and constellation deployment of microsatellites in low Earth orbit.

[0069] The above are merely preferred embodiments of the present invention and are not intended to limit the implementation methods and protection scope of the present invention. Those skilled in the art should recognize that any equivalent substitutions and obvious changes made based on the description and illustrations of the present invention should be included within the protection scope of the present invention.

Claims

1. A microsatellite structure based on a VPX standard module, characterized in that, include: The satellite main structure and solar panels include: a VPX standard equipment compartment (1) and a non-standard equipment compartment (2). The VPX standard equipment compartment (1) is located on top of the non-standard equipment compartment (2). The VPX standard equipment compartment (1) and the non-standard equipment compartment (2) are detachably connected. The solar array is a top-mounted solar array (3) or a rear-mounted solar array (4), which is detachably mounted on the main structure of the satellite.

2. The microsatellite structure based on the VPX standard module according to claim 1, characterized in that, The non-standard equipment compartment (2) includes: a front end plate (21), a rear end plate (22), a left side plate (23), a right side plate (24), a top side plate (25), and a bottom side plate (26). The top side plate (25) and the bottom side plate (26) are parallel. The top side plate (25) is located above the bottom side plate (26). The front end faces of the top side plate (25) and the bottom side plate (26) are connected by the front end plate (21). The rear end faces of the top side plate (25) and the bottom side plate (26) are connected by the rear end plate (22). The left end faces of the top side plate (25) and the bottom side plate (26) are connected by the left side plate (23). The right end faces of the top side plate (25) and the bottom side plate (26) are connected by the right side plate (24).

3. The microsatellite structure based on the VPX standard module according to claim 2, characterized in that, The non-standard equipment compartment (2) also includes: limit switch (27) and limit switch bracket (28). Two limit switch brackets (28) are installed on the bottom surface of the bottom side plate (26), and one limit switch (27) is installed on each limit switch bracket (28).

4. The microsatellite structure based on the VPX standard module according to claim 3, characterized in that, The non-standard equipment compartment (2) also includes: flanges (29), with a flange (29) installed at each corner of the bottom surface of the bottom side plate (26).

5. The microsatellite structure based on the VPX standard module according to claim 4, characterized in that, Also includes: The separation mechanism (5) includes: a base (51), a damper (52), a shape memory alloy puller (53), and a pin connector (54). Each corner of the top of the base (51) is connected to a flange (29), and each corner of the bottom of the base (51) is equipped with a damper (52). A shape memory alloy puller (53) and two pin connectors (54) are installed on the base (51).

6. The microsatellite structure based on the VPX standard module according to claim 2, characterized in that, The VPX standard equipment compartment (1) includes: an upper cover plate (11), a lower cover plate (12), a left cover plate (13), a right cover plate (14), and a back plate (15). The lower cover plate (12) is installed on the upper surface of the top side plate (25). The upper cover plate (11) is located above the upper cover plate (11). The left end face of the upper cover plate (11) and the left end face of the lower cover plate (12) are connected through the left cover plate (13). The right end face of the upper cover plate (11) and the right end face of the lower cover plate (12) are connected through the right cover plate (14). The rear end faces of the upper cover plate (11), the lower cover plate (12), the left cover plate (13), and the right cover plate (14) are all connected to the back plate (15).

7. The microsatellite structure based on the VPX standard module according to claim 6, characterized in that, When the solar panel is a top-mounted solar panel (3), the VPX standard equipment compartment (1) also includes: a rear cover plate (16), which is installed on the rear side of the back panel (15); and a rear end plate (22) that extends upward to be on the same plane as the upper surface of the upper cover plate (11). The top-side mounted solar wing (3) includes: a solar wing body mounting plate one (31), a solar wing deployment plate one (32), a solar wing deployment plate two (33), a solar wing connector one (34), and a solar wing hinge one (35). The solar wing body mounting plate one (31) is located on the top side of the satellite's main structure. The top side end faces of the upper cover plate (11) and the rear end plate (22) are connected to the solar wing body mounting plate one (31). The two solar wing deployment plates one (32) are respectively located on the satellite's main body. On the left and right sides of the structure, a solar wing deployment plate 2 (33) is provided on the outside of each solar wing deployment plate 1 (32). The left and right end faces of the solar wing body mounting plate 1 (31) are respectively hinged to the top end face of a solar wing deployment plate 1 (32) through a solar wing hinge 1 (35). The bottom end face of each solar wing deployment plate 1 (32) is hinged to the bottom end face of the solar wing deployment plate 2 (33) located outside it through a solar wing hinge 1 (35). At least one solar wing connector (34) is installed on the outer side of both the left side plate (23) and the right side plate (24). The solar wing deployment plate (32) and the solar wing deployment plate (33) on the left side are connected to the solar wing connector (34) on the left side plate (23) in an openable and closable manner. The solar wing deployment plate (32) and the solar wing deployment plate (33) on the right side are connected to the solar wing connector (34) on the right side plate (24) in an openable and closable manner, thereby realizing the folding and unfolding of the solar wing (3) for top side mounting.

8. The microsatellite structure based on the VPX standard module according to claim 6, characterized in that, When the solar panel is a rear-mounted solar panel (4), it also includes: an equipment compartment connector (6), an equipment compartment connector (6) is detachably connected between the left side plate (23) and the left cover plate (13), and another equipment compartment connector (6) is detachably connected between the right side plate (24) and the right end cover, and the two equipment compartment connectors (6) are symmetrically arranged on the left and right sides; The rear-mounted solar wing (4) includes: solar wing body mounting plate two (41), solar wing deployment plate three (42), solar wing connector two (43) and solar wing hinge two (44). The solar wing body mounting plate two (41) is located on the rear side of the satellite main structure and is connected to the rear end plate (22). The two solar wing deployment plates three (42) are respectively located on the left and right sides of the satellite main structure. The left and right end faces of the solar wing body mounting plate two (41) are respectively hinged to the rear end face of one solar wing deployment plate three (42) through the solar wing hinge two (44). Each equipment compartment connector (6) is equipped with a solar wing connector (43), and each solar wing connector (43) is connected to a solar wing deployment plate (42) in an openable and closable manner, thereby enabling the folding and unfolding of the rear-mounted solar wing (4).

9. The microsatellite structure based on the VPX standard module according to claim 6, characterized in that, The satellite's main structure is equipped with multiple standard interfaces, lightweight weight-reduction grooves, and local reinforcing ribs.

10. The microsatellite structure based on the VPX standard module according to claim 6, characterized in that, The VPX standard equipment compartment (1) and the non-standard equipment compartment (2) are made of aluminum alloy or magnesium alloy. The inner surface of the VPX standard equipment compartment (1) is treated with black anodizing or blackening, and the outer surface of the VPX standard equipment compartment (1) is treated with natural conductive oxidation.

Images