A type of ultra-large planar array space computing center

Abstract

This invention discloses an ultra-large planar array-type space computing center, belonging to the field of aerospace technology and space infrastructure. Deployed in a diurnal orbit, the space computing center comprises multiple computing satellites physically docked in orbit via a docking mechanism. Each computing satellite includes a satellite platform and its onboard computing payload, a concentrated solar power system, and a space pump-driven fluid loop system. The satellite platform is also equipped with a docking mechanism for inter-satellite physical docking. Multiple computing satellites form a rigid planar array through the docking mechanism, enabling energy scheduling among the concentrated solar power systems. This invention's space computing center improves system energy utilization efficiency and power supply reliability through multi-satellite physical docking and coordinated energy scheduling.

Description

Technical Field

[0001] This invention relates to the fields of aerospace technology and space infrastructure technology, specifically to an ultra-large planar array type space computing center. Background Technology

[0002] Against the backdrop of the rapid development of artificial intelligence technology, the global demand for computing power is growing exponentially, continuously driving the construction of ultra-large-scale computing centers. However, traditional ground-based computing centers face a series of systemic problems in their construction and operation, mainly including extremely high energy consumption, huge consumption of cooling water resources, increased greenhouse gas emissions, and challenges such as dense land use and difficulty in rapid and flexible deployment.

[0003] Space-based computing centers are considered a crucial development direction for overcoming the aforementioned limitations. They can directly utilize space-based solar energy for power and achieve efficient heat dissipation by leveraging the cold, dark environment of space. Among these, the twilight orbit, because its orbital plane largely coincides with the Earth's terminator, allows satellites to receive near-continuous solar radiation, making it an ideal orbit for deploying space-based computing systems.

[0004] Currently, the deployed space computing centers mainly take two forms:

[0005] The first type is the monolithic computing satellite. This type of satellite highly integrates computing payloads, energy systems, and thermal control systems within a single satellite platform. Its technological development is limited by the physical envelope of the single satellite platform, launch mass, and the balance between power and heat dissipation. Specifically, the area of ​​the solar array that a single satellite can carry is limited, resulting in an upper limit to its power generation capacity; simultaneously, the area of ​​its radiating heat dissipation surface also determines the upper limit of the system's heat dissipation. These physical limitations collectively constrain the scale and power density of the computing payload that a single satellite can carry, making it difficult to achieve a breakthrough in on-orbit computing power.

[0006] The second type is the distributed satellite constellation. This type of system consists of multiple satellites carrying computing payloads, forming a network through orbital configurations and inter-satellite links. In existing constellations of this type, the satellites are physically independent, primarily exchanging data and coordinating tasks through wireless communication networks. However, this loosely structured networking approach has inherent limitations: the energy systems of each satellite are independent, preventing direct power sharing and cross-satellite scheduling. This means that energy shortages for a satellite during Earth's shadow period or when its power generation system fails cannot be directly compensated for by neighboring satellites, limiting the overall system's energy efficiency and power supply reliability.

[0007] Therefore, there is an urgent need in this field for a new space computing system architecture that can break through the limitations of single-star physics and achieve deep collaboration and integration of energy systems among multiple stars. Summary of the Invention

[0008] The purpose of this invention is to propose an ultra-large planar array type space computing center to solve the problems in the existing technology where the computing power of a single computing satellite is limited by its physical envelope and the distributed satellite constellation cannot achieve deep synergy of energy systems due to physical separation.

[0009] To achieve the above objectives, this invention proposes an ultra-large planar array type space computing center, deployed in a dawn-dusk orbit, including multiple computing satellites physically docked in orbit;

[0010] Each of the computing satellites includes a satellite platform, and a computing payload, a concentrated light energy system, and a space pump-driven fluid loop system mounted on the satellite platform. The satellite platform is equipped with a docking mechanism for on-orbit physical docking between satellites.

[0011] Multiple computing satellites form a large planar array-type rigid whole through the docking mechanism, and energy scheduling is performed between the concentrated energy systems based on the docking mechanism.

[0012] Compared with existing technologies, the ultra-large planar array type space computing center proposed in this invention has the following beneficial effects:

[0013] 1) Effective integration of system resources: By physically docking multiple computing satellites in orbit into a planar array structure, the physical resources originally scattered across multiple independent satellites are integrated into a unified system. This architecture breaks through the limitations of a single satellite in terms of structural space and power capacity, expanding the total area of ​​the solar array and the total scale of the computing payload, providing a foundation for supporting high-power-density computing payloads.

[0014] 2) Enhancing the coordination capabilities of energy systems: The physical connection established based on the satellite docking mechanism provides a transmission channel for coordination between energy systems. By constructing a unified power supply bus across satellites, direct power transmission from energy-surplus satellites to energy-scarce satellites can be achieved. This scheduling mechanism improves the reliability of energy supply under different operating conditions and increases the overall energy utilization efficiency.

[0015] In summary, this invention provides a scalable and highly reliable technical solution for space computing systems through an architecture of multi-satellite on-orbit physical docking and energy coordinated scheduling. Attached Figure Description

[0016] The accompanying drawings, which form part of this specification, are used to provide a further understanding of the invention. The illustrative embodiments of the invention and their descriptions are used to explain the invention and do not constitute an undue limitation of the invention. In the drawings:

[0017] Figure 1 This is a schematic diagram illustrating the effect of an ultra-large planar array type space computing center proposed in this invention;

[0018] Figure 2 This is a schematic diagram of the architecture of a concentrating energy system proposed in this invention;

[0019] Figure 3 This is a schematic diagram of the on-orbit deployment structure of a concentrating solar cell array proposed in this invention.

[0020] Figure 4 This is a schematic diagram of the structure of a solar cell array panel proposed in this invention;

[0021] Figure 5 This is a schematic diagram of a focusing array proposed in this invention;

[0022] Figure 6 This is a schematic diagram of the structure of a light-concentrating module proposed in this invention;

[0023] Figure 7 This is a schematic diagram of the heat dissipation path of the solar cell in the concentrating module proposed in this invention;

[0024] Figure 8 This is a schematic diagram of the outward thermal radiation path of the primary mirror substrate of the focusing module proposed in this invention.

[0025] Figure 9 This is a schematic diagram of the architecture of a space pump-driven fluid loop system proposed in this invention;

[0026] Figure 10 This is a schematic diagram of the structure of the single-layer cooling plate proposed in this invention;

[0027] Figure 11 This is a schematic diagram of the unfolding process of the flexible radial cooling plate proposed in this invention.

[0028] In the diagram: 10. Satellite platform; 11. Docking mechanism; 13. Flexible radiant cooling plate; 20. Solar cell array panel; 21. Thin-walled carbon fiber tube frame; 22. Concentrating array; 23. Concentrating module; 1. Primary mirror; 101. Primary mirror substrate; 102. Reflective layer; 103. Transparent infrared radiation layer; 104. Thermally conductive radiation layer; 3. Secondary mirror; 4. Secondary mirror support; 5. Solar cell; 6. Light funnel; 7. Heat diffuser; 30. Single radiant cooling plate; 31. Rigid fluid pipeline; 32. Lightweight high-strength frame; 33. Flexible radiant membrane; 41. Mounting plane; 42. Hollowed-out section. Detailed Implementation

[0029] The technical solutions in the embodiments of the present invention will now be clearly and completely described with reference to the accompanying drawings. The described embodiments are only a part of the implementation of the present invention, not all of them. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0030] It should be noted that, in the description of this invention, "an embodiment" or "an embodiment of the invention" refers to a specific feature, structure, or characteristic that can be included in at least one implementation of the invention. The phrase "in one embodiment" appearing in different places in the specification does not necessarily refer to the same embodiment, nor are they mutually exclusive individual or selective embodiments. This invention can also be implemented in ways other than those described herein, and any equivalent modifications made by those skilled in the art without departing from the concept of the invention fall within the scope of protection of this invention.

[0031] Before describing the specific solution of this invention, the underlying technical background should first be explained. Currently, deploying high-performance computing facilities in dawn-dusk orbits faces the dual constraints of single-satellite resource limitations and insufficient multi-satellite coordination. Individual satellites are limited by launch mass and structural envelope, resulting in a clear upper limit to their power generation capacity. While satellite constellations networked via wireless links are logically interconnected, they remain physically independent entities. Their energy systems struggle to overcome communication latency and bandwidth limitations, hindering deep integration such as direct power sharing.

[0032] Against this backdrop, the design concept of this invention lies in constructing a modular on-orbit physical architecture that directly connects multiple standardized computing satellites through docking mechanisms to form a structurally unified planar array. The core of this design is to establish a cross-satellite energy scheduling physical foundation through rigid mechanical connections and power interconnection interfaces, thereby improving energy utilization efficiency and power supply reliability at the system level.

[0033] See Figure 1 This invention proposes an ultra-large planar array type space computing center, deployed in a dawn-dusk orbit, including multiple computing satellites physically docked in orbit;

[0034] Each of the computing satellites includes a satellite platform 10, and a computing payload, a concentrated light energy system, and a space pump-driven fluid loop system disposed on the satellite platform. The satellite platform 10 is equipped with a docking mechanism 11 for on-orbit physical docking between satellites.

[0035] Multiple computing satellites form a large planar array-type rigid whole through the docking mechanism 11, and energy scheduling is performed between the concentrated energy systems based on the docking mechanism 11.

[0036] This invention constructs an ultra-large planar array-type space computing center by integrating multiple computing satellites in orbit. The realization of this ultra-large planar array structure is based on the following two-level design:

[0037] At the single-satellite level, a modular functional planar array design is adopted. Each computing satellite's docking surface is equipped with a high-precision, high-rigidity mechanical docking mechanism, ensuring a stable and deformably controllable rigid whole after multi-satellite docking. Its planar characteristics are jointly guaranteed by the positioning accuracy of the docking mechanism and the structural rigidity of the satellite itself. Specifically, the dual solar arrays in the concentrated energy system adopt a multi-level array configuration of "concentrating module-concentrating array-solar plate-solar array," with each solar array forming a large-area power generation structure with high planarity. The radiant cooling plate assembly in the space pump-driven fluid loop system also adopts an array design, consisting of multiple identical radiant cooling plates connected by hinges, forming a large-area continuous radiative heat dissipation surface after in-orbit deployment. At the system level, a planar array is constructed through multi-satellite physical interconnection. Based on the standardized and modular design of the single satellites, multiple computing satellites form a structurally continuous planar array on a macroscopic scale through physical docking, integrating discrete single satellites into a unified space infrastructure. After docking, the solar arrays of each satellite collectively constitute a continuous power generation plane.

[0038] Through a combination of standardized satellite modules and rigid physical docking, the system constructs a space computing center that evolves from an independent satellite into a massive planar array. The entire array operates as a unified whole, enabling centralized command reception and resource management via an inter-satellite interconnection network and a unified management system.

[0039] This invention connects multiple standardized computing satellites in orbit using a rigid mechanical docking mechanism, constructing a planar array with a unified structure, and achieves the following technical effects at the system architecture level:

[0040] 1) Modular scalability: The system uses standardized computing satellites as basic modules and is assembled in orbit through a unified satellite docking mechanism, enabling scalability. This system architecture is not dependent on single launch capability and can linearly expand the total power generation and computing output scale of the system by adding standard modules according to mission requirements.

[0041] 2) System Resource Integration: This architecture integrates the physical resources originally scattered across multiple independent satellites into a unified system, overcoming the limitations of a single satellite in terms of structural space and mass. Through physical docking, it achieves the integration of the total area of ​​the solar cell array and the total scale of computing power payload.

[0042] 3) Coordinated Energy Dispatch: Based on the physical channel established by the satellite docking mechanism, energy coordination capabilities beyond wireless links are achieved. By integrating power interconnection interfaces within the docking mechanism, a unified power supply bus across satellites is constructed, enabling direct power transmission from energy-surplus satellites to energy-scarce satellites, thus improving the reliability of the system's energy supply under different operating conditions.

[0043] In summary, this invention achieves an architectural transformation from independent satellites to an integrated space computing infrastructure through the combined design of standardized satellite modules, rigid physical docking, and resource interconnection networks, providing a system-level solution for large-scale space computing missions.

[0044] In one embodiment, the docking mechanism is equipped with a power interconnection interface. When multiple computing satellites are physically docked, their respective power interconnection interfaces are electrically connected to each other to form an inter-satellite power supply bus, which is a unified power supply bus covering the entire satellite array.

[0045] The space computing center also has a cross-satellite energy convergence system. Based on this power supply bus, energy scheduling between concentrated energy systems is achieved through the cross-satellite energy convergence system.

[0046] The inter-satellite energy convergence system includes multiple local power collectors, a power dispatch manager, and a power switch array. The local power collectors are located on each computing satellite and are used to collect the electrical parameters output by the satellite's concentrated solar power system. The power dispatch manager is located on one or more computing satellites and is communicatively connected to each local power collector, used to generate dispatch commands based on the electrical parameters. The power switch array is located on the power supply bus and is controlled by the power dispatch manager; each satellite access point in the power switch array has an independently controllable switch.

[0047] This embodiment achieves energy dispatching in the following way: A power interconnection interface is configured in the docking mechanism to form a unified power supply bus, establishing a direct power transmission path between satellites. Relevant electrical parameters, such as voltage and current values, are collected from the concentrated energy systems of each satellite using local power collectors. The power dispatch manager generates dispatch instructions based on the electrical parameter data collected by multiple local power collectors, controlling the power switch array to open or close the power transmission switches at the corresponding satellite access points, thus enabling cross-satellite power transmission from satellites with surplus computing power to satellites with insufficient computing power.

[0048] This scheduling mechanism integrates the independent energy systems of each satellite into an interconnected power grid, making the surplus power of any satellite a shared resource of the system. This helps to mitigate individual energy differences caused by orbital position, local faults, or load fluctuations, and optimizes energy allocation at the system level.

[0049] In summary, this embodiment achieves energy dispatching based on physical connection through the cooperation of power interconnection interface, power supply bus and inter-satellite energy convergence system.

[0050] In one embodiment, the power scheduling manager is deployed on multiple computing satellites, and the cross-satellite energy convergence system adopts a distributed architecture with master-slave backup. In this architecture, one of the multiple computing satellites dynamically elects a master node to run the power scheduling manager; the master node further designates another computing satellite as a hot backup node and keeps its state synchronized with the master node. When the master node fails, the hot backup node automatically becomes the new master node and continues to perform energy scheduling functions.

[0051] This master-slave backup architecture effectively avoids single points of failure in the system. Even if the master node fails, the energy scheduling and management functions can continue to operate, thereby ensuring uninterrupted cross-satellite energy scheduling and maintaining the overall stability of the computing center.

[0052] The dynamic election mechanism enables the system to flexibly select the most suitable node as the master node based on the health status, computing load, or energy status of each satellite, avoiding the long-term fixation of control responsibilities on a node that may have degraded performance or be in a poor position, thereby improving the system's adaptability and resource utilization efficiency.

[0053] The state synchronization mechanism between the hot backup node and the master node ensures that in the event of a master-slave switchover, the new master node can immediately assume scheduling responsibilities based on the latest system state, without undergoing a lengthy data reconstruction process. This significantly shortens fault recovery time and reduces the risk of scheduling interruptions or decision-making errors caused by state inconsistencies.

[0054] In summary, the embodiments of the present invention combine dynamic election, master-slave backup, state synchronization and automatic fault switching to construct a distributed energy management system with fault self-recovery and operational self-adaptation capabilities, which enhances the reliability of long-term on-orbit operation of the space computing center and provides support for system scaling.

[0055] In one embodiment, each computing satellite is equipped with a standardized on-orbit maintenance interface; the space computing center is also equipped with one or more space operation robots for performing on-orbit maintenance tasks of the computing satellites through the standardized on-orbit maintenance interface.

[0056] Standardized on-orbit maintenance interfaces include cabinet power supply interfaces, liquid cooling interfaces, and communication interfaces. Space-based operational robots can perform one or more of the following on-orbit maintenance tasks: power-off isolation and replacement of faulty computing satellites; replenishment or sealing of fluid loop systems via the liquid cooling interface; and software injection or system debugging via the communication interface.

[0057] By configuring standardized on-orbit maintenance interfaces on each computing satellite, a unified physical docking and operational benchmark is provided for space operation robots, solving the problem of difficult on-orbit service caused by the lack of dedicated maintenance interfaces for traditional satellites, and establishing the infrastructure for implementing systematic maintenance.

[0058] Standardized interfaces ensure consistency in maintenance operations across different satellites, enabling space operations robots to maintain multiple satellites using the same tools and procedures. This standardized design improves maintenance efficiency, reduces operational complexity, and provides technical assurance for the sustainable operation of large-scale satellite arrays.

[0059] By using space-operated robots to perform on-orbit maintenance tasks, the system can complete operations such as component replacement and interface maintenance without interrupting power or service. This capability effectively extends the on-orbit lifespan of individual satellites and the entire system, and reduces system performance degradation or functional loss caused by component failures.

[0060] In summary, the embodiments of the present invention provide a systematic on-orbit maintenance capability for space computing centers through the collaborative design of standardized maintenance interfaces and space operation robots, supporting their long-term stable operation.

[0061] In one embodiment, see Figure 2 The concentrated solar power system includes a power generation unit, an energy storage unit, and a power control unit. The power generation unit uses a deployable concentrated solar array to power the computing payload and platform equipment during the sunshine period and to charge the energy storage unit. The energy storage unit powers the computing payload and platform equipment during the shadow period. The power control unit is electrically connected between the power generation unit and the energy storage unit and is configured to control the energy transfer between the power generation unit and the energy storage unit according to the satellite's orbital phase and dynamically allocate the power supply to the computing payload and platform equipment.

[0062] The concentrated solar power system, through the coordinated operation of three functional units—power generation, energy storage, and power control—constructs an energy supply architecture adapted to the periodic changes in sunlight during the day and night orbit. The power generation unit utilizes a deployable concentrated solar array, which improves power generation efficiency per unit area through concentrating technology. Furthermore, during launch, its foldable structure meets envelope constraints, helping to balance high power consumption requirements with the satellite platform's mass and volume constraints. The energy storage unit refers to devices used to store electrical energy, such as lithium-ion battery packs or supercapacitor banks, to store surplus energy during the sunlight period, providing energy support for operation during the shadow period.

[0063] The power control unit, as the core of the system's management, implements energy transfer management and dynamic power allocation by monitoring the output power of the power generation unit, the state of charge of the energy storage unit, and the load power consumption requirements in real time. Its core functions are mainly reflected in two aspects:

[0064] ① In terms of energy transfer management, the power control unit adaptively controls the energy transfer between the power generation unit and the energy storage unit based on the satellite orbit phase. For example, the energy storage unit is charged at the beginning of the solar radiation period, and the charging strategy is optimized at the end based on the predicted length of the shadow period; during the shadow period, the discharge parameters are dynamically adjusted based on the remaining power of the energy storage unit. This approach helps to extend the lifespan of the energy storage unit and maintain the continuity of energy supply.

[0065] ② Regarding dynamic power allocation, the power control unit manages the power supply to computing loads and platform equipment according to preset priorities based on rail phase and system status. During periods of energy shortage, priority is given to ensuring the platform's basic functions and critical computing tasks, while non-critical loads are operated at reduced capacity, thereby reducing the risk of power outages, achieving system-level on-demand energy allocation, and improving the flexibility of task execution.

[0066] In actual operation, during the sunshine period, the power generation unit converts solar energy into electrical energy, giving priority to powering the load and equipment, and the remaining electrical energy charges the energy storage unit; after entering the shadow period, the system switches to the energy storage unit discharge mode and dynamically adjusts the power supply allocation according to the task priority; when the energy storage unit's state of charge is lower than the threshold, the power consumption of non-core equipment is automatically limited to ensure the operation of the platform's basic functions.

[0067] In summary, this invention establishes an energy supply system that integrates power generation, energy storage, and a control unit by decomposing the architecture of a concentrated solar power system. The power generation unit balances power demand and volume constraints through concentrated solar power and deployable design; the energy storage unit provides energy support during the shadow period; and the power control unit achieves optimized scheduling and allocation of energy through dynamic management, ensuring the continuous operation of the computing load throughout the entire orbital cycle.

[0068] In one embodiment, see Figures 3-5 As shown, after the concentrating solar array is deployed in orbit, each solar panel includes at least one solar array panel 20. The solar array panel includes a thin-walled carbon fiber tube frame 21 and m×m concentrating arrays 22 disposed on the frame. Each concentrating array consists of n×n concentrating modules 23, where m and n are both integers ≥2.

[0069] The design establishes a three-tiered modular structure consisting of solar array panels 20, concentrator array 22, and concentrator modules 23. This hierarchical structure supports large-scale system configuration, allows for adjustment of solar array size according to mission requirements, and facilitates mass production and integration of the product.

[0070] The solar array panel 20 uses a thin-walled carbon fiber tube frame 21 as its main support structure, which ensures the necessary stiffness and strength while contributing to the lightweight design of the system. This thin-walled carbon fiber tube frame 21 utilizes the high specific strength and high specific stiffness characteristics of carbon fiber composite materials to provide the necessary support stiffness for the entire panel while achieving structural lightweighting, enabling it to withstand the mechanical loads during launch and providing a stable mounting reference for the concentrator array 22 after on-orbit deployment.

[0071] The focusing array 22 adopts a distributed modular design. When a single focusing module 23 or a local array fails due to spatial environmental factors, the impact can be controlled locally, avoiding the loss of overall system functionality and thus improving the system's fault tolerance. At the same time, the standardized array arrangement provides a consistent working environment for each focusing module, which is conducive to achieving uniform energy collection across the entire light-collecting surface.

[0072] In one specific implementation example, after the concentrating solar array is deployed in orbit, it forms a pair of symmetrically arranged solar panels on both sides of the satellite platform 10. Each solar panel includes two solar array panels 20, and each solar array panel 20 can contain 16 concentrating arrays 22 in a 4×4 pattern. Each concentrating array 22 can be composed of 25 concentrating modules 23 in a 5×5 pattern. During orbital illumination, the concentrating solar array converts solar energy into electrical energy through its distributed concentrating modules, and after accumulator regulation, forms a power output that meets the load requirements.

[0073] In summary, the embodiments of the present invention have achieved a multi-level modular design of a concentrating solar cell array through the above technical solutions, which improves the power redundancy and operational reliability of the energy system while ensuring a lightweight structure.

[0074] In one embodiment, see Figure 6 As shown, the focusing module 23 includes a primary mirror 1, a secondary mirror 3, a light funnel 6, a solar cell 5, and a heat spreader 7. The primary mirror 1 includes a primary mirror base 101, which forms the sun-facing opening of the focusing module 23. Its optical working surface faces the secondary mirror 3, and it is used to receive and reflect incident sunlight. The secondary mirror 3 is fixed above the primary mirror 1 by a secondary mirror bracket 4 and is located in the converging area of ​​the reflected light path of the primary mirror 1. Its optical working surface faces the entrance of the light funnel 6, and it is used to perform secondary reflection of the light reflected by the primary mirror 1. The light funnel 6 has a conical cavity structure. Its entrance receives the reflected light path of the secondary mirror 3, and its exit corresponds to the light-receiving surface of the solar cell 5, and it is used to guide the secondary reflected light from the secondary mirror 3 to the solar cell 5. The solar cell 5 is located below the light path exit of the light funnel 6 and is used to convert light energy into electrical energy. The heat-dissipating plate 7 is tightly attached between the battery cell 5 and the main mirror substrate 101 to conduct the heat generated by the battery cell 5 to the main mirror substrate 101, and then the main mirror substrate 101 radiates heat to the space environment.

[0075] In the concentrating module 23, the various optical components cooperate spatially to form a complete concentrated solar power generation unit. During operation, sunlight is transmitted sequentially through the primary mirror 1, secondary mirror 3, and light funnel 6, ultimately forming a high-energy-density light spot on the surface of the solar cell 5. After absorbing photons, the solar cell 5 generates photogenerated carriers, forming an effective photocurrent and achieving photoelectric conversion. During the orbital illumination period, multiple concentrating modules 23 distributed on both sides of the solar array work together, and the electrical energy generated by each module is collected and combined through a cable network to form a power output that meets the requirements.

[0076] The concentrating module 23 adopts a secondary reflection optical path design. The primary mirror 1 acts as a light collector, reflecting and converging sunlight to the secondary mirror 3. The secondary mirror 3 then shapes the beam and guides the light to the light funnel 6. This optical path structure helps to improve the concentrating efficiency, achieving higher energy density while reducing the area of ​​the required semiconductor cells.

[0077] The conical cavity structure of the light funnel 6 has a certain ability to adapt to small changes in the solar incidence angle. When sunlight is incident perpendicularly, the reflected light directly illuminates the solar cells; when the incidence angle is slightly deviated, some light can be reflected through the conical sidewalls and directed to the central region of the solar cells, which helps to enhance the system's adaptability to on-orbit attitude fluctuations.

[0078] The heat spreader 7, as a key heat-conducting element, is tightly bonded between the solar cell 5 and the primary mirror substrate 101, constructing a heat conduction path from the high-density point heat source of the solar cell 5 to the large-area heat sink of the primary mirror substrate 101. The heat spreader 7 reduces the local heat flux density of the solar cell through lateral diffusion, helping to reduce the risk of localized high temperatures and ensuring the solar cell operates within a safe temperature range, while simultaneously promoting heat transfer to the primary mirror substrate. Ultimately, the heat is radiated into space through the surface of the primary mirror. This integration of optical and heat dissipation functions avoids the need for a separate, bulky heat dissipation component for the solar cell, contributing to a more compact and lightweight structure.

[0079] In summary, the embodiments of the present invention achieve synergistic performance of the focusing module in terms of light energy collection, heat dissipation, and environmental adaptability through the integrated design of optical path structure optimization and thermal management path.

[0080] In one embodiment, see still Figure 6 The primary mirror substrate 101 adopts a near-parabolic spherical structure with a quadrilateral outline. The near-parabolic spherical surface reflects and converges incident sunlight to a predetermined area, providing the optical basis for light concentration. The quadrilateral outline provides a structural boundary for modular integration, enabling multiple concentrating modules to be arranged in a matrix on the surface of the solar array.

[0081] The quadrilateral outer shape facilitates stable connections with adjacent modules and support frames, enhancing overall structural rigidity. Furthermore, this configuration is adaptable to manufacturing processes, facilitating processing, assembly, and testing. By employing a near-parabolic spherical structure with a quadrilateral outer shape, the primary mirror substrate 101 achieves optical converging functionality while supporting modular integration and structural load-bearing capacity.

[0082] The concave surface of the primary mirror substrate serves as the optical working surface, and a reflective layer 102 and a transparent infrared radiation layer 103 covering the reflective layer 102 are sequentially provided. The reflective layer 102 provides high reflectivity to sunlight, while the transparent infrared radiation layer 103, while maintaining optical transmittance, can dissipate the heat absorbed by the primary mirror substrate 101 into the space environment in the form of infrared radiation.

[0083] A thermally conductive radiation layer 104 is provided on the convex surface of the primary mirror substrate. Utilizing its high in-plane thermal conductivity, this thermally conductive radiation layer 104 can diffuse the heat absorbed by the primary mirror substrate, especially the localized heat flow conducted from the solar cells, along the plane of the primary mirror substrate 101. Simultaneously, the thermally conductive radiation layer 104 also serves as a radiative heat dissipation surface, radiating heat into the space environment in the form of infrared radiation.

[0084] Through the synergistic effect of the concave and convex functional layers, a double-sided heat dissipation mechanism is formed in the primary mirror substrate 101. The transparent infrared radiation layer 103 on the concave side and the thermally conductive radiation layer 104 on the convex side together constitute the radiation heat dissipation surface, increasing the effective heat dissipation area. This design helps improve the temperature distribution uniformity of the primary mirror substrate, reduces the operating temperature, and provides support for maintaining optical surface accuracy.

[0085] During operation, sunlight passes through the concave transparent infrared radiation layer 103 and is reflected by the reflective layer 102 to the secondary mirror 3. The heat generated by the solar cell 5 during operation is conducted to the primary mirror substrate 101 through the heat spreader 7. Part of this heat is directly radiated and dissipated through the concave transparent infrared radiation layer 103, while the other part is radiated and dissipated laterally through the convex heat-conducting radiation layer 104. This dual-path heat dissipation mechanism helps maintain the thermal stability of the primary mirror substrate.

[0086] It should be noted that the reflective layer 102 is preferably a silver-plated layer. Silver has a high reflectivity (typically >95%) in the visible to near-infrared band, which can reflect most of the solar energy to the secondary mirror, while reducing the absorption of light energy by the primary mirror substrate 101, thus helping to reduce the heat generation of the primary mirror substrate itself.

[0087] The transparent infrared radiation layer 103 is a functional coating deposited on the reflective layer 102, which has high visible light transmittance and infrared emissivity. This coating not only provides space environment protection for the reflective layer 102 and slows down its performance degradation in the space environment, but also effectively dissipates the heat absorbed by the primary mirror substrate through infrared radiation, so that the concave surface has both high reflectivity and efficient heat dissipation functions.

[0088] The thermally conductive radiation layer 104 is preferably a graphene film layer. This graphene film is a two-dimensional nanomaterial composed of carbon atoms and possesses the following excellent properties: ① The graphene film has high in-plane thermal conductivity (>1500 W / m·K), allowing it to work in conjunction with the heat spreader to rapidly diffuse the high-density heat generated by the solar cell and laterally conducted through the heat spreader to the entire primary mirror substrate, thus achieving uniform temperature, helping to avoid localized overheating, ensuring the safety of the solar cell, and providing a large surface area for heat radiation into space. ② The graphene film itself is also an excellent infrared radiation material with a certain infrared emissivity (>0.9), allowing it to directly radiate some heat. ③ The graphene film has a thickness on the micrometer scale and its mass is almost negligible, having a minimal impact on the system's weight gain. ④ The flexible nature of the graphene film allows it to adapt to thermal deformation under temperature changes, avoiding interfacial delamination caused by thermal stress.

[0089] Therefore, the concave surface of the primary mirror substrate, composed of a reflective layer 102 and a transparent infrared radiating layer 103, possesses both high reflectivity and infrared emissivity, facilitating the integration of the optical surface and the radiative heat dissipation surface. The thermally conductive radiating layer 104 on the convex surface of the primary mirror substrate works in conjunction with the heat spreader 7 to construct a low thermal resistance heat conduction path from the battery cell 5 to the primary mirror substrate 101. The specific heat transfer process includes:

[0090] ① First-stage lateral heat diffusion: The high-density heat flow generated by the battery cell 5 is first absorbed by the heat diffusion plate 7 attached to it and diffused laterally to reduce the heat flow density and avoid the formation of local hot spots under the battery cell 5.

[0091] ② Secondary lateral temperature uniformity: After the heat is conducted to the primary mirror substrate 101 through the heat spreader 7, the heat-conducting radiation layer 104 attached to the convex surface of the primary mirror substrate uses its ultra-high in-plane thermal conductivity to rapidly diffuse the heat laterally to the entire primary mirror substrate to achieve temperature uniformity.

[0092] The heat is ultimately radiated into the space environment through the convex and concave dual surfaces of the primary mirror substrate 101, forming an efficient heat dissipation path for the solar cell from "solar cell → heat spreader → primary mirror substrate (convex surface + concave surface) → space environment". See [link to solar cell heat dissipation path] for details. Figure 7 As shown.

[0093] This dual-layer design of the primary lens allows it to perform its light-gathering function while also possessing good heat dissipation capabilities. (See also...) Figure 8 As shown, when the focusing module is working, sunlight penetrates the transparent infrared radiation layer 103, is reflected by the reflective layer 102 to the secondary mirror 3, and is then reflected a second time by the secondary mirror 3 to the solar cell 5. The high-density heat flow generated by the photoelectric conversion of the solar cell 5 is first absorbed by the heat spreader 7 immediately below it and undergoes primary lateral diffusion. Then, the heat is conducted to the primary mirror substrate 101 via the heat spreader 7. The solar radiation heat absorbed by the primary mirror substrate 101 itself and the heat conducted from the solar cell 5 are dissipated through two paths: one is direct outward thermal radiation through the concave transparent infrared radiation layer 103 (thermal path B1); the other is outward thermal radiation through the convex thermal conductive radiation layer 104 after secondary lateral temperature equalization (thermal path B2). This dual-sided heat dissipation collaborative design of "concave reflection + heat dissipation, convex thermal conduction + heat dissipation" improves heat dissipation efficiency while ensuring the optical performance of the primary mirror, helps to suppress the thermal deformation of the primary mirror, and ensures the stability of the focusing optical path.

[0094] In summary, the primary mirror substrate of this invention integrates effective thermal management capabilities while achieving the function of focusing and reflecting light. The combination of the concave reflective layer and the transparent infrared radiating layer provides heat dissipation while ensuring reflective performance, while the convex thermally conductive radiating layer enhances lateral heat conduction and radiative heat dissipation capabilities. This integrated design avoids the use of additional heat dissipation components, contributing to a compact and lightweight structure and providing a technical foundation for the stable operation of the focusing module.

[0095] In one embodiment, the secondary mirror has a near-hyperboloid surface structure, with a reflective medium film on its optical working surface and a heat dissipation coating on its non-optical working surface.

[0096] The secondary mirror adopts a near-hyperboloid design, forming a synergistic optical system with the primary mirror's near-parabolic spherical surface. This effectively corrects the optical path, improves beam converging effect, and helps optimize beam quality, forming a suitable light spot on the surface of the solar cell.

[0097] The reflective medium film on the optical working surface of the secondary mirror is a multilayer interference thin film prepared through processes such as vacuum deposition. This reflective medium film is optimized for the solar spectrum and has high reflectivity, which can reduce the loss of light energy during the reflection process of the secondary mirror.

[0098] The heat dissipation coating on the non-optical working surface of the secondary mirror uses materials with high infrared emissivity and low solar absorptivity, such as aerospace-grade inorganic white paint. Although the reflective dielectric film on its optical working surface significantly reduces light absorption, the secondary mirror will still heat up due to external heat flow and residual energy absorption. This heat dissipation coating enhances the infrared emissivity of the non-optical working surface of the secondary mirror, promoting the dissipation of accumulated heat from the secondary mirror into the space environment.

[0099] In summary, the embodiments of the present invention, by adopting a near-hyperbolic surface shape and combining it with a double-sided functional coating design, achieve beam conduction while also taking into account the requirements for operating temperature control, which helps to maintain the optical performance stability of the focusing module.

[0100] In one embodiment, see still Figure 6 The primary mirror base 101 has mechanical mounting surfaces 41 at its four corners. The secondary mirror support 4 adopts a long, rod-shaped structure and has a hollowed-out portion 42. The two ends of the support are fixedly connected to two mounting surfaces on a pair of diagonals of the primary mirror base 101.

[0101] The mounting planes at the four vertices of the quadrilateral outline of the primary mirror base 101 provide defined mounting interfaces for the secondary mirror bracket 4. Fixing both ends of the secondary mirror bracket to the diagonal mounting planes creates a support structure with a large span. The hollow design of the secondary mirror bracket 4 helps reduce obstruction of the incident and reflected light paths while maintaining structural functionality.

[0102] The secondary mirror 3 and the secondary mirror support 4 are secured together by a threaded connector with an anti-loosening structure. This connection method can suppress loosening of the threaded pair that may be caused by emission vibration and on-orbit temperature changes, maintaining the stability of the connection. The threaded connection allows for appropriate adjustment of the angle of the secondary mirror during assembly to meet optical alignment requirements.

[0103] In summary, the embodiments of the present invention provide a stable support structure for the secondary mirror through the combination design of the mounting plane, the bracket with the hollow part, and the anti-loosening threaded connection, while providing solutions for reducing light shading and maintaining alignment accuracy.

[0104] In one embodiment, see Figure 9 As shown, the present invention further proposes a space pump-driven fluid loop system, including: a radiant cooling plate assembly, a drive assembly, and a loop controller. The radiant cooling plate assembly and the drive assembly form a closed fluid loop through a fluid pipeline. The loop controller is electrically connected to the radiant cooling plate assembly and the drive assembly, respectively.

[0105] The radiant cooling plate assembly is a deployable flexible radiant cooling plate used to radiate the heat carried by the working fluid flowing through its internal fluid pipes into space.

[0106] The drive assembly includes a circulation pump for providing pump drive power to the fluid circuit and driving the working fluid to circulate within the fluid pipeline.

[0107] The loop controller is used to monitor the operating status of the fluid loop system, and control the deployment status of the cooling plate assembly and adjust the operating parameters of the drive assembly based on the operating status; wherein, the operating status includes temperature, pressure, flow rate, equipment speed and valve position status information in the system.

[0108] The space pump-driven fluid loop system of this invention adopts a modular architecture, with three core components functionally interconnected through standardized interfaces. In operation, the working fluid flows along a closed loop driven by a circulating pump, sequentially passing through the heating device and the radiant cooling plate assembly, completing a "drive-absorption-radiation-recirculation" cycle. This architecture clearly defines the functional division of each module, providing a foundation for coordinated system operation.

[0109] The design of the deployable flexible radiant cooling plate allows it to be folded and retracted during the launch phase to adapt to the envelope constraints of the launch vehicle; after being deployed in orbit, it forms a large area of ​​radiative heat dissipation surface, providing the necessary conditions for heat dissipation.

[0110] The loop controller monitors system operating parameters in real time to adjust parameters such as the circulating pump speed and control the deployment status of the cooling plates. This monitoring mechanism enables the system to adjust its operating status according to environmental changes and load requirements.

[0111] In summary, the embodiments of the present invention provide a corresponding solution for the thermal management of high-power space equipment by combining a modular system architecture, deployable heat dissipation surface design, and status monitoring functions. The modular architecture supports the separate development and testing of the system, the deployable design balances launch constraints and on-orbit functional requirements, and the status monitoring function helps maintain the system's operational stability under different operating conditions.

[0112] In one embodiment, the space pump-driven fluid loop system is further equipped with a heating element. This heating element is in the form of thin-film heating plates distributed on the outer surface of the components in the fluid loop system that require temperature control. A loop controller is electrically connected to the heating element and controls its on / off state and power based on the monitored temperature.

[0113] Thin-film heating elements are characterized by their thinness and good adhesion, and can be arranged according to the shape of different mounting surfaces. In fluid system layouts, heating components are mainly distributed in key parts of the fluid circuit: in radiant cooling plate assemblies, heating elements are axially bonded or wrapped around the surface of metal flexible hoses along the rigid fluid pipeline; in drive assemblies, they are attached to the housings of equipment such as circulating pumps, pump drivers, and liquid receivers.

[0114] The loop controller drives the heating element via a switching circuit and, together with a temperature sensor, forms a temperature regulation system. This system can be configured with redundancy to improve control reliability.

[0115] The installation of heating components can bring the following effects to the fluid loop system:

[0116] 1) Under low-temperature conditions, by directional heating of pipelines and key equipment, the working fluid state can be maintained, preventing operation interruption or component damage caused by freezing, improving the system's operational stability under extreme temperature conditions, and enhancing its adaptability to low-temperature environments.

[0117] 2) Implementing on-demand heating based on real-time temperature feedback avoids continuous full-power operation and helps achieve a balance between precise temperature control and energy consumption.

[0118] 3) The heating control function is integrated into the loop controller, realizing unified management of the main heat dissipation loop and auxiliary heating, and enhancing the system's thermal regulation capability under different operating conditions.

[0119] In summary, by incorporating a heating component, this embodiment of the invention provides the system with corresponding low-temperature protection, which helps maintain the system's operational stability under different temperature environments.

[0120] In one embodiment, see Figure 10 As shown, the deployable flexible radiant cooling plate includes at least one radiant cooling plate 30, which is folded and retracted during the launch phase and unfolded to the working position during the on-orbit phase.

[0121] Each radiant cooling plate 30 comprises three main components: a lightweight, high-strength frame 32, multiple parallel rigid fluid conduits 31 integrated on the frame, and a flexible radiant membrane 33 covering the outer surface of the rigid fluid conduits 31. The function of the flexible radiant membrane 33 is to radiate the heat carried by the working fluid flowing through the rigid fluid conduits 31 into space.

[0122] The design employs a deployable modular configuration. The radiant cooling plate 30 serves as the basic functional unit, which can be folded and retracted during the launch phase to adapt to the envelope constraint, and deployed during the on-orbit phase to obtain the required heat dissipation area.

[0123] In terms of structural design, a composite structure combining rigidity and flexibility is adopted. The lightweight, high-strength frame 32 serves as the main load-bearing structure, providing structural support for the rigid fluid pipelines 31 integrated on it. Multiple parallel rigid fluid pipelines 31 constitute a fluid channel system, and a flexible radiating membrane 33 serves as the main heat dissipation surface covering the outer surface of the pipelines.

[0124] In terms of heat dissipation path, a heat transfer channel is established from the working fluid to space. The rigid fluid pipeline 31 undertakes the functions of working fluid transportation and basic heat conduction, while the flexible radiant membrane 33 realizes the lateral diffusion of heat and completes heat dissipation through its surface radiation.

[0125] Each radiant cooling plate (30mm) employs a parallel flow channel layout, including a main inlet pipe, a main return pipe, and multiple parallel branch pipes. This structure helps achieve uniform distribution of the working fluid, improving heat dissipation efficiency and temperature uniformity. The parallel flow channel design can maintain some heat dissipation capacity even when a single pipe experiences localized abnormalities, while also helping to reduce fluid flow resistance.

[0126] The lightweight, high-strength frame 32 provides stable structural support for the rigid fluid pipeline 31 and the flexible radiation membrane 33. Preferably, the lightweight, high-strength frame 32 is made of materials such as carbon fiber composite materials or titanium alloy honeycomb sandwich structures, which can achieve lightweighting of single panels while ensuring structural performance. This frame can withstand mechanical loads during the launch phase, maintain the dimensional and positional stability of the heat dissipation surface during the on-orbit phase, and provide support for the positioning and fixation of the pipeline.

[0127] By integrating the rigid fluid pipeline 31, the flexible radiation membrane 33, and the lightweight high-strength frame 32 into a modular heat dissipation unit, the system assembly process is simplified, and a structural basis is provided to adapt to different heat dissipation requirements.

[0128] In summary, the embodiments of the present invention provide corresponding solutions in terms of emission constraint adaptability, structural stability and thermal management performance through the design of deployable modular configuration, rigid-flexible composite structure and parallel flow channel heat transfer path.

[0129] In one embodiment, the deployable flexible radiant cooling plate includes a plurality of odd-numbered radiant cooling plate panels, wherein the central radiant cooling plate panel is fixedly connected to the satellite platform; during the launch phase, all radiant cooling plate panels are folded and compressed and fixed to the satellite platform; during the on-orbit phase, after all radiant cooling plate panels are deployed, they are arranged symmetrically with the central radiant cooling plate panel as the axis of symmetry and remain coplanar.

[0130] See Figure 11 As shown, 30 represents a single radiant cooling plate, and 13 represents a flexible radiant cooling plate, referring here to all the radiant cooling plate panels compressed into a single unit. During the initial launch phase, all the radiant cooling plate panels are compressed into a single unit by multiple separation nuts and folded into the satellite platform 10. During in-orbit deployment, all the radiant cooling plate panels, initially folded, are first rotated 90° by a motor to be perpendicular to the satellite platform 10, then rotated 180° to reach a predetermined position parallel to the satellite platform 10. The separation nuts are then energized to unlock, allowing the radiant cooling plate panels to separate from each other. The deployment hinges provide the deployment torque to unfold the panels. After in-orbit deployment, all the radiant cooling plate panels are symmetrically arranged and remain coplanar.

[0131] During launch, all 30-fold radiated cooling plates are folded and stowed within the satellite platform, forming a stable force transmission path based on the satellite platform 10. This design helps resist vibration and shock during launch, reduces harmful relative motion or deformation within the radiated cooling plate assembly, and improves the mechanical environmental adaptability during launch.

[0132] After deployment in orbit, all the radiating cooling plates 30 are arranged symmetrically and remain coplanar. The symmetrical configuration helps control changes in the center of mass and reduces interference with the satellite's attitude; the coplanar arrangement ensures that all heat dissipation surfaces effectively face the cold black space, avoiding mutual shading and thus improving heat dissipation efficiency. In addition, the coplanar structure has high overall stiffness, which helps suppress the effects of on-orbit micro-vibrations and thermal deformation, forming a stable heat dissipation platform.

[0133] The layout employs an odd number of deployable radial cooling plates, which fold and retract to accommodate the envelope constraints of the rocket fairing. Furthermore, the deployed heat dissipation area during the on-orbit phase is significantly larger than that of the satellite platform, thus providing a structural basis for handling kilowatt-level heat dissipation. The symmetrical layout also promotes uniform fluid distribution within the piping, reducing the risk of localized overheating or uneven heat dissipation, further enhancing the system's heat dissipation reliability and temperature field uniformity. Moreover, using an odd number of plates (such as three) facilitates the construction of a stable symmetrical structure, avoiding the configuration uncertainties and hinge design complexity that may arise from an even number of plates.

[0134] In summary, the embodiments of the present invention, by limiting the number of blocks, fixing the center, and using a symmetrical and coplanar unfolding configuration, form a coordinated design scheme in terms of structural stability, unfolding reliability, heat dissipation efficiency, and system scalability.

[0135] The above are merely preferred embodiments of the present invention and are not intended to limit the scope of protection of the present invention. For those skilled in the art, the present invention can have various modifications and variations. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A super large planar array type space computing power center, characterized in that, Deployed in dawn-dusk orbit, a plurality of computing satellites are physically docked on orbit, including a plurality of computing satellites; Each of the computing satellites comprises a satellite platform, and an computing load, a light-concentrating energy system and a space-pumped fluid loop system arranged on the satellite platform, and the satellite platform is equipped with a docking mechanism for inter-satellite physical docking on orbit, and the docking mechanism is configured with an electrical interconnection interface; When the plurality of computing satellites are physically docked through the docking mechanism, the electrical interconnection interfaces of the computing satellites are electrically connected to each other to form an inter-satellite power supply bus, and the expandable light-concentrating solar cell array of the light-concentrating energy system and the expandable flexible radiating plate of the space-pumped fluid loop system are respectively deployed on orbit to form a super-large planar array type rigid whole. The space computing center is also configured with a cross-satellite energy bus system for energy scheduling between the light-concentrating energy systems based on the inter-satellite power supply bus; the cross-satellite energy bus system adopts a master-slave backup distributed architecture, comprising: A plurality of local power collectors are arranged on each computing satellite for collecting electrical parameters of the light-concentrating energy system of the satellite; A power scheduling manager is arranged on the plurality of computing satellites and is in communication connection with the plurality of local power collectors for generating scheduling instructions according to the electrical parameters; A power switch array is arranged on the inter-satellite power supply bus and is controlled by the power scheduling manager; the power scheduling manager realizes cross-satellite transmission of electrical energy from a surplus computing satellite to a short computing satellite by controlling the power switch array; Among them, the plurality of computing satellites where the power scheduling manager is arranged dynamically elect a computing satellite as a master node to run the power scheduling manager, and one of the other computing satellites is designated by the master node as a hot backup node to keep state synchronization with the master node; when the master node fails, the hot backup node automatically becomes a new master node to run the power scheduling manager.

2. The super-mesh array space computing power center according to claim 1, wherein, Each of the computing satellites is configured with a standardized on-orbit maintenance interface; The space computing center is also configured with one or more space operation robots for performing on-orbit maintenance tasks of the computing satellites through the standardized on-orbit maintenance interface.

3. The super-mesh array space computing power center according to claim 1, wherein, The light-concentrating energy system comprises a power generation unit, an energy storage unit and a power supply control unit; The power generation unit adopts the expandable light-concentrating solar cell array for supplying power to the computing load and platform equipment during the light period and charging the energy storage unit; The energy storage unit is used to supply power to the computing load and platform equipment during the eclipse period; The power supply control unit is electrically connected between the power generation unit and the energy storage unit, and is configured to control the energy transmission between the power generation unit and the energy storage unit according to the orbit phase of the satellite, and dynamically allocate the power supply power to the computing load and platform equipment.

4. The super-mesh array space staking center according to claim 1, wherein, The deployable concentrating solar cell array, after being deployed in orbit, forms a symmetrically arranged double-sided solar wing on both sides of the satellite platform; each side of the solar wing comprises at least one solar cell array panel, which comprises a thin-walled carbon fiber tube frame and m×m concentrating arrays arranged on the frame; each concentrating array is composed of n×n concentrating modules, wherein m and n are both integers greater than or equal to 2.

5. The super-mesh array space staking center according to claim 4, wherein, The concentrating module comprises a primary mirror, a secondary mirror, a light funnel, a cell sheet and a heat spreading sheet. The primary mirror comprises a primary mirror base, which forms an open sun-facing surface of the concentrating module, and the optical working surface thereof faces the secondary mirror for receiving and reflecting incident sunlight. The secondary mirror is fixed above the primary mirror by a secondary mirror support and located in the reflection light path convergence area of the primary mirror, and the optical working surface thereof faces the entrance of the light funnel for performing the second reflection of the light reflected by the primary mirror. The light funnel has a conical cavity structure, the entrance thereof receives the reflection light path of the secondary mirror, and the exit thereof corresponds to the light receiving surface of the cell sheet for guiding the light twice reflected by the secondary mirror to the cell sheet. The cell sheet is arranged below the light path exit of the light funnel for converting light energy into electrical energy. The heat spreading sheet is closely attached between the cell sheet and the primary mirror base for conducting the heat generated by the operation of the cell sheet to the primary mirror base, and the primary mirror base radiates the heat to the space environment.

6. The superplane array space staking center of claim 5, wherein, The primary mirror adopts a near-parabolic spherical structure with a quadrilateral outline. The concave surface thereof is provided with a reflective layer and a transparent infrared radiation layer covering the reflective layer in sequence, and the transparent infrared radiation layer is used for radiating the heat absorbed by the primary mirror base to the space environment in the form of infrared rays. The convex surface thereof is provided with a heat-conducting radiation layer for conducting the heat absorbed by the primary mirror base in the plane and radiating the heat to the space environment in the form of infrared rays.

7. The superplane array type space staking center according to claim 5, wherein, The surface structure of the secondary mirror is a near hyperboloid, the optical working surface thereof is provided with a reflective medium film, and the non-optical working surface thereof is provided with a heat dissipation coating.

8. The superplane array type space staking center according to claim 6, wherein, The primary mirror base is provided with mechanical mounting planes at four corners thereof; The secondary mirror support adopts a long strip-shaped rod structure, and is provided with a hollow part thereon; the two ends of the secondary mirror support are fixedly connected to two mounting planes on a pair of diagonal lines of the primary mirror base; The secondary mirror and the secondary mirror support are fixed by a threaded connecting piece with an anti-loosening structure.

9. The super-mesh array space staking center according to claim 1, wherein, The space pump-driven fluid circuit system comprises a radiative cooling plate assembly, a driving assembly and a circuit controller, the radiative cooling plate assembly and the driving assembly form a closed fluid circuit through fluid pipelines, and the circuit controller is electrically connected with the radiative cooling plate assembly and the driving assembly respectively; The radiative cooling plate assembly adopts the deployable flexible radiative cooling plate for radiating the heat carried by the working medium flowing through the internal fluid pipeline to space; The driving assembly comprises a circulating pump for providing pump driving power for the fluid circuit to push the working medium to circulate in the fluid pipeline. The loop controller is used for monitoring the running state of the fluid loop system, and controlling the unfolding state of the radiating panel assembly and adjusting the running parameters of the driving assembly based on the running state; wherein the running state includes temperature, pressure, flow, equipment rotating speed and valve position state information in the system.

10. The super-mesh array space staking center according to claim 9, wherein, The fluid loop system is also configured with a heating assembly in the form of a thin film heating sheet, which is distributed on the outer surface of the components in the fluid loop system that need temperature control; The loop controller is electrically connected with the heating assembly, and controls the on-off and power of the heating assembly based on the monitored temperature.

11. The super-mesh array space staking center according to claim 9, wherein, The deployable flexible radiating panel includes at least one radiating panel single panel, which is folded and stored in the launch phase, and is unfolded to the working position in the on-orbit phase. Each radiating panel single panel includes a lightweight high-strength frame, a plurality of parallel rigid fluid pipes integrated on the frame, and a flexible radiating film covering the outer surface of the rigid fluid pipes, which is used for radiating the heat carried by the working medium flowing through the rigid fluid pipes to space.

12. The super-mesh array space staking center according to claim 11, wherein, The deployable flexible radiating panel includes a plurality of odd-numbered radiating panel single panels, wherein the middle radiating panel single panel is fixedly connected with the satellite platform; all the radiating panel single panels are folded and stored and fixedly pressed on the satellite platform in the launch phase; after being unfolded, all the radiating panel single panels are symmetrically arranged and kept coplanar with the middle radiating panel single panel as the symmetric axis in the on-orbit phase.

Images