Outboard integrated power supply system applied to spacecraft and manufacturing and control method of outboard integrated power supply system

Abstract

The invention discloses an outboard integrated power supply system applied to a spacecraft and a manufacturing and control method thereof.The outboard integrated power supply system comprises a solar cell array, a winding drum mechanism, a high-heat-conduction shell, a heat dissipation cover plate, a storage battery pack and a power supply controller, and the solar cell array is a flexible rolling type solar wing and is wound around the winding drum mechanism; the winding drum mechanism is installed on the sunny side of the high-heat-conduction shell, the solar cell array is in a folded state in the satellite assembling stage, after the satellite is launched to a preset orbit, the solar cell array is in an unfolded power generation state and supplies power to the satellite, and the heat dissipation cover plate is installed on the sunny side of the high-heat-conduction shell in a matched mode. A preset mounting position for mounting the storage battery pack and the power supply controller is arranged in the high-heat-conductivity shell, and the storage battery pack and the power supply controller are mounted in the preset mounting position in parallel. The problems that a traditional distributed satellite power supply system occupies the space in a satellite cabin, the system integration and test complexity is high, and the thermal management efficiency is low can be effectively solved.

Description

Technical Field

[0001] This invention relates to the technical field of spacecraft power systems, and in particular to an integrated external power system for spacecraft and its manufacturing and control methods. Background Technology

[0002] In the field of spacecraft power systems, a distributed satellite power system with a "cabin-exterior" separation layout is typically adopted, where the battery pack and power controller are located inside the satellite cabin, while the solar array is deployed outside. This architecture has revealed a series of inherent technical drawbacks in long-term spacecraft design and application.

[0003] First, the battery pack and power controller occupy a significant amount of valuable space within the satellite cabin, severely limiting the installation volume of the payload. Furthermore, additional structural supports are needed to support and secure these internal devices, further increasing the satellite's structural complexity and weight. Second, system integration and testing are highly complex. The battery pack, power controller, and solar array—the three core components—are delivered and accepted as independent products. During the satellite assembly phase, extensive mechanical installation, electrical cable connections, and interface matching tests are required. This discrete integration process is complex, time-consuming, and costly, and the numerous interfaces introduce a higher potential risk of failure. Finally, thermal management is inefficient. The large amount of waste heat generated by the battery pack and power controller during operation is initially released into the satellite cabin. The cabin thermal control system must first collect the heat, then conduct it to the cabin walls via heat pipes, and finally dissipate it into space through radiators. This long, multi-stage, and inefficient process necessitates a heavier and more power-consuming thermal control system to maintain the equipment within a suitable temperature range, becoming a significant burden on the system.

[0004] In summary, traditional distributed satellite power systems have many problems in terms of space occupation, system integration and testing, and thermal management. There is an urgent need for a new technical solution to solve these drawbacks and improve the overall performance and reliability of spacecraft power systems. Summary of the Invention

[0005] The purpose of this invention is to overcome the shortcomings of the prior art and provide an integrated external power system for spacecraft and its manufacturing and control methods, which can effectively solve the problems of traditional distributed satellite power systems, such as occupying internal space, high complexity of system integration and testing, and low thermal management efficiency.

[0006] The objective of this invention is achieved through the following technical solution:

[0007] An integrated extravehicular power system for spacecraft includes a solar array, a roll mechanism, a high thermal conductivity shell, a heat dissipation cover, a battery pack, and a power controller. The solar array is a flexible, rollable solar array wound around the roll mechanism, which is installed on the sun-facing side of the high thermal conductivity shell. During satellite assembly, the solar array is in a retracted state. After the satellite is launched and reaches its predetermined orbit, the solar array unfolds to generate power for the satellite. The heat dissipation cover is installed on the shaded side of the high thermal conductivity shell. The high thermal conductivity shell has pre-set mounting positions for the battery pack and power controller. The battery pack and power controller are installed side-by-side in the pre-set mounting positions, which effectively shortens the high-voltage, high-current transmission path. The power controller optimizes the electrical energy transmitted from the solar array, transmitting a portion of the energy to the battery pack and adjusting the other portion to a stable bus voltage, which is then output to the satellite cabin through the satellite's multi-functional transcabin interface. Simultaneously, the high thermal conductivity shell and heat dissipation cover directly dissipate the heat generated by the power controller and battery pack into the deep space environment through radiation.

[0008] Furthermore, the outer surface of the high thermal conductivity housing and heat dissipation cover is provided with a high emissivity thermal control white paint. The high emissivity thermal control white paint radiates heat outward during the illumination period and reduces the absorption of heat, thereby reducing the temperature rise of the battery pack and power controller.

[0009] Furthermore, the high emissivity thermal control white paint has a high hemispherical emissivity ε>0.85.

[0010] Furthermore, both the high thermal conductivity shell and the heat dissipation cover are made of high thermal conductivity carbon fiber composite material or metal matrix composite material.

[0011] Furthermore, thermally conductive mounting areas are respectively provided between the battery pack and the high thermal conductivity casing and heat dissipation cover. The thermally conductive mounting areas are provided with thermally conductive fillers or thermally conductive silicone pads to ensure good thermal contact. The remaining areas between the battery pack and the high thermal conductivity casing and heat dissipation cover, except for the thermally conductive mounting areas, are respectively covered with multi-layer thermal insulation components. The remaining areas between the battery pack and the high thermal conductivity casing and heat dissipation cover, except for the thermally conductive mounting areas, are thermally insulated by polyimide pads.

[0012] Furthermore, thermally conductive mounting areas are respectively provided between the power controller and the high thermal conductivity housing and the heat dissipation cover. The thermally conductive mounting areas are provided with thermally conductive fillers or thermally conductive silicone pads to ensure good thermal contact. The remaining areas between the power controller and the high thermal conductivity housing and the heat dissipation cover, except for the thermally conductive mounting areas, are respectively covered with multi-layer thermal insulation components. The remaining areas between the power controller and the high thermal conductivity housing and the heat dissipation cover, except for the thermally conductive mounting areas, are thermally insulated by polyimide pads.

[0013] Furthermore, the roll mechanism is provided with extension rods on both sides for deploying the solar cell array. The extension rods are made of superelastic memory composite material. One end of the extension rod is fixed to the sun-facing side of the high thermal conductivity shell, and the other end is wound to the end of the roll mechanism and fixed by a constraint structure. In the initial state, the extension rod stores elastic strain energy. After the satellite is launched and reaches the predetermined orbit, the constraint structure is released by the satellite or ground command, so that the extension rod releases the strain energy instantly and quickly springs back to the extended state, thereby driving the solar cell array to deploy smoothly.

[0014] Furthermore, the power controller includes a maximum power point tracking and charging regulator unit for optimizing the electrical energy output from the solar array, a bus control and regulation unit for regulating the electrical energy to a stable bus voltage, and a fault judgment module for determining whether there are solar array output short circuits, battery overvoltage / undervoltage, output overcurrent, or internal overheating operation faults based on the output of the solar array, the voltage / current / temperature of the battery pack, and the bus voltage parameters.

[0015] A method for manufacturing an integrated external power system for spacecraft, comprising the following steps:

[0016] The outer surfaces of the power controller and battery pack, except for the thermally conductive mounting area, are covered with multiple layers of thermal insulation components. The power controller and battery pack are then fixed to the preset mounting positions of the high thermal conductivity housing using polyimide pads, and all electrical connections are completed.

[0017] After the circuit layer of the solar cell array is laid, it is wound onto the roll mechanism, and then the roll mechanism is installed on the sun-facing side of the high thermal conductivity shell. Then the electrical interconnection between the solar cell array, the battery pack and the power controller is completed.

[0018] High emissivity thermal control white paint is sprayed or pasted on the outer surface of the high thermal conductivity shell and heat dissipation cover, and the heat dissipation cover is installed on the sun-facing side of the high thermal conductivity shell to complete the integrated assembly of the external power system.

[0019] Electromagnetic compatibility testing, vibration testing, and thermal vacuum testing were conducted sequentially on the external integrated power system to complete its fabrication.

[0020] A control method for an integrated external power system applied to a spacecraft includes the following steps:

[0021] The integrated external power system is installed as a whole at the pre-designated installation location outside the satellite cabin;

[0022] After the satellite is successfully launched and enters its predetermined orbit, the external integrated power system is in a power-off hibernation or safe standby mode.

[0023] The ground control center sends a power-on command to the integrated electronic system inside the satellite cabin via the uplink. The integrated electronic system sends a wake-up signal to the power controller outside the satellite cabin or directly connects the power supply switch via the command line of the multi-functional through-cabin interface. The solar array is deployed under ground control.

[0024] After powering on, the power controller performs a self-test. It controls the solar array to charge the battery pack. At the same time, the power controller continuously monitors the output of the solar array, the voltage / current / temperature of the battery pack, and the bus voltage parameters. Based on the monitored parameters, the power controller performs maximum power point tracking, battery charge and discharge management, and bus voltage regulation. It also reports the battery pack voltage, the internal temperature of the power controller, and the fault flag data to the integrated electronic system periodically or when there are changes through the telemetry channel of the multi-functional through-cabin interface.

[0025] After receiving data, the integrated electronic system adjusts the battery pack charging current threshold according to the overall satellite mission plan, and sends macro commands to the power controller through the remote control channel to enable the solar array to power the load and charge the battery pack during the sunshine period, and the battery to discharge and power the load during the shadow period, ultimately realizing the intelligent management mode of "on-board decision-making and off-board execution".

[0026] Compared with the prior art, the present invention has the following advantages and beneficial effects:

[0027] 1. Lightweighting and Space Liberation: By integrating high-power, high-heat-generating units such as battery packs and power controllers with the solar cell array on a common support outside the cabin, structural, thermal, and electrical integration is carried out outside the cabin to form an independent, pre-integrated "power system module". This directly and significantly reduces the weight of the satellite platform and completely frees up the space originally occupied inside the satellite cabin, allowing these resources to be used for payloads and directly improving the satellite's mission capabilities.

[0028] 2. Improved integration efficiency and reliability: Through modular design of the power system, all internal electrical connections and functional tests can be completed before the power system leaves the factory, which greatly simplifies the assembly, integration and testing process of the satellite platform, shortens the development cycle and improves reliability;

[0029] 3. Leap in thermal management performance: Make full use of the space environment for efficient thermal management. Through a high thermal conductivity structure, waste heat is directly radiated away. The heat dissipation path is extremely short and the efficiency is extremely high, which fundamentally solves the heat dissipation problem of equipment inside the cabin and simplifies the thermal control system. Attached Figure Description

[0030] Figure 1 This is a schematic diagram of the overall structure of the power supply system of the present invention.

[0031] Figure 2 This is an exploded view of the power supply system of the present invention.

[0032] Figure 3 This is a schematic diagram of the structure of the high thermal conductivity shell of the present invention.

[0033] Figure 4 This is a schematic diagram of the battery pack structure of the present invention.

[0034] Figure 5 This is a schematic diagram of the working principle of the power supply system of the present invention. Detailed Implementation

[0035] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are some embodiments of the present invention, but not all embodiments. 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.

[0036] Example 1:

[0037] like Figures 1 to 2 As shown, this embodiment provides an integrated external power system for spacecraft, including a solar array 1, a high thermal conductivity shell 2, a battery pack 3, a heat dissipation cover 4, a power controller 5, and a roll mechanism 6. The solar array 1 is a flexible rollable solar array, specifically a solar blanket covered with solar cells. The flexible rollable solar array is wound around the roll mechanism 6, which is installed on the sun-facing side of the high thermal conductivity shell 2. During satellite assembly, the solar array is in a retracted state. After the satellite is launched and reaches its predetermined orbit, the solar array is deployed to generate power for the satellite. The heat dissipation cover 4 is installed on the shaded side of the high thermal conductivity shell 2. The interior of body 2 is provided with a preset mounting position for the battery pack 3 and the power controller 5. The battery pack 3 and the power controller 5 are installed side by side in the preset mounting position, with the power controller 5 placed close to the battery pack 3. This effectively shortens the high-voltage and high-current transmission path. The power controller optimizes the power transmitted from the solar array and transmits part of the power to the battery pack. The other part of the power is regulated to a stable bus voltage and output to the satellite cabin through the satellite's multi-functional through-cabin interface. At the same time, the heat generated by the power controller and the battery pack is directly dissipated into the deep cold space environment in the form of radiation through the high thermal conductivity shell and heat dissipation cover.

[0038] In this embodiment, the outer surfaces of the high thermal conductivity shell and heat dissipation cover are provided with high emissivity thermal control white paint. The high emissivity thermal control white paint radiates heat outward and reduces heat absorption during the illumination period, thereby reducing the temperature rise of the battery pack and power controller. The high emissivity thermal control white paint has a high hemispherical emissivity ε>0.85.

[0039] like Figure 3 As shown, the high thermal conductivity shell serves as the main load-bearing structure and installation platform of the entire system. It provides installation points and structural support for the drum mechanism and solar cell array, as well as installation space and heat dissipation function for the battery pack and power controller. Both the high thermal conductivity shell and the heat dissipation cover are made of high thermal conductivity carbon fiber composite material or metal matrix composite material, which have the functions of installation, force transmission and heat dissipation.

[0040] like Figure 4 As shown, the battery pack is composed of multiple lithium-ion battery cells connected in series and parallel. Thermally conductive mounting areas are provided between the battery pack and the high thermal conductivity casing and heat dissipation cover. Thermally conductive mounting areas are provided with thermally conductive fillers or thermally conductive silicone pads 7 to ensure good thermal contact. The remaining areas between the battery pack and the high thermal conductivity casing and heat dissipation cover, except for the thermally conductive mounting areas, are covered with multi-layer thermal insulation components (not shown in the figure). The remaining areas between the battery pack and the high thermal conductivity casing and heat dissipation cover, except for the thermally conductive mounting areas, are thermally insulated by polyimide pads 8.

[0041] A thermally conductive mounting area is provided between the power controller and the high thermal conductivity housing and heat sink cover. The thermally conductive mounting area is provided with thermally conductive filler or thermally conductive silicone pad 9 to ensure good thermal contact. The remaining areas between the power controller and the high thermal conductivity housing and heat sink cover, except for the thermally conductive mounting area, are covered with multi-layer thermal insulation components (not shown in the figure). The remaining areas between the power controller and the high thermal conductivity housing and heat sink cover, except for the thermally conductive mounting area, are thermally insulated by polyimide pads 8.

[0042] The roll mechanism 6 has extension rods 10 on both sides for unfolding the solar cell array 1. The extension rods 10 are made of super-elastic memory composite material. One end of the extension rod 10 is fixed to the sun-facing side of the high thermal conductivity shell 2, and the other end is wound to the end of the roll mechanism 6 and fixed by a constraint structure (not shown in the figure). In the initial state, the extension rod stores elastic strain energy. After the satellite is launched and reaches the predetermined orbit, the constraint structure is released by the satellite or ground command, so that the extension rod releases the strain energy instantly and quickly springs back to the extended state, thereby driving the solar cell array to unfold smoothly.

[0043] The power controller includes a maximum power point tracking and charge regulator unit (MPPT and charge regulator) for optimizing the output power of the solar array, a bus control and regulation unit (bus control and regulator) for regulating the power to a stable bus voltage, and a fault diagnosis module for determining whether there are solar array output short circuits, battery overvoltage / undervoltage, output overcurrent, or internal overheating faults based on the output power of the solar array, the voltage / current / temperature of the battery pack, and the bus voltage parameters. The power controller uses radiation-resistant electronic components and incorporates redundant circuitry to improve reliability.

[0044] like Figure 5 As shown, the three critical paths of energy flow, data flow, and heat flow in the power system of this embodiment are as follows:

[0045] 1) Energy flow path:

[0046] Power generation: The solar cell array acts as a power generation unit, outputting the electrical energy it generates to the power controller.

[0047] Regulation and storage: The power controller optimizes the power energy through the maximum power point tracking and charging regulator unit. On the one hand, it charges and stores energy for the battery pack, and on the other hand, it regulates the power energy to a stable bus voltage through the bus control and regulation unit.

[0048] Power supply: The power controller outputs a stable bus power supply to the satellite cabin through the satellite's multi-functional through-cabin interface to power various payloads and platform equipment inside the satellite cabin.

[0049] 2) Data flow path:

[0050] Status monitoring: The operating status outside the satellite cabin (such as voltage, current, and temperature) is transmitted to the integrated electronic system inside the satellite cabin via telemetry signal lines in the multi-functional through-cabin interface.

[0051] Command control: The integrated electronic system generates control commands (such as power on / off and mode switching) based on the received status information, and then sends them to the power controller for execution through the telemetry channel of the same interface, thereby realizing closed-loop management of the power system.

[0052] 3) Heat flow path:

[0053] Heat generation: The power controller (electrical energy conversion) and the battery pack (electrochemical reaction) generate waste heat during operation.

[0054] Conduction: The generated waste heat is efficiently conducted to the highly thermally conductive shell and heat dissipation cover that are closely connected to it.

[0055] Radiation-induced heat dissipation: The high thermal conductivity shell and heat dissipation cover serve as large heat dissipation surfaces, directly dissipating the heat absorbed into the deep cold space environment in the form of radiation, thereby completing the entire thermal management cycle.

[0056] In addition, the present invention can also integrate the battery pack and the power controller into an external module, which is installed independently of the solar array and connected to it by a short cable, so as to free up internal space and improve heat dissipation.

[0057] Example 2:

[0058] This embodiment provides a method for manufacturing an integrated external power system for spacecraft, including the following steps:

[0059] The outer surfaces of the power controller and battery pack, except for the thermally conductive mounting area, are covered with multiple layers of thermal insulation components. The power controller and battery pack are then fixed to the preset mounting positions of the high thermal conductivity housing using polyimide pads, and all electrical connections are completed.

[0060] After the circuit layer of the solar cell array is laid, it is wound onto the roll mechanism, and then the roll mechanism is installed on the sun-facing side of the high thermal conductivity shell. Then the electrical interconnection between the solar cell array, the battery pack and the power controller is completed.

[0061] High emissivity thermal control white paint is sprayed or pasted on the outer surface of the high thermal conductivity shell and heat dissipation cover, and the heat dissipation cover is installed on the sun-facing side of the high thermal conductivity shell to complete the integrated assembly of the external power system.

[0062] Electromagnetic compatibility tests and vibration tests were conducted on the integrated external power system in sequence to ensure that it could meet the requirements of the launch mechanical environment and the on-orbit electromagnetic environment. Then, it was sent into a hot vacuum chamber for a hot vacuum test to verify its working performance and heat dissipation capacity in a vacuum and high and low temperature alternating environment. After all tests were passed, the integrated external power system was completed.

[0063] Example 3:

[0064] This embodiment provides a control method for an integrated external power system applied to a spacecraft, including the following steps:

[0065] The integrated external power system is installed as a whole at the pre-designated installation location outside the satellite cabin;

[0066] After the satellite is successfully launched and enters its predetermined orbit, the external integrated power system is in a power-off hibernation or safe standby mode.

[0067] The ground control center sends a power-on command to the integrated electronic system inside the satellite cabin via the uplink. The integrated electronic system sends a wake-up signal to the power controller outside the satellite cabin or directly connects the power supply switch via the command line of the multi-functional through-cabin interface. The solar array is deployed under ground control.

[0068] After powering on, the power controller performs a self-test. It controls the solar array to charge the battery pack. At the same time, the power controller continuously monitors the output of the solar array, the voltage / current / temperature of the battery pack, and the bus voltage parameters. Based on the monitored parameters, the power controller performs maximum power point tracking, battery charge and discharge management (constant current charging, constant voltage charging, trickle charging), and bus voltage regulation. It also periodically or when there are changes, reports the battery pack voltage, the internal temperature of the power controller, and the fault flag data to the integrated electronic system through the telemetry channel of the multi-functional through-cabin interface.

[0069] After receiving data, the integrated electronic system adjusts the battery pack charging current threshold according to the overall satellite mission plan. It then sends macro commands (including "enter shadow mode", "enter maximum power charging mode", and "disable a fault channel") to the power controller via the remote control channel. This enables the solar array to power the load and charge the battery pack during the sunshine period, and the battery to discharge and power the load during the shadow period, ultimately achieving an intelligent management mode of "on-board decision-making and off-board execution".

[0070] In this embodiment, the ground control center or integrated electronic system only needs to periodically receive the status telemetry of the external integrated power system without sending any control commands, and the system can operate stably for a long time. The power controller has a built-in fault diagnosis module, which can determine whether there are operational faults such as solar array output short circuit, battery overvoltage / undervoltage, output overcurrent, or internal overheating based on the output of the solar cell array, the voltage / current / temperature of the battery pack, and the bus voltage parameters. Once a fault is detected, the power controller will immediately execute a preset autonomous response strategy according to the fault level, such as: Level 1 fault (recoverable) attempt autonomous reset; Level 2 fault (affecting performance) degrade operation and report to the integrated electronic system; Level 3 fault (severe) immediately disconnect the fault path and report to the integrated electronic system to request intervention, ensuring the system's rapid response and isolation of faults.

[0071] For maintenance (such as adjusting the charging strategy due to battery capacity degradation after long-term operation), ground personnel can remotely update the control parameters of the power controller (such as charging termination voltage and maximum charging current) via uplink remote control commands, thereby achieving on-orbit software upgrades and maintenance.

[0072] The above description is merely a preferred embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any equivalent substitutions or modifications made by those skilled in the art within the scope disclosed in the present invention, based on the technical solution and inventive concept of the present invention, shall fall within the scope of protection of the present invention.

Claims

1. An integrated external power system for spacecraft, characterized in that: The system includes a solar array, a roll mechanism, a high thermal conductivity shell, a heat dissipation cover, a battery pack, and a power controller. The solar array is a flexible, rollable solar array wound around a roll mechanism, which is installed on the sun-facing side of the high thermal conductivity shell. During satellite assembly, the solar array is in a retracted state. After the satellite is launched and reaches its predetermined orbit, the solar array unfolds to generate power for the satellite. The heat dissipation cover is installed on the shaded side of the high thermal conductivity shell. The high thermal conductivity shell has pre-set mounting positions for the battery pack and power controller. The battery pack and power controller are installed side-by-side in the pre-set mounting positions, which effectively shortens the high-voltage, high-current transmission path. The power controller optimizes the electrical energy transmitted from the solar array, transmitting a portion of the energy to the battery pack and adjusting the other portion to a stable bus voltage, which is then output to the satellite cabin through the satellite's multi-functional trans-cabin interface. Simultaneously, the high thermal conductivity shell and heat dissipation cover directly dissipate the heat generated by the power controller and battery pack into the deep space environment through radiation.

2. The integrated external power system for spacecraft according to claim 1, characterized in that: The outer surfaces of the high thermal conductivity housing and heat dissipation cover are coated with high emissivity thermal control white paint. The high emissivity thermal control white paint radiates heat outward and reduces heat absorption during sunlight exposure, thereby reducing the temperature rise of the battery pack and power controller.

3. The integrated external power system for spacecraft according to claim 2, characterized in that: The high emissivity thermal control white paint has a high hemispherical emissivity ε>0.

85.

4. The integrated external power system for spacecraft according to claim 1, characterized in that: Both the high thermal conductivity shell and the heat dissipation cover are made of high thermal conductivity carbon fiber composite material or metal matrix composite material.

5. The integrated external power system for spacecraft according to claim 1, characterized in that: A thermally conductive mounting area is provided between the battery pack and the high thermal conductivity casing and heat dissipation cover. The thermally conductive mounting area is provided with thermally conductive filler or thermally conductive silicone pads to ensure good thermal contact. The remaining areas between the battery pack and the high thermal conductivity casing and heat dissipation cover, except for the thermally conductive mounting areas, are covered with multi-layer thermal insulation components. The remaining areas between the battery pack and the high thermal conductivity casing and heat dissipation cover, except for the thermally conductive mounting areas, are thermally insulated by polyimide pads.

6. The integrated external power system for spacecraft according to claim 1, characterized in that: A thermally conductive mounting area is provided between the power controller and the high thermal conductivity housing and heat dissipation cover plate. The thermally conductive mounting area is provided with thermally conductive filler or thermally conductive silicone pad to ensure good thermal contact. The remaining areas between the power controller and the high thermal conductivity housing and heat dissipation cover plate, except for the thermally conductive mounting area, are covered with multi-layer thermal insulation components. The remaining areas between the power controller and the high thermal conductivity housing and heat dissipation cover plate, except for the thermally conductive mounting area, are thermally insulated by polyimide pads.

7. The integrated external power system for spacecraft according to claim 1, characterized in that: The roll mechanism has extension rods on both sides for deploying the solar array. The extension rods are made of superelastic memory composite material. One end of the extension rod is fixed to the sun-facing side of the high thermal conductivity shell, and the other end is wound to the end of the roll mechanism and fixed by a constraint structure. In the initial state, the extension rod stores elastic strain energy. After the satellite is launched and reaches the predetermined orbit, the constraint structure is released by the satellite or ground command, so that the extension rod releases the strain energy instantly and quickly springs back to the extended state, thereby driving the solar array to deploy smoothly.

8. The integrated external power system for spacecraft according to claim 1, characterized in that: The power controller includes a maximum power point tracking and charging regulator unit for optimizing the electrical energy output from the solar array, a bus control and regulation unit for regulating the electrical energy to a stable bus voltage, and a fault judgment module for determining whether there are solar array output short circuits, battery overvoltage / undervoltage, output overcurrent, or internal overheating operation faults based on the output of the solar array, the voltage / current / temperature of the battery pack, and the bus voltage parameters.

9. A method for manufacturing an integrated external power system for a spacecraft as described in any one of claims 1 to 8, characterized in that, Including steps, The outer surfaces of the power controller and battery pack, except for the thermally conductive mounting area, are covered with multiple layers of thermal insulation components. The power controller and battery pack are then fixed to the preset mounting positions of the high thermal conductivity housing using polyimide pads, and all electrical connections are completed. After the circuit layer of the solar cell array is laid, it is wound onto the roll mechanism, and then the roll mechanism is installed on the sun-facing side of the high thermal conductivity shell. Then the electrical interconnection between the solar cell array, the battery pack and the power controller is completed. High emissivity thermal control white paint is sprayed or pasted on the outer surface of the high thermal conductivity shell and heat dissipation cover, and the heat dissipation cover is installed on the sun-facing side of the high thermal conductivity shell to complete the integrated assembly of the external power system. Electromagnetic compatibility testing, vibration testing, and thermal vacuum testing were conducted sequentially on the external integrated power system to complete its fabrication.

10. A control method for an integrated external power system applied to a spacecraft according to any one of claims 1 to 8, characterized in that, Including steps, The integrated external power system is installed as a whole at the pre-designated installation location outside the satellite cabin; After the satellite is successfully launched and enters its predetermined orbit, the external integrated power system is in a power-off hibernation or safe standby mode. The ground control center sends a power-on command to the integrated electronic system inside the satellite cabin via the uplink. The integrated electronic system sends a wake-up signal to the power controller outside the satellite cabin or directly connects the power supply switch via the command line of the multi-functional through-cabin interface. The solar array is deployed under ground control. After powering on, the power controller performs a self-test. It controls the solar array to charge the battery pack. At the same time, the power controller continuously monitors the output of the solar array, the voltage / current / temperature of the battery pack, and the bus voltage parameters. Based on the monitored parameters, the power controller performs maximum power point tracking, battery charge and discharge management, and bus voltage regulation. It also reports the battery pack voltage, the internal temperature of the power controller, and the fault flag data to the integrated electronic system periodically or when there are changes through the telemetry channel of the multi-functional through-cabin interface. After receiving data, the integrated electronic system adjusts the battery pack charging current threshold according to the overall satellite mission plan, and sends macro commands to the power controller through the remote control channel to enable the solar array to power the load and charge the battery pack during the sunshine period, and the battery to discharge and power the load during the shadow period, ultimately realizing the intelligent management mode of "on-board decision-making and off-board execution".

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