General circuit system of all-electric satellite

Abstract

The application provides a general circuit system of an all-electric propulsion satellite, comprising a power supply control and distribution unit, integrated with a power supply conversion module, a power distribution management module, a heater management module and a pyrotechnic device management module; an electric propulsion system high-voltage isolation power supply path, comprising a 100V DC bus and a cable distribution device drawn from the power supply control and distribution unit, the input end of the cable distribution device is connected with the 100V DC bus, the cable distribution device provides multiple independent 100V output channels, which are respectively connected to the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit at the end; a low-frequency cable network, used for realizing electrical connection between on-board devices; and a satellite-rocket interface. Thus, the application can meet the unique requirements of high-orbit satellites using multiple electric propulsion systems for orbit transfer and position keeping, such as high voltage, high power, light weight and high safety.

Description

Technical Field

[0001] This invention relates to the field of spacecraft overall design and power supply technology, and in particular to an overall circuit system for an all-electric propulsion satellite, which is widely applicable to spacecraft systems with multi-level and complex load voltage requirements. Background Technology

[0002] The overall circuit system is the satellite's energy hub and power distribution network. It is responsible for transmitting the primary power output from the power generation units (solar arrays), energy storage units (battery banks), and power control unit (PCU) to the electrical loads after power conversion, distribution, protection, and isolation. It provides electrical energy to all platforms, payload equipment, heaters, pyrotechnics, etc., meeting different voltage, power, timing, and quality requirements. Its design directly affects the satellite's mission capabilities, reliability, weight, and lifespan, and is a core component of the satellite's overall design. The overall circuit system architecture directly determines the overall satellite weight, bus efficiency, mission reliability, and electromagnetic compatibility.

[0003] Traditional high-orbit communication or remote sensing satellites typically use chemical propellants for orbit transfer and position maintenance, and their overall circuit system design is relatively mature. These designs often employ a combined distributed power distribution architecture, with multiple independent units such as platform power distributors, payload power distributors, thermal control managers, and pyrotechnics managers having clearly defined functions, interconnected through complex cable networks. However, with the development of all-electric propulsion technology, all-electric propulsion satellites that use electric propulsion systems to perform all or most of their orbital maneuvers are becoming the development trend. Such satellites present unprecedented new demands and challenges to the overall circuit system: First, power requirements have increased dramatically. Electric thrusters (such as ion thrusters and Hall thrusters) often operate at power levels ranging from thousands to tens of watts and require drive voltages of up to hundreds of volts, which traditional power distribution architectures cannot efficiently and safely handle. Second, power supply complexity has increased sharply. The satellite platform needs to simultaneously and efficiently manage multiple loads, including a single 100V bus unit, multi-stage voltage / power heaters, pyrotechnics, and high-voltage propulsion systems, placing higher demands on the integration and reliability of the overall circuit system. Third, there is an urgent need for lightweight design. To compensate for the additional weight of the electric thrusters and extend their lifespan, the overall circuit system itself must be significantly lighter to increase the payload ratio. Fourth, safety requirements are extreme. Fault isolation under high-voltage and high-current environments, secondary insulation of cable networks, and avoiding systemic risks of satellite failure caused by single-point faults are crucial.

[0004] Clearly, traditional distributed power distribution architectures are no longer able to meet the urgent needs of all-electric propulsion satellites for lightweight and high power density in terms of equipment configuration, management strategies, system weight, and reliability design.

[0005] A review of domestic and international literature and patent databases reveals that current research on all-electric propulsion satellites largely focuses on the electric thruster itself or its control algorithms. There are no publicly available solutions specifically tailored to the overall circuit system architecture design for all-electric propulsion satellites. Existing satellite power system patents primarily focus on localized improvements, such as efficiency enhancements for individual power modules or specific power distribution protection circuits. There is a lack of comprehensive solutions from a system-level perspective that address issues such as high-voltage propulsion power supply, integrated management of multi-level loads, system weight reduction, and isolation of high-risk components.

[0006] The limitations of existing traditional distributed overall circuit architectures for satellites in adapting to all-electric propulsion satellite applications include: (1) Traditional multi-unit distributed architecture (power controller + platform distributor + payload distributor + thermal control manager, etc.) is heavy, has a complex cable network and a long power path, which is not conducive to the lightweight design and reliability improvement of all-electric propulsion satellites.

[0007] (2) Traditional architecture cannot safely and efficiently meet the power supply and distribution requirements of 10,000-watt, high-voltage (hundreds to thousands of volts) electric propulsion systems of both ion and Hall types, and lacks dedicated isolation measures for high-voltage paths.

[0008] (3) Traditional cable network design has a large design margin in terms of conductor selection, connection redundancy and branch length in order to ensure the reliability and safety of the cable network, resulting in an unnecessary increase in the weight of the entire cable network.

[0009] (4) Traditional satellite-rocket interface relies on satellite cables inside the rocket. In the dynamic environment before separation of the satellite and rocket, there is a systemic risk that cable damage may cause short circuit failure of the entire satellite bus. In addition, the launch vehicle and satellite-related accessories occupy a large part of the total system weight and consume additional launch capacity resources.

[0010] Therefore, the industry urgently needs an innovative overall circuit system design to systematically address the unique requirements of all-electric propulsion satellites for an efficient, highly reliable, lightweight, and highly safe overall circuit system. Summary of the Invention

[0011] To address the shortcomings of existing technologies where traditional distributed overall circuit architectures for satellites cannot adapt to all-electric propulsion satellite applications, this invention proposes an overall circuit system for all-electric propulsion satellites. Through centralized integrated power distribution, thermal control and pyrotechnics management design, high-voltage propulsion power supply compatibility design, lightweight low-frequency cable network design, and optimized satellite-rocket interface design, this system meets the unique requirements of high-orbit satellites using multiple electric propulsion systems for orbit transfer and position holding, which require high voltage, high power, lightweight design, and high safety.

[0012] To achieve the above-mentioned technical effects, the present invention provides an overall circuit system for an all-electric propulsion satellite, comprising: The power control and distribution unit is a centralized integrated hardware platform that integrates a power conversion module, a power distribution management module, a heater management module, and a pyrotechnics management module. The high-voltage isolated power supply path of the electric propulsion system includes a 100V DC bus and a cable splitter led out from the power control and distribution unit. The input end of the cable splitter is connected to the 100V DC bus. The cable splitter provides multiple independent 100V output channels, which are respectively connected to the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit at the end. A low-frequency cable network is used to realize electrical connections between on-board equipment; wherein, the high-voltage isolation power supply path of the electric propulsion system adopts AWG8 high-power conductors; The satellite-rocket interface is configured with the following electrical connection path: no internal satellite cable is installed between the satellite and the launch vehicle, running from the launch vehicle fairing surface to the satellite; the satellite is powered by a ground umbilical cable during the pre-launch phase.

[0013] Furthermore, the power conversion module is used to convert and regulate the primary power generated by the solar cell array and the battery to produce a stable 100V DC bus. The power distribution management module provides two power distribution paths, 260W and 520W, for distributing power to 100V loads. Each power distribution branch is equipped with a fuse to rigidly isolate the load or line short circuit fault at the power source. The heater management module provides two heating voltage outputs, 7V and 100V, and integrates a temperature acquisition circuit. Each of its power distribution branches is equipped with a safety switch and a fuse for bus short circuit protection. The pyrotechnics management module provides a 28V pyrotechnics pulse bus for generating pyrotechnics detonation pulses, and the pyrotechnics bus is equipped with a current-limiting protection circuit and multi-level switch control.

[0014] Furthermore, in the high-voltage isolation power supply path of the electric propulsion system, the 100V DC bus is directly led out from the power control and distribution unit through the SUN interface to achieve physical isolation from the 100V power distribution load; The ion electric propulsion power processing unit and the Hall electric propulsion power processing unit are configured to independently perform voltage boosting conversion within themselves to generate their respective required high voltages; the power control and distribution unit and the cable splitter are configured to operate only at a voltage level of 100V.

[0015] Furthermore, the high-voltage isolation power supply path of the electric propulsion system adopts the following high-voltage path isolation measures: Primary isolation: PEEK insulating skeleton and epoxy vacuum infusion process are used inside the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit; Secondary isolation: At the input front end of the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit, all 100V power supply paths use AWG8 double-insulated wires; at the rear end of the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit, the high-voltage output uses armored high-voltage shielded cables and implements shielding grounding; the armored high-voltage shielded cables are spatially isolated from other types of cables on the satellite by adopting independent routing channels and separate binding methods.

[0016] Furthermore, the lightweight low-frequency cable network is achieved through any one or more of the following methods: The designated electrical nodes of the on-board equipment adopt a two-point, two-wire connection method. While meeting the conductor derating requirements, conductors should be selected based on the principle of the lightest wire type. A tailless design was developed for high-power electrical connectors and the J30JH type electrical connector used on satellites. The pyrotechnic pathway is optimized so that the output of the pyrotechnic management module is directly connected to the electrical pathway of the pyrotechnic via the pyrotechnic star table. During the layout design phase, the shortest path of various cables is optimized to make the main cable channel have an I-shaped layout, forming a fishbone-shaped coverage of the equipment on the north and south plates of the satellite and the equipment on the partition.

[0017] Furthermore, in the high-voltage isolated power supply path of the electric propulsion system, the voltage drop of the power supply path of the AWG8 high-power conductor under normal operating conditions is less than 1V, and the overall high-voltage leakage current of this path is less than 2.1mA.

[0018] Furthermore, the overall circuit system is configured to support the following launch site operation modes: During a preset time period before rocket launch, the electrical connection path of the satellite-rocket interface is powered and inspected by ground test equipment via a ground umbilical cable; at the first preset moment before launch, the electrical connection path is switched to power supply from the onboard battery, and the ground umbilical cable is in a detached state.

[0019] Furthermore, the overall circuit system is also configured to support an emergency power-off operation mode, specifically including: The pre-launch process retains an emergency power-off command injection point available up to the second preset time before rocket liftoff; after receiving the emergency power-off command before the second preset time, the system responds to the emergency power-off command and performs an emergency power-off operation.

[0020] Compared with the prior art, the technical solution provided by the present invention has the following significant advantages: (1) Significantly lightweight design and high power density: Through integrated design, the shells of multiple individual units, electrical connectors and a large number of interconnecting cables are eliminated, and the overall circuit system itself is reduced in weight by about 70%; the elimination of internal cables reduces the weight of the star-rocket system by about 30kg; shorter power distribution paths and lightweight design improve system power density.

[0021] (2) Improve the safety and reliability of the power supply of the system: The source fuse enables rapid and rigid isolation of short circuit faults in the load or cable network, preventing the fault from spreading to the primary bus and ensuring the energy safety of the entire satellite; the 100V bus of the electric propulsion high-voltage path is directly powered, and the voltage is boosted by the splitter and PPU, with a voltage drop of <1V and a leakage current of <2.1mA in the high-voltage path (better than the 5mA limit). The risk of high voltage crosstalk is eliminated from the architecture, and the safety of power supply and distribution is significantly improved, laying the architectural foundation for grid connection of multiple propulsion systems at the 10,000-watt level.

[0022] (3) Eliminate systemic risks: Eliminating the internal satellite cable fundamentally eliminates the catastrophic risk of satellite primary bus failure caused by short circuit due to cable damage during launch; the man-hours of the three parties at the launch site are reduced from 84 man-hours to 4 man-hours, and the AIT man-hours are reduced from 36 man-hours to 0.5 man-hours. The supporting joint operation and emergency power outage schemes provide comprehensive countermeasures, which are conducive to the mass production and rapid networking of satellites. Attached Figure Description

[0023] Figure 1 This is a block diagram of the overall circuit system architecture of the all-electric propulsion satellite provided in an embodiment of the present invention; Figure 2 This is a block diagram of a traditional distributed power controller. Figure 3 This is a structural block diagram of the power management and power distribution unit of the overall circuit system of the all-electric propulsion satellite provided in an embodiment of the present invention; Figure 4 A circuit diagram of the high-voltage isolated power supply path of the electric propulsion system in the overall circuit system of the all-electric propulsion satellite provided in an embodiment of the present invention; Figure 5 A comparative schematic diagram of the optimization of pyrotechnic pathways in the overall circuit system of the all-electric propulsion satellite according to an embodiment of the present invention; Figure 6 This is a schematic diagram of the routing of high-power cables and high-voltage cables in the overall circuit system of the all-electric propulsion satellite provided in an embodiment of the present invention; Figure 7 A schematic diagram showing the comparison of the star-rocket-ground interface of the overall circuit system of the all-electric propulsion satellite before and after optimization, provided for an embodiment of the present invention; Figure 8The launch site joint operation timing diagram of the overall circuit system of the all-electric propulsion satellite provided in an embodiment of the present invention. Detailed Implementation

[0024] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the invention.

[0025] It should be noted that references to "an embodiment," "embodiment," "example embodiment," etc., in this specification refer to the described embodiment including specific features, structures, or characteristics, but not every embodiment must include these specific features, structures, or characteristics. Furthermore, such expressions do not refer to the same embodiment. Moreover, when describing specific features, structures, or characteristics in conjunction with embodiments, whether or not explicitly described, it is indicated that incorporating such features, structures, or characteristics into other embodiments is within the knowledge of those skilled in the art.

[0026] Furthermore, certain terms are used in the specification and subsequent claims to refer to specific components or parts. Those skilled in the art will understand that manufacturers may use different names or terms to refer to the same component or part. This specification and subsequent claims do not distinguish components or parts by differences in name, but rather by differences in function. The terms "comprising" and "including" used throughout the specification and subsequent claims are open-ended and should be interpreted as "including but not limited to." Additionally, the term "connection" here includes any direct and indirect electrical connection means. Indirect electrical connection means include connections made through other means.

[0027] The overall circuit system of an all-electric propulsion satellite provided in one embodiment of the present invention has the following overall architecture: Figure 1 As shown, the system consists of four main parts: a power control and distribution unit, a high-voltage isolation power supply path for the electric propulsion system, a low-frequency cable network, and a star-rocket interface.

[0028] like Figure 3 As shown, the power control and distribution unit 10 is a physically centralized, integrated single hardware platform that integrates and replaces, for example, [other components]. Figure 2The traditional design separates the core power distribution and drive functions of the power controller (PCU), platform distributor, load distributor, thermal control manager (responsible for heater power distribution and temperature acquisition), and pyrotechnics manager. This design minimizes the 100V bus power distribution path, reduces intermediate connectors and cables, improves power supply reliability, and achieves significant weight reduction through structural integration. The unit employs a modular design internally, specifically comprising the following modules: The power conversion module 11 is used to convert and regulate the primary power generated by the solar cell array, lithium-ion battery, etc., to produce a stable 100V DC bus.

[0029] The power distribution management module 12 provides two power distribution paths, 260W and 520W, to manage the power distribution of 100V loads. Each power distribution branch of the module is equipped with a fuse to rigidly isolate load or line short-circuit faults at the power source end, preventing the fault from spreading to the primary bus and ensuring the overall energy safety of the satellite.

[0030] The heater management module 13 provides two heating voltage outputs: 7V and 100V, and integrates a temperature acquisition circuit. Safety switches and fuses are installed in the heater power distribution branch for busbar short-circuit protection.

[0031] The pyrotechnics management module 14 provides a 28V pyrotechnics pulse bus to emit pyrotechnics detonation pulses, thereby controlling the detonation of pyrotechnics such as solar panels and antennas. The pyrotechnics bus is equipped with current-limiting protection circuits and comprehensive multi-stage switching control to ensure the reliability and safety of pyrotechnics detonation control.

[0032] The high-voltage isolated power supply path of the electric propulsion system includes a 100V DC bus and a cable splitter derived from the power control and distribution unit. The input end of the cable splitter is connected to the 100V DC bus, and the cable splitter provides multiple independent 100V output channels, which are respectively connected to the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit at the end; for example... Figure 4 As shown, the composition and working principle of the high-voltage isolated power supply path for the electric propulsion system are as follows: First, the stable 100V DC bus generated by the power conversion module 11 in the power control and distribution unit 10 is directly led out through the SUN interface to achieve physical isolation from the 100V power distribution load.

[0033] Secondly, a cable splitter is set up as the core distribution node. The splitter receives a 100V bus input and provides multiple independent 100V output channels, which are respectively connected to the Ion Electric Propulsion Power Processing Unit (PPU) and the Hall Electric Propulsion Power Processing Unit (PPU).

[0034] Regarding high voltage generation, the approximately 1420V high voltage of the ion electric propulsion power processing unit and the approximately 400V high voltage of the Hall electric propulsion power processing unit are both independently boosted and converted within their respective PPUs. The power control and distribution unit 10 and the splitter body only operate at the 100V voltage level and do not directly generate or carry high voltage, thus eliminating the risk of high voltage crosstalk from the architectural perspective.

[0035] Regarding high-voltage path isolation measures, a two-level protection system of primary isolation and secondary isolation was adopted: Primary isolation: PEEK insulation skeleton and epoxy vacuum infusion process are used inside the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit; specifically, PEEK insulation skeleton and epoxy vacuum infusion process are used inside the high voltage PPU to ensure that the solid insulation meets the technical requirements of 1kVDC / mm and 400VAC / mm, and the creepage distance is not greater than 300V / mm.

[0036] Secondary isolation: At the input front end of the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit, all 100V power supply paths use AWG8 double-insulated wires; at the rear end of the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit, the high-voltage output uses armored high-voltage shielded cables and implements shielding grounding; the armored high-voltage shielded cables are spatially isolated from other types of cables on the satellite by adopting independent routing channels and separate binding methods, effectively suppressing electromagnetic interference (EMI) and potential high-voltage coupling risks.

[0037] The low-frequency cable network is used to realize the electrical connection between on-board equipment; wherein, the high-voltage isolation power supply path of the electric propulsion system adopts AWG8 high-power conductor; the voltage drop of the power supply path of the AWG8 high-power conductor under normal operating conditions is less than 1V, and the overall high-voltage leakage current of the path is less than 2.1mA.

[0038] The lightweight low-frequency cable network is achieved through any one or more of the following methods: The designated electrical nodes of the on-board equipment adopt a two-point, two-wire connection method. While meeting the conductor derating requirements, conductors should be selected based on the principle of the lightest wire type. A tailless design was developed for high-power electrical connectors and the J30JH type electrical connector used on satellites. The pyrotechnic pathway is optimized so that the output of the pyrotechnic management module is directly connected to the electrical pathway of the pyrotechnic via the pyrotechnic star table. During the layout design phase, the shortest path of various cables is optimized to make the main cable channel have an I-shaped layout, forming a fishbone-shaped coverage of the equipment on the north and south plates of the satellite and the equipment on the partition.

[0039] Specifically, the low-frequency cable network in this embodiment achieves system weight reduction and performance optimization through the following refined design methods: In terms of contact design, the configuration was optimized and the number of electrical contacts was simplified based on the actual conditions on the satellite. For important contacts of each subsystem that require redundant design, only a two-point, two-wire connection method was used, eliminating unnecessary multiple redundancies.

[0040] In terms of conductor design, conductor selection rules were optimized, and the selection range of double-insulated and shielded conductors was strictly controlled. The lightest possible conductor type was selected whenever possible while meeting derating requirements. In particular, AWG8 high-power conductors were used in the high-voltage electric propulsion distribution path, replacing the traditional solution of multiple 20-gauge conductors with a single thick conductor, significantly reducing the total weight of the cable network conductors.

[0041] In terms of electrical connector design, tailless designs were implemented for the AWG8 high-power electrical connector and the J30JH type electrical connector widely used on the satellite. Taking into account both the signal reception performance of the individual unit and the electromagnetic compatibility (EMC) performance of the entire satellite, tailless optimization was carried out for connectors without shielding requirements.

[0042] In terms of pyrotechnic product distribution channel design, such as Figure 5 As shown, the traditional "pyrotechnics management module - pyrotechnics star table, transition head, socket - pyrotechnics" path is optimized into a direct electrical path of "pyrotechnics management module - pyrotechnics star table - pyrotechnics", eliminating intermediate transition links and achieving a compatible design for path testing, pyrotechnics protection, and path optimization.

[0043] In terms of cable routing design, such as Figure 6 As shown, the shortest path optimization for the routing of various cables was carried out during the layout design phase. The main cable routing is in the shape of an "I" to achieve a herringbone-like coverage of the equipment on the north and south panels and the equipment on the partitions, minimizing the length of the cables.

[0044] The electrical connection path of the satellite-rocket interface is configured such that no internal satellite cable is installed between the satellite and the launch vehicle, running from the surface of the launch vehicle fairing to the satellite; the satellite is powered by a ground umbilical cable during the pre-launch phase.

[0045] like Figure 7 As shown, this embodiment optimizes the traditional "satellite-rocket-ground" connection relationship, eliminating the traditional internal satellite cable and corresponding satellite-rocket electrical interface that runs from the launch vehicle fairing surface to the satellite. By eliminating the internal satellite cable, the total weight of the satellite-rocket system is reduced by approximately 30 kg. More importantly, it fundamentally eliminates the catastrophic risk of the satellite's primary busbar failing directly due to short circuits caused by cable damage during launch.

[0046] The corresponding launch site operation strategy, such as Figure 8As shown, this includes the pre-launch joint operation procedures and the satellite emergency power-off plan: Regarding the pre-launch joint operation procedures, at a specific time point before rocket liftoff (e.g., T-8 hours to T-45 minutes), ground test equipment supplies power to the satellite via the umbilical cable and performs final checks. At T-45 minutes before liftoff (the first preset time), the system reliably switches to onboard battery power and disconnects the umbilical cable. This procedure reduces the man-hours for all three parties at the launch site from the traditional 84 man-hours to 4 man-hours, and the AIT (Autonomous In-Service) man-hours from 36 man-hours to 0.5 man-hours.

[0047] Regarding the satellite emergency power-off plan, an emergency power-off command injection point is set in the pre-launch process. The latest operable time for this injection point is T-5 seconds before rocket liftoff (the second preset time). In the event of an emergency requiring an emergency power-off operation on the satellite, the disconnect plug can be restored to the plugged-in state within 5 hours. By sending a ground command to disconnect the onboard battery circuit, the entire satellite can be powered off, providing comprehensive countermeasures for launch site safety.

[0048] The overall circuit system of the all-electric propulsion satellite constructed through the above specific embodiments achieves the following technical effects: In terms of weight reduction, the integrated design eliminates the need for multiple individual unit casings, electrical connectors, and a large number of interconnecting cables, resulting in a weight reduction of approximately 70% for the overall circuit system; the elimination of internal cables reduces the weight of the entire rocket and spacecraft system by approximately 30 kg.

[0049] In terms of power supply performance, the electric propulsion high-voltage path directly draws power from the 100V bus, and after being boosted by the splitter and PPU, the voltage drop of the power supply path is <1V, and the leakage current of the high-voltage path is <2.1mA (better than the 5mA limit), which significantly improves the safety of power supply and distribution and lays the architectural foundation for grid connection of multiple propulsion types at the 10,000-watt level.

[0050] In terms of reliability, the installation of source-end fuses enables rapid and rigid isolation of load or cable network short-circuit faults, preventing the fault from spreading to the primary busbar; the elimination of the satellite-borne cable fundamentally eliminates the risk of single-point failure due to short circuit of the entire satellite busbar during the launch phase.

[0051] The above description is merely a preferred embodiment of the present invention and is not intended to limit the invention. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

[0052] It should be noted that, in this document, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes that element. Furthermore, it should be noted that the scope of the methods and apparatuses in the embodiments of the present invention is not limited to performing functions in the order shown or discussed, but may also include performing functions substantially simultaneously or in the reverse order, depending on the functions involved. For example, the described methods may be performed in a different order than described, and various steps may be added, omitted, or combined. Additionally, features described with reference to certain examples may be combined in other examples.

[0053] Of course, the present invention may have other various embodiments. Without departing from the spirit and essence of the present invention, those skilled in the art can make various corresponding changes and modifications according to the present invention, but these corresponding changes and modifications should all fall within the protection scope of the appended claims.

Claims

1. A general circuit system for an all-electric propulsion satellite, characterized in that, Including: The power control and distribution unit is a centralized integrated hardware platform that integrates a power conversion module, a power distribution management module, a heater management module, and a pyrotechnics management module. The high-voltage isolated power supply path of the electric propulsion system includes a 100V DC bus and a cable splitter led out from the power control and distribution unit. The input end of the cable splitter is connected to the 100V DC bus. The cable splitter provides multiple independent 100V output channels, which are respectively connected to the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit at the end. A low-frequency cable network is used to realize electrical connections between on-board equipment; wherein, the high-voltage isolation power supply path of the electric propulsion system adopts AWG8 high-power conductors; The satellite-rocket interface is configured with the following electrical connection path: no internal satellite cable is installed between the satellite and the launch vehicle, running from the launch vehicle fairing surface to the satellite; the satellite is powered by a ground umbilical cable during the pre-launch phase.

2. The system according to claim 1, characterized in that, The power conversion module is used to convert and regulate the primary power generated by the solar cell array and the battery to produce a stable 100V DC bus. The power distribution management module provides two power distribution paths, 260W and 520W, for distributing power to 100V loads. Each power distribution branch is equipped with a fuse to rigidly isolate the load or line short circuit fault at the power source. The heater management module provides two heating voltage outputs, 7V and 100V, and integrates a temperature acquisition circuit. Each of its power distribution branches is equipped with a safety switch and a fuse for bus short circuit protection. The pyrotechnics management module provides a 28V pyrotechnics pulse bus for generating pyrotechnics detonation pulses, and the pyrotechnics bus is equipped with a current-limiting protection circuit and multi-level switch control.

3. The system according to claim 1, characterized in that, In the high-voltage isolation power supply path of the electric propulsion system, the 100V DC bus is directly led out from the power control and distribution unit through the SUN interface to achieve physical isolation from the 100V power distribution load. The ion electric propulsion power processing unit and the Hall electric propulsion power processing unit are configured to independently perform voltage boosting conversion within themselves to generate their respective required high voltages; the power control and distribution unit and the cable splitter are configured to operate only at a voltage level of 100V.

4. The system according to claim 1, characterized in that, The high-voltage isolation power supply path of the electric propulsion system adopts the following high-voltage path isolation measures: Primary isolation: PEEK insulating skeleton and epoxy vacuum infusion process are used inside the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit; Secondary isolation: At the input front end of the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit, all 100V power supply paths use AWG8 double-insulated wires; at the rear end of the ion electric propulsion power processing unit and the Hall electric propulsion power processing unit, the high-voltage output uses armored high-voltage shielded cables and implements shielding grounding; the armored high-voltage shielded cables are spatially isolated from other types of cables on the satellite by adopting independent routing channels and separate binding methods.

5. The system according to claim 1, characterized in that, The lightweight low-frequency cable network is achieved through any one or more of the following methods: The designated electrical nodes of the on-board equipment adopt a two-point, two-wire connection method. While meeting the conductor derating requirements, conductors should be selected based on the principle of the lightest wire type. A tailless design was developed for high-power electrical connectors and the J30JH type electrical connector used on satellites. The pyrotechnic pathway is optimized so that the output of the pyrotechnic management module is directly connected to the electrical pathway of the pyrotechnic via the pyrotechnic star table. During the layout design phase, the shortest path of various cables is optimized to make the main cable channel have an I-shaped layout, forming a fishbone-shaped coverage of the equipment on the north and south plates of the satellite and the equipment on the partition.

6. The system according to claim 1, characterized in that, In the high-voltage isolated power supply path of the electric propulsion system, the voltage drop of the power supply path of the AWG8 high-power conductor under normal operating conditions is less than 1V, and the overall high-voltage leakage current of the path is less than 2.1mA.

7. The system according to any one of claims 1 to 6, characterized in that, The overall circuit system is configured to support the following launch site operation modes: During a preset time period before rocket launch, the electrical connection path of the satellite-rocket interface is powered and inspected by ground test equipment via a ground umbilical cable; at the first preset moment before launch, the electrical connection path is switched to power supply from the onboard battery, and the ground umbilical cable is in a detached state.

8. The system according to any one of claims 1 to 6, characterized in that, The overall circuit system is also configured to support an emergency power failure operation mode, specifically including: The pre-launch process retains an emergency power-off command injection point available up to the second preset time before rocket liftoff; after receiving the emergency power-off command before the second preset time, the system responds to the emergency power-off command and performs an emergency power-off operation.

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