Minimum safety system for lightweight spacecraft and its implementation method

Abstract

This invention provides a minimum safety system and implementation method for lightweight spacecraft. Addressing the reliable operation of spacecraft in orbit, it optimizes the configuration of the minimum safety system, designing the orderly power-up, cold / hot start-up, and control timing configurations for each component. This proposes a system configuration and construction method suitable for the reliable operation of lightweight spacecraft, realizing the spacecraft's safe start-up control logic and safe mode switching control logic. It ensures the smooth flow of spacecraft energy, control, and information, enabling the orderly operation of the spacecraft based on the minimum safety system. This provides effective support for the spacecraft's autonomous, safe, and stable operation during faults and for ground emergency rescue. Through this invention's minimum safety system configuration and construction method for lightweight spacecraft, the reliability and safety of autonomous execution in orbit can be improved, and the impact of on-orbit faults on spacecraft safety can be reduced.

Description

Technical Field

[0001] This invention relates to the field of spacecraft on-orbit reliability and safety assurance technology, specifically to a lightweight spacecraft minimum safety system and its implementation method. Background Technology

[0002] With the increasing deployment of lightweight spacecraft constellation systems in the future, relying solely on ground-based constellation control will place a heavy burden on the system. Furthermore, the limited telemetry and control (TT&C) arcs in low-Earth orbit (LEO) spacecraft result in insufficient real-time ground control, making it difficult to meet the requirements for long-term reliable, stable, and autonomous operation of the constellation in orbit. Particularly in the event of emergency failures outside the TT&C arcs, traditional lightweight spacecraft, due to insufficient system robustness and adaptability, may be unable to develop targeted and effective contingency plans in advance to address on-orbit power outages, lack of telemetry, and other unforeseen circumstances. Therefore, spacecraft need to possess a certain degree of autonomous fault handling capability. Based on this, a minimum safety system configuration and rapid construction method for lightweight spacecraft is proposed, breaking through the technologies of minimum safety system forward design and rapid construction, and formulating a minimum safety system forward design and rapid construction strategy. This method can be used for the design, testing, verification, and evaluation of the reliability and safety of lightweight spacecraft throughout their entire life cycle and across all functions, and is of great significance for improving on-orbit fault emergency handling and autonomous recovery capabilities.

[0003] No patents related to the minimum safety system configuration and construction method based on lightweight spacecraft have been found.

[0004] Patent document CN117319838A (application number: 202311085423.3) discloses a SpaceVPX high-performance single-unit minimum safety system and device based on management plane control, which realizes a minimum safety system based on the SpaceVPX hardware architecture. The minimum safety system is a spaceborne single-unit device. The minimum safety system configuration and construction method of this invention is aimed at lightweight spacecraft.

[0005] Patent document CN117112035A (application number: 202310928201.7) discloses a flexible onboard minimum safety system and a rapid construction method, which solves the problems of high efficiency and high reliability of software-defined satellites. The CPU with an operating system in the minimum system is used to boot the program and complete the initialization configuration. The minimum system is used for on-orbit rapid reconfiguration. The construction method of this invention is mainly used for the construction of the minimum system itself of lightweight spacecraft. The key focus of the integrated electronic computer software is to realize functions such as attitude control and sun orientation, load reduction after power failure and cold / hot restart, and restoration of telemetry and control links.

[0006] Patent document CN115663783A (application number: 202211351466.7) discloses a bus power interactive power supply and distribution system based on the minimum system on the satellite. It improves the reliability of the satellite by using a bus power interactive power supply and distribution system based on the minimum system. This is quite different from the present invention, which achieves orderly power supply and distribution of each unit within the minimum system according to the start-up sequence while meeting the current demand.

[0007] Patent document CN113085822B (application number: 202110470705.X) discloses a minimum safety system, control method, device and storage medium for intelligent driving mining trucks, the composition and control method of which are quite different from the present invention.

[0008] Patent document CN108427394B (application number: CN201810147358.5) discloses a spacecraft safety mode analysis method, which focuses on how to analyze safety modes. The method reduces the types of safety modes, improves design efficiency, and improves the reliability of spacecraft autonomous execution. It is quite different from the minimum safety system hardware and software configuration and construction method of this invention.

[0009] Patent document CN108399300A (application number: 201810175233.3) discloses a method for configuring a safety mode for a spacecraft. It analyzes the dangers generated by the spacecraft, determines whether it is in the minimum mode, and configures the corresponding safety mode according to the risk. This method is quite different from how the minimum safety system of this invention is configured and constructed. Summary of the Invention

[0010] In view of the deficiencies in the prior art, the purpose of this invention is to provide a lightweight spacecraft minimum safety system and its implementation method.

[0011] A lightweight spacecraft minimum safety system according to the present invention includes: a power supply and distribution unit, an integrated electronic unit, an attitude sensing, measurement and execution unit, a telemetry and control unit, and an electric propulsion assembly;

[0012] The aforementioned power supply and distribution unit is used for the stable supply, generation, storage, distribution, control, and management of energy for spacecraft.

[0013] The integrated electronic unit is used to collect spacecraft telemetry data, receive commands, control attitude sensing and measurement, and execute motors and electric propulsion components, so as to realize the functions of attitude control for solar orientation, load reduction after power failure and cold / hot restart, and restoration of telemetry and control links.

[0014] The attitude sensing, measurement and execution unit is used to measure and determine the attitude angle and angular velocity of the spacecraft, and to complete the three-axis attitude stabilization control with the cooperation of the execution unit.

[0015] The aforementioned telemetry and control unit is used for receiving remote control data from spacecraft and transmitting telemetry data downlink.

[0016] The electric propulsion assembly is used for safe solar control in emergency situations.

[0017] Preferably, the power supply and distribution unit includes: a solar panel, a battery pack, and a power controller;

[0018] The power controller is used to control the solar panels and battery pack to achieve isolated power supply to the internal components of the minimum safety system and other external loads; while meeting the current requirements, it enables orderly power supply and distribution to each unit within the minimum safety system according to the startup sequence.

[0019] The startup sequence is as follows: power supply and distribution unit > integrated electronic unit > measurement and control unit > attitude sensing, measurement and execution unit > electric propulsion assembly.

[0020] Preferably, the attitude sensing, measurement, and execution unit includes: a miniature star sensor, a miniature fiber optic gyroscope, and a magnetic torque converter, to achieve attitude determination and control of lightweight spacecraft;

[0021] Among them, the micro-star sensor is used to achieve high-precision solar safety control of spacecraft using stars in emergency mode;

[0022] The aforementioned miniature fiber optic gyroscope enables high-precision stabilization and alignment control of spacecraft under continuous attitude changes.

[0023] In emergency mode, the magnetic torque device is used to perform attitude control and momentum management of the spacecraft using the Earth's magnetic field, achieving preliminary safety control over the Sun.

[0024] Preferably, the system further includes: triggering a low-power emergency communication mode in the event of a spacecraft malfunction; and using the integrated electronic unit control to achieve information interaction with the ground in the low-power emergency communication mode.

[0025] Preferably, the electric propulsion assembly includes an electric propulsion unit with a cold gas propulsion system and some piping. The cold gas propulsion unit serves as an emergency backup in case of spacecraft failure and, under the integrated electronic and electrical control, enables safe solar control.

[0026] Preferably, key telemetry command resources for spacecraft are allocated, and telemetry and commands are configured separately. This enables the integrated electronic unit to obtain key telemetry data of each unit within the minimum safety system and to have the ability to drive external load commands in fault mode without relying on external bus interfaces.

[0027] Preferably, a graphical approach is used to implement the spacecraft safe start-up control logic and safe mode switching control logic.

[0028] A method for implementing a lightweight spacecraft minimum safety system according to the present invention includes:

[0029] Step S1: Utilize a power supply and distribution unit to achieve stable energy supply, generation, storage, distribution, control, and management for the spacecraft;

[0030] Step S2: Utilize the integrated electronic unit to collect spacecraft telemetry data, receive commands, control attitude sensing and measurement, and execute motors and electric propulsion components to achieve the functions of attitude control for solar orientation, load reduction after power failure and cold / hot restart, and restoration of the telemetry and control link;

[0031] Step S3: Utilize attitude sensing measurement and execution unit to measure and determine the spacecraft's attitude angle and angular velocity, and complete the three-axis attitude stabilization control with the assistance of the execution unit;

[0032] Step S4: Use the telemetry and control unit to realize the remote control reception and telemetry downlink of the spacecraft;

[0033] Step S5: Utilize the electric propulsion assembly to achieve safe sun control in emergency situations.

[0034] Preferably, the power supply and distribution unit includes: a solar panel, a battery pack, and a power controller;

[0035] The power controller is used to control the solar panels and battery pack to achieve isolated power supply to the internal components of the minimum safety system and other external loads; while meeting the current requirements, it enables orderly power supply and distribution to each unit within the minimum safety system according to the startup sequence.

[0036] The startup sequence is as follows: power supply and distribution unit > integrated electronic unit > measurement and control unit > attitude sensing, measurement and execution unit > electric propulsion assembly;

[0037] The attitude sensing, measurement and execution unit includes: a miniature star sensor, a miniature fiber optic gyroscope and a magnetic torque converter, to realize attitude determination and control of lightweight spacecraft;

[0038] Among them, the micro-star sensor is used to achieve high-precision solar safety control of spacecraft using stars in emergency mode;

[0039] The aforementioned miniature fiber optic gyroscope enables high-precision stabilization and alignment control of spacecraft under continuous attitude changes.

[0040] In emergency mode, the magnetic torque device is used to perform attitude control and momentum management of the spacecraft using the Earth's magnetic field, achieving preliminary safety control over the Sun.

[0041] Preferably, the method further includes:

[0042] In the event of a spacecraft malfunction, a low-power emergency communication mode is triggered; in the low-power emergency communication mode, the integrated electronic unit control is used to achieve information interaction with the ground.

[0043] The electric propulsion assembly includes an electric propulsion unit with a cold gas propulsion system and some piping. The cold gas propulsion unit serves as an emergency backup in case of spacecraft failure and, under the integrated electronic and electrical control, enables safe control of the sun.

[0044] Allocate key telemetry command resources for spacecraft, configure telemetry and commands separately, and enable the integrated electronic unit to obtain key telemetry data of each unit within the minimum safety system and have the ability to drive external load commands in fault mode without relying on external bus interfaces.

[0045] Compared with existing technologies, this invention has the following beneficial effects: Addressing the emergency safety requirements of spacecraft, which endure severe overloads and vibrations during launch and face the influence of high vacuum, microgravity, alternating high and low temperatures, and space ion radiation after entering orbit, some fault modes, if not addressed with specific measures, can pose safety risks to the entire spacecraft. This invention addresses issues such as insufficient spacecraft robustness, lack of system safety design and verification capabilities, inadequate autonomous recovery capabilities of the minimum safety system, and insufficient impact on the spacecraft's ability to handle emergencies. It proposes a lightweight spacecraft minimum safety system and its implementation method, enabling the configuration and rapid construction of a lightweight spacecraft minimum safety system. It breaks through the technology of forward design and rapid construction of minimum safety systems, and formulates a strategy for forward design and rapid construction of the minimum safety system around the spacecraft's safe startup control logic and safe mode switching control logic, improving the reliability of autonomous execution in emergency situations. It can be used for the design, testing, verification, and evaluation of the reliability and safety of lightweight spacecraft throughout their entire life cycle and full functionality, and is of great significance for improving on-orbit fault emergency handling and autonomous recovery capabilities. Attached Figure Description

[0046] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0047] Figure 1 Configuration of the minimum safety system for spacecraft;

[0048] Figure 2 A method for constructing a minimum safety system for spacecraft;

[0049] Figure 3 To ensure the safe activation of spacecraft control logic;

[0050] Figure 4 This is the control logic for switching safety modes on spacecraft. Detailed Implementation

[0051] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the scope of protection of the present invention.

[0052] Example 1

[0053] According to the present invention, a minimum safety system and implementation method for lightweight spacecraft are provided. For the reliable operation of spacecraft in orbit, a system and method suitable for the reliable operation of lightweight spacecraft are constructed by optimizing the configuration of minimum safety components and designing the power-on, startup and control timing configuration of each component.

[0054] like Figure 1 As shown, the lightweight spacecraft minimum safety system includes:

[0055] Power supply and distribution unit, integrated electronic unit, attitude sensing, measurement and execution unit, measurement and control unit, and electric propulsion unit components;

[0056] The aforementioned power supply and distribution unit is used for the stable supply, generation, storage, distribution, control, and management of energy for spacecraft.

[0057] The integrated electronic unit is used to collect spacecraft telemetry data, receive commands, control attitude sensing and measurement, and execute motors and electric propulsion components, so as to realize the functions of attitude control for solar orientation, load reduction after power failure and cold / hot restart, and restoration of telemetry and control links.

[0058] The attitude sensing, measurement and execution unit is used to measure and determine the attitude angle and angular velocity of the spacecraft, and to complete the three-axis attitude stabilization control with the cooperation of the execution unit.

[0059] The aforementioned telemetry and control unit is used for receiving remote control data from spacecraft and transmitting telemetry data downlink.

[0060] The electric propulsion assembly is used for safe solar control in emergency situations.

[0061] The power supply and distribution unit includes: solar panels, battery packs, and a power controller;

[0062] The power controller is used to control the solar panels and battery pack to achieve isolated power supply to the internal components of the minimum safety system and other external loads; while meeting the current requirements, it enables orderly power supply and distribution to each unit within the minimum safety system according to the startup sequence.

[0063] The startup sequence is as follows: power supply and distribution unit > integrated electronic unit > measurement and control unit > attitude sensing, measurement and execution unit > electric propulsion assembly.

[0064] The attitude sensing, measurement and execution unit includes: a miniature star sensor, a miniature fiber optic gyroscope and a magnetic torque converter, to realize attitude determination and control of lightweight spacecraft;

[0065] Among them, the micro-star sensor is used to achieve high-precision solar safety control of spacecraft using stars in emergency mode;

[0066] The aforementioned miniature fiber optic gyroscope enables high-precision stabilization and alignment control of spacecraft under continuous attitude changes.

[0067] In emergency mode, the magnetic torque device is used to perform attitude control and momentum management of the spacecraft using the Earth's magnetic field, achieving preliminary safety control over the Sun.

[0068] The system also includes: triggering a low-power emergency communication mode in the event of a spacecraft malfunction; and in the low-power emergency communication mode, using the integrated electronic unit control to achieve information interaction with the ground.

[0069] The electric propulsion assembly includes an electric propulsion unit with a cold gas propulsion system and some piping. The cold gas propulsion unit serves as an emergency backup in case of spacecraft failure and, under the integrated electronic and electrical control, enables safe solar control.

[0070] Allocate key telemetry command resources for spacecraft, configure telemetry and commands separately, and enable the integrated electronic unit to obtain key telemetry data of each unit within the minimum safety system and have the ability to drive external load commands in fault mode without relying on external bus interfaces.

[0071] The spacecraft safety startup control logic and safety mode switching control logic are implemented using a graphical interface. These two logics are separate internal control logics: one for startup control after a power failure, and the other for switching to a safe mode without power failure. Both logics can be switched autonomously or remotely from the ground.

[0072] The method for configuring and constructing the minimum safety system for lightweight spacecraft, such as Figure 2 Shown, including:

[0073] Step S1: Select lightweight supporting single-unit products, including power supply and distribution unit, integrated electronic computer, attitude sensing and measurement motor, attitude execution unit, measurement and control unit and electric propulsion actuator assembly.

[0074] In this embodiment, each of the above-mentioned supporting stand-alone products can be a separate stand-alone product or a comprehensive modular product.

[0075] Step S2: Configure the DC / DC converter. The operating voltage of the basic product can be obtained through DC / DC selection. By configuring and selecting the external control circuit of the DC / DC converter, the starting voltage values ​​and priority sequence of each unit can be obtained, achieving orderly power supply to each unit within the minimum system. Considering the order of ensuring energy supply, monitoring and control links, and control, lower starting voltages are easier to power on, indicating higher priority. The priority sequence of starting voltages for orderly power supply is: V1 <V2<V3<V4<V5;

[0076] In this embodiment, the priority order of the starting voltage of each single unit in step S2 above can be statistically analyzed according to Table 1, and the current I of the solar panel is ensured to be greater than I1+I2+I3+I4+I5.

[0077] Table 1 Startup Priority Allocation Table

[0078] Serial Number Single-player name Start-up voltage Current Power consumption Remark 1 Power control stand-alone unit V1 I1 P1 2 Integrated electronic single machine V2 I2 P2 3 Measurement and control stand-alone machine V3 I3 P3 4 Attitude sensing, measurement and execution stand-alone V4 I4 P4 5 Electric propulsion components V5 I5 P5

[0079] Step S3: Configure integrated electronic computer hardware and software products. On the software side, an operating system architecture is adopted to realize telemetry acquisition, command reception, attitude control sensing measurement and execution control, and to complete the collaborative loading of monitoring software, management software and application software. On the hardware side, it realizes the acquisition of spacecraft telemetry and reception of key commands, and has the ability to drive and control the electric propulsion cold thruster unit, realizing functions such as attitude control for sun orientation, load reduction after power failure restart and restoration of telemetry and control links.

[0080] In this embodiment, step S3 above needs to take into account both cold and hot start-up and emergency start-up control judgment logic, specifically as follows: Figure 3 As shown, this ensures that the integrated electronic computer's judgment logic is safe and reliable before and after the separation of the satellite and rocket, and after cold start and hot start.

[0081] Step S4: Allocate key telemetry command resources for the spacecraft to enable the integrated electrical unit to obtain the minimum key telemetry data within the system and have command-driven capabilities in fault mode without relying on external bus interfaces.

[0082] Step S5: Select attitude sensing measurement and execution units, including miniature star sensors, miniature optical fibers, and magnetic torquers.

[0083] Step S6: Select a propulsion unit, including a cold thrust assembly. The cold thrust assembly serves as an emergency backup in case of spacecraft failure and enables solar control under the integrated electrical unit control.

[0084] In step S6 above, the spacecraft does not autonomously use cold gas propulsion. Once the ground determines that the spacecraft's condition is faulty, the injection command chain activates cold gas propulsion. After activating the cold gas propulsion function, if a satellite power failure and restart or an integrated electronic reset fault occurs, the software sends a valve closure command to prohibit jet propulsion.

[0085] Step S7: Select typical safety control modes and construct safe integrated electrical and attitude control mode switching conditions. The switching between modes is divided into two types according to the design and on-orbit reliability requirements: autonomous switching and ground-based remote control switching.

[0086] The above step S7 uses Figure 4 The graphical representation shows the selection and expression of state machine switching, and provides the conditions and methods for switching spacecraft safety modes.

[0087] Those skilled in the art will understand that, besides implementing the system and its various devices, modules, and units provided by this invention in the form of purely computer-readable program code, the same functions can be achieved entirely through logical programming of the method steps, making the system and its various devices, modules, and units of this invention function in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers. Therefore, the system and its various devices, modules, and units provided by this invention can be considered as a hardware component, and the devices, modules, and units included therein for implementing various functions can also be considered as structures within the hardware component; alternatively, the devices, modules, and units for implementing various functions can be considered as both software modules implementing the method and structures within the hardware component.

[0088] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A lightweight spacecraft minimum safety system, characterized in that, include: Power supply and distribution unit, integrated electronic unit, attitude sensing, measurement and execution unit, measurement and control unit, and electric propulsion unit components; The aforementioned power supply and distribution unit is used for the stable supply, generation, storage, distribution, control, and management of energy for spacecraft. The integrated electronic unit is used to collect spacecraft telemetry data, receive commands, control attitude sensing and measurement, and execute motors and electric propulsion components, so as to realize the functions of attitude control for solar orientation, load reduction after power failure and cold / hot restart, and restoration of telemetry and control links. The attitude sensing, measurement and execution unit is used to measure and determine the attitude angle and angular velocity of the spacecraft, and to complete the three-axis attitude stabilization control with the cooperation of the execution unit. The aforementioned telemetry and control unit is used for receiving remote control data from spacecraft and transmitting telemetry data downlink. The electric propulsion assembly is used for safe solar control in emergency situations.

2. The lightweight spacecraft minimum safety system according to claim 1, characterized in that, The power supply and distribution unit includes: solar panels, battery packs, and a power controller; The power controller is used to control the solar panels and battery pack to achieve isolated power supply to the internal components of the minimum safety system and other external loads; while meeting the current requirements, it enables orderly power supply and distribution to each unit within the minimum safety system according to the startup sequence. The startup sequence is as follows: power supply and distribution unit > integrated electronic unit > measurement and control unit > attitude sensing, measurement and execution unit > electric propulsion assembly.

3. The lightweight spacecraft minimum safety system according to claim 1, characterized in that, The attitude sensing, measurement and execution unit includes: a miniature star sensor, a miniature fiber optic gyroscope and a magnetic torque converter, to realize attitude determination and control of lightweight spacecraft; Among them, the micro-star sensor is used to achieve high-precision solar safety control of spacecraft using stars in emergency mode; The aforementioned miniature fiber optic gyroscope enables high-precision stabilization and alignment control of spacecraft under continuous attitude changes. In emergency mode, the magnetic torque device is used to perform attitude control and momentum management of the spacecraft using the Earth's magnetic field, achieving preliminary safety control over the Sun.

4. The lightweight spacecraft minimum safety system according to claim 1, characterized in that, The system also includes: triggering a low-power emergency communication mode in the event of a spacecraft malfunction; and in the low-power emergency communication mode, using the integrated electronic unit control to achieve information interaction with the ground.

5. The lightweight spacecraft minimum safety system according to claim 1, characterized in that, The electric propulsion assembly includes an electric propulsion unit with a cold gas propulsion system and some piping. The cold gas propulsion unit serves as an emergency backup in case of spacecraft failure and, under the control of the integrated electronic unit, enables safe control of the sun.

6. The lightweight spacecraft minimum safety system according to claim 1, characterized in that, Allocate key telemetry command resources for spacecraft, configure telemetry and commands separately, and enable the integrated electronic unit to obtain key telemetry data of each unit within the minimum safety system and have the ability to drive external load commands in fault mode without relying on external bus interfaces.

7. The lightweight spacecraft minimum safety system according to claim 1, characterized in that, The spacecraft safe start-up control logic and safe mode switching control logic are implemented in a graphical manner.

8. A method for implementing a minimum safety system for lightweight spacecraft, characterized in that, include: Step S1: Utilize a power supply and distribution unit to achieve stable energy supply, generation, storage, distribution, control, and management for the spacecraft; Step S2: Utilize the integrated electronic unit to collect spacecraft telemetry data, receive commands, control attitude sensing and measurement, and execute motors and electric propulsion components to achieve the functions of attitude control for solar orientation, load reduction after power failure and cold / hot restart, and restoration of the telemetry and control link; Step S3: Utilize attitude sensing measurement and execution unit to measure and determine the spacecraft's attitude angle and angular velocity, and complete the three-axis attitude stabilization control with the assistance of the execution unit; Step S4: Use the telemetry and control unit to realize the remote control reception and telemetry downlink of the spacecraft; Step S5: Utilize the electric propulsion assembly to achieve safe sun control in emergency situations.

9. The method for implementing a lightweight spacecraft minimum safety system according to claim 8, characterized in that, The power supply and distribution unit includes: solar panels, battery packs, and a power controller; The power controller is used to control the solar panels and battery pack to achieve isolated power supply to the internal components of the minimum safety system and other external loads; while meeting the current requirements, it enables orderly power supply and distribution to each unit within the minimum safety system according to the startup sequence. The startup sequence is as follows: power supply and distribution unit > integrated electronic unit > measurement and control unit > attitude sensing, measurement and execution unit > electric propulsion assembly; The attitude sensing, measurement and execution unit includes: a miniature star sensor, a miniature fiber optic gyroscope and a magnetic torque converter, to realize attitude determination and control of lightweight spacecraft; Among them, the micro-star sensor is used to achieve high-precision solar safety control of spacecraft using stars in emergency mode; The aforementioned miniature fiber optic gyroscope enables high-precision stabilization and alignment control of spacecraft under continuous attitude changes. In emergency mode, the magnetic torque device is used to perform attitude control and momentum management of the spacecraft using the Earth's magnetic field, achieving preliminary safety control over the Sun.

10. The method for implementing a lightweight spacecraft minimum safety system according to claim 8, characterized in that, The method further includes: In the event of a spacecraft malfunction, a low-power emergency communication mode is triggered; in the low-power emergency communication mode, the integrated electronic unit control is used to achieve information interaction with the ground. The electric propulsion assembly includes an electric propulsion unit with a cold gas propulsion system and some piping. The cold gas propulsion unit serves as an emergency backup in case of spacecraft failure and, under the control of the integrated electronic unit, enables safe control of the sun. Allocate key telemetry command resources for spacecraft, configure telemetry and commands separately, and enable the integrated electronic unit to obtain key telemetry data of each unit within the minimum safety system and have the ability to drive external load commands in fault mode without relying on external bus interfaces.

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