Lightweight spacecraft minimum safety system and implementation method

Abstract

The invention provides a lightweight spacecraft minimum safety system and an implementation method, aiming at reliable operation of a spacecraft in orbit, through configuration optimization of the minimum safety system, orderly power-up, cold and hot start and control time sequence configuration design are carried out on each component unit, and the reliability of the spacecraft is improved. The system configuration and construction method suitable for reliable operation of the lightweight spacecraft is provided, safe starting control logic and safe mode switching control logic of the spacecraft are achieved, communication of spaceflight energy flow, control flow and information flow can be ensured, the spacecraft operates orderly based on a minimum safety system, and the reliability of the spacecraft is improved. And effective support is provided for autonomous, safe and stable operation of spacecraft faults and ground emergency rescue. Through the configuration and construction method for the minimum safety system of the lightweight spacecraft, the reliability and safety of on-orbit autonomous execution of the lightweight spacecraft can be improved, and the influence of on-orbit faults on the safety of the spacecraft is reduced.

Description

Technical Field

[0001] The present invention relates to the technical field of on-orbit reliability and safety assurance of spacecraft, and in particular to a minimum safety system for a lightweight spacecraft and an implementation method thereof. Background Art

[0002] With the deployment of more and more lightweight spacecraft constellation systems in the future, purely relying on the ground for constellation control will cause a serious burden. At the same time, due to the limited measurement and control arc of low-orbit aerospace, the real-time performance of ground control is insufficient, and it is difficult to meet the needs of long-term reliable, stable and autonomous operation of the constellation in orbit. In particular, when encountering emergency faults in the non-measurement and control arc, traditional lightweight spacecraft may not be able to formulate targeted and effective fault plans in advance due to insufficient system robustness and adaptability. To deal with sudden situations on the orbit such as power failure and no telemetry, the spacecraft needs to have a certain autonomous fault handling capability. Based on this, a minimum safety system configuration and rapid construction of lightweight spacecraft is proposed, breaking through the minimum safety system forward design and rapid construction technology, and formulating the minimum safety system forward design and rapid construction strategy; It can be used for the design, testing, verification and evaluation of the reliability and safety of lightweight spacecraft throughout the life cycle and full-function performance, which is of great significance for improving the emergency handling and autonomous recovery capabilities of on-orbit faults.

[0003] Currently, no patents related to the configuration and construction methods of the minimum safety system based on lightweight spacecraft have been found.

[0004] Patent document CN117319838A (application number: 202311085423.3) discloses a SpaceVPX high-performance stand-alone minimum safety system and equipment based on management plane control, which realizes a minimum safety system based on the SpaceVPX hardware architecture. Its minimum safety system is a satellite-borne stand-alone device. The minimum safety system configuration and construction method of the present 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 a rapid on-orbit reconstruction method. The construction method of the present invention is mainly used for the construction of the lightweight spacecraft minimum system itself, in which the focus of the integrated electronic computer software is to realize the functions of attitude control solar orientation, load reduction after power failure and hot restart, and measurement and control link recovery.

[0006] Patent document CN115663783A (application number: 202211351466.7) discloses a bus power interactive power supply and distribution system based on the minimum system on board, which improves the reliability of the satellite through the bus power interactive power supply and distribution system based on the minimum system. This is quite different from the present invention, which realizes orderly power supply and distribution of each unit in the minimum system according to the startup 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 an intelligent driving mine car, and its composition and control method are significantly 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 the safety mode. The method reduces the types of safety modes, improves design efficiency, and improves the reliability of autonomous execution of spacecraft. It is quite different from the minimum safety system hardware and software configuration and construction method of the present invention.

[0009] Patent document CN108399300A (application number: 201810175233.3) discloses a method for configuring a safety mode for a spacecraft, which analyzes the dangers generated by the spacecraft, determines whether it is in the minimum mode, and configures the corresponding safety mode based on the risks, which is quite different from how the minimum safety system of the present invention is configured and constructed. Summary of the invention

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

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

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

[0013] The integrated electronic stand-alone is used to collect spacecraft telemetry data, receive instructions, control attitude sensor measurement and execution motors and electric propulsion components, and realize attitude control solar orientation, load reduction after power failure and hot and cold restart, and measurement and control link recovery functions;

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

[0015] The measurement and control unit is used for receiving remote control of the spacecraft and transmitting telemetry data;

[0016] The electric propulsion unit assembly is used for safe sun 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 the battery pack to realize the isolated power supply of the minimum safety system and other external loads; when the current demand is met, the power supply and distribution of each unit in the minimum safety system is realized according to the startup sequence;

[0019] Among them, the startup sequence is: power supply and distribution unit>integrated electronics unit>measurement and control unit>attitude sensing measurement and execution unit>electric propulsion unit assembly.

[0020] Preferably, the attitude sensing measurement and execution unit includes: a micro star sensor, a micro fiber optic gyroscope and a magnetic torquer, so as to realize the attitude determination and control of the lightweight spacecraft;

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

[0022] The micro-fiber gyroscope is used to perform high-precision stabilization and alignment control on a spacecraft under continuous attitude changes;

[0023] The magnetic torquer is used to perform attitude control and momentum management on the spacecraft using the geomagnetic field in an emergency mode, thereby achieving preliminary safety control over the sun.

[0024] Preferably, the system further comprises: triggering a low-power emergency communication mode in the event of a spacecraft failure; and in the low-power emergency communication mode, utilizing the integrated electronic stand-alone control to achieve information interaction with the ground.

[0025] Preferably, the electric propulsion unit assembly includes an electric propulsion cold gas propulsion unit and part of the pipelines. The cold gas propulsion unit serves as an emergency backup in case of spacecraft failure, and achieves safe solar control under the control of the integrated electronic motor.

[0026] Preferably, the spacecraft telemetry key command resources are allocated, and the telemetry and commands are configured separately, so that the integrated electronic unit can obtain the key telemetry data of each unit in the minimum safety system and have the ability to drive external load commands in a fault mode without relying on an external bus interface.

[0027] Preferably, the spacecraft safety startup control logic and safety mode switching control logic are implemented in a graphical manner.

[0028] A method for implementing a minimum safety system for a lightweight spacecraft provided by the present invention comprises:

[0029] Step S1: Using a single power supply and distribution machine to achieve stable energy supply, generation, storage, distribution, control and management for spacecraft;

[0030] Step S2: using the integrated electronic stand-alone machine to collect spacecraft telemetry data, receive commands, control attitude sensor measurement and execution motors and electric propulsion components, to achieve attitude control solar orientation, load reduction after power failure and hot and cold restart, and measurement and control link recovery functions;

[0031] Step S3: using the attitude sensor measurement and execution unit to measure and determine the attitude angle and angular velocity of the spacecraft, and completing the three-axis attitude stabilization control in cooperation with the execution unit;

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

[0033] Step S5: Using the electric propulsion unit 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 the battery pack to realize the isolated power supply of the minimum safety system and other external loads; when the current demand is met, the power supply and distribution of each unit in the minimum safety system is realized according to the startup sequence;

[0036] The startup sequence is: power supply and distribution unit > integrated electronics unit > measurement and control unit > attitude sensor measurement and execution unit > electric propulsion unit;

[0037] The attitude sensing measurement and execution unit includes: a micro star sensor, a micro fiber optic gyroscope and a magnetic torquer, which realizes the attitude determination and control of the lightweight spacecraft;

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

[0039] The micro-fiber gyroscope is used to perform high-precision stabilization and alignment control on a spacecraft under continuous attitude changes;

[0040] The magnetic torquer is used to perform attitude control and momentum management on the spacecraft using the geomagnetic field in an emergency mode, thereby achieving preliminary safety control over the sun.

[0041] Preferably, the method further comprises:

[0042] triggering a low-power emergency communication mode in the event of a spacecraft failure; and utilizing the integrated electronic stand-alone control to achieve information interaction with the ground in the low-power emergency communication mode;

[0043] The electric propulsion unit assembly includes an electric propulsion cold gas propulsion unit and some pipelines. The cold gas propulsion unit serves as an emergency backup in the event of a spacecraft failure. Under the control of the integrated electronic motor, safe solar control is achieved;

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

[0045] Compared with the prior art, the present invention has the following beneficial effects: the present invention focuses on the emergency safety mission requirements of spacecraft. During the launch process, the spacecraft will be subjected to severe overload, vibration and other environments. After entering orbit, it will also face the influence of space environments such as high vacuum, microgravity, high and low temperature alternation, and space ion radiation. If targeted measures are not taken for some failure modes, safety problems will arise for the entire spacecraft. The present invention aims at the problems of insufficient robustness of spacecraft, lack of system safety design and verification capabilities, lack of autonomous recovery capabilities of the minimum safety system, and insufficient impact on the spacecraft's emergency handling capabilities. A lightweight spacecraft minimum safety system and an implementation method are proposed to realize the configuration and rapid construction of the lightweight spacecraft minimum safety system, break through the minimum safety system forward design and rapid construction technology, and formulate a minimum safety system forward design and rapid construction strategy around the spacecraft safety startup control logic and the spacecraft safety mode switching control logic, thereby improving the reliability of autonomous execution of the spacecraft under emergency conditions. The system can be used for the design, testing, verification and evaluation of the reliability and safety of lightweight spacecraft throughout its life cycle and full-function performance, which is of great significance for improving the emergency handling and autonomous recovery capabilities of on-orbit faults. BRIEF DESCRIPTION OF THE DRAWINGS

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

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

[0048] Figure 2 Methods for constructing minimum safety systems for spacecraft;

[0049] Figure 3 Safely start control logic for spacecraft;

[0050] Figure 4 Control logic for spacecraft safe mode switching. DETAILED DESCRIPTION

[0051] The present invention is described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those of ordinary skill in the art, several changes and improvements can also be made without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0052] Example 1

[0053] According to a lightweight spacecraft minimum safety system and implementation method provided by the present invention, aimed at the reliable operation of the spacecraft on orbit, through the optimization configuration of the minimum safety component units, the power-on, startup and control timing configuration design of each component unit, a system and method suitable for the reliable operation of the lightweight spacecraft are constructed.

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

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

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

[0057] The integrated electronic stand-alone is used to collect spacecraft telemetry data, receive instructions, control attitude sensor measurement and execution motors and electric propulsion components, and realize attitude control solar orientation, load reduction after power failure and hot and cold restart, and measurement and control link recovery functions;

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

[0059] The measurement and control unit is used for receiving remote control of the spacecraft and transmitting telemetry data;

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

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

[0062] The power controller is used to control the solar panels and the battery pack to realize the isolated power supply of the minimum safety system and other external loads; when the current demand is met, the power supply and distribution of each unit in the minimum safety system is realized according to the startup sequence;

[0063] Among them, the startup sequence is: power supply and distribution unit>integrated electronics unit>measurement and control unit>attitude sensing measurement and execution unit>electric propulsion unit assembly.

[0064] The attitude sensing measurement and execution unit includes: a micro star sensor, a micro fiber optic gyroscope and a magnetic torquer, which realizes the attitude determination and control of the lightweight spacecraft;

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

[0066] The micro-fiber gyroscope is used to perform high-precision stabilization and alignment control on a spacecraft under continuous attitude changes;

[0067] The magnetic torquer is used to perform attitude control and momentum management on the spacecraft using the geomagnetic field in an emergency mode, thereby 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 failure; in the low-power emergency communication mode, utilizing the integrated electronic stand-alone control to achieve information interaction with the ground.

[0069] The electric propulsion unit assembly includes an electric propulsion cold gas propulsion unit and some pipelines. The cold gas propulsion unit serves as an emergency backup in the event of a spacecraft failure. Under the control of the integrated electronic motor, safe solar control is achieved.

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

[0071] The spacecraft safety startup control logic and safety mode switching control logic are implemented in a graphical way. The spacecraft safety startup control logic and safety mode switching control logic are two different internal control logics, one for the startup control after the whole satellite power failure, and the other for non-power failure safety mode switching. Both mode control logics can choose autonomous switching and ground injection remote control switching.

[0072] The configuration and construction method of the minimum safety system of the lightweight spacecraft, such as Figure 2 As shown, including:

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

[0074] In this embodiment, the above-mentioned supporting stand-alone products can be independent stand-alone products or comprehensive module products.

[0075] Step S2: Configure DC / DC. The working voltage of the basic product can be obtained through DC / DC selection. The starting voltage value of each unit and the priority sequence of the starting voltage can be obtained through DC / DC external control circuit configuration and selection, so as to realize the orderly power supply of each unit in the minimum system. Considering the order of energy, measurement and control link, and control, the lower the starting voltage, the easier it is to power on, indicating the principle of higher priority. The starting voltage priority sequence of orderly power supply and distribution is: V1 <V2<V3<V4<V5;

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

[0077] Table 1 Startup priority allocation table

[0078] Serial number Single machine name Starting voltage Current Power consumption Remark 1 Power control standalone V1 I1 P1 2 Comprehensive electronic stand-alone V2 I2 P2 3 Single machine for measurement and control 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 comprehensive electronic computer hardware and software products. The software adopts the operating system architecture to realize telemetry collection, command reception, attitude control and attitude sensor measurement and execution control, and complete the coordinated loading of monitoring software, management software and application software; the hardware realizes the collection of spacecraft telemetry and the reception of key commands, and has the ability to drive and control the electric propulsion cold thruster, realize attitude control solar orientation, load reduction after power failure and restart, and measurement and control link recovery.

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

[0081] Step S4: Allocate spacecraft telemetry key instruction resources to enable the integrated electrical stand-alone machine to obtain the minimum system key telemetry data and have instruction driving capability in the fault mode without relying on the external bus interface.

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

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

[0084] In the above step S6, the spacecraft does not autonomously use cold air propulsion. When the ground determines that the spacecraft is in a faulty state, the instruction chain enables cold air propulsion. After the cold air propulsion function is turned on, if the whole satellite loses power and restarts, or the integrated electronic reset fails, the software sends a valve closing instruction to prohibit jetting.

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

[0086] In the above step S7, Figure 4 The graphical representation shown selects and expresses the state machine switching, and provides the conditions and methods for switching the spacecraft safety mode.

[0087] Those skilled in the art know that, in addition to realizing the system and its various devices, modules, and units provided by the present invention in a purely computer-readable program code, it is entirely possible to realize the same functions in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers by logically programming the method steps. Therefore, the system and its various devices, modules, and units provided by the present invention can be considered as a hardware component, and the devices, modules, and units included therein for realizing various functions can also be regarded as structures within the hardware component; the devices, modules, and units for realizing various functions can also be regarded as both software modules for realizing the method and structures within the hardware component.

[0088] The above describes the specific embodiments of the present invention. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essence of the present invention. In the absence of conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A minimum safety system for lightweight spacecraft, It is characterized in that include: Power supply and distribution unit, integrated electronics unit, attitude sensing measurement and execution unit, measurement and control unit and electric propulsion unit components; The power supply and distribution unit is used for stable energy supply, generation, storage, distribution, control and management of spacecraft; The integrated electronic stand-alone is used to collect spacecraft telemetry data, receive instructions, control attitude sensor measurement and execution motors and electric propulsion components, and realize attitude control solar orientation, load reduction after power failure and hot and cold restart, and measurement and control link recovery functions; The attitude sensing measurement and execution unit is used to complete the measurement and determination of the attitude angle and angular velocity of the spacecraft, and complete the three-axis attitude stabilization control in cooperation with the execution unit; The measurement and control unit is used for receiving remote control of the spacecraft and transmitting telemetry data; The electric propulsion unit assembly is used for safe sun control in emergency situations.

2. The lightweight spacecraft minimum safety system according to claim 1, It is characterized in that The power supply and distribution unit includes: a solar panel, a battery pack and a power controller; The power controller is used to control the solar panels and the battery pack to realize the isolated power supply of the minimum safety system and other external loads; when the current demand is met, the power supply and distribution of each unit in the minimum safety system is realized according to the startup sequence; Among them, the startup sequence is: power supply and distribution unit>integrated electronics unit>measurement and control unit>attitude sensing measurement and execution unit>electric propulsion unit assembly.

3. The lightweight spacecraft minimum safety system according to claim 1, It is characterized in that The attitude sensing measurement and execution unit includes: a micro star sensor, a micro fiber optic gyroscope and a magnetic torquer, which realizes the attitude determination and control of the lightweight spacecraft; The micro star sensor is used to realize high-precision solar safety control of the spacecraft using stars in emergency mode; The micro-fiber gyroscope is used to perform high-precision stabilization and alignment control on a spacecraft under continuous attitude changes; The magnetic torquer is used to perform attitude control and momentum management on the spacecraft using the geomagnetic field in an emergency mode, thereby achieving preliminary safety control over the sun.

4. The lightweight spacecraft minimum safety system according to claim 1, It is characterized in that The system also includes: triggering a low-power emergency communication mode in the event of a spacecraft failure; in the low-power emergency communication mode, utilizing the integrated electronic stand-alone control to achieve information interaction with the ground.

5. The lightweight spacecraft minimum safety system according to claim 1, It is characterized in that The electric propulsion unit assembly includes an electric propulsion cold gas propulsion unit and some pipelines. The cold gas propulsion unit serves as an emergency backup in the event of a spacecraft failure. Under the control of the integrated electronic motor, safe solar control is achieved.

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

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

8. A method for realizing a minimum safety system for a lightweight spacecraft. It is characterized in that include: Step S1: Using a single power supply and distribution machine to achieve stable energy supply, generation, storage, distribution, control and management for spacecraft; Step S2: using the integrated electronic stand-alone machine to collect spacecraft telemetry data, receive commands, control attitude sensor measurement and execution motors and electric propulsion components, to achieve attitude control solar orientation, load reduction after power failure and hot and cold restart, and measurement and control link recovery functions; Step S3: using the attitude sensor measurement and execution unit to measure and determine the attitude angle and angular velocity of the spacecraft, and completing the three-axis attitude stabilization control in cooperation with the execution unit; Step S4: using the measurement and control unit to realize the spacecraft remote control reception and telemetry downlink; Step S5: Using the electric propulsion unit to achieve safe sun control in emergency situations.

9. The method for realizing the minimum safety system of a lightweight spacecraft according to claim 8, It is characterized in that The power supply and distribution unit includes: a solar panel, a battery pack and a power controller; The power controller is used to control the solar panels and the battery pack to realize the isolated power supply of the minimum safety system and other external loads; when the current demand is met, the power supply and distribution of each unit in the minimum safety system is realized according to the startup sequence; The startup sequence is: power supply and distribution unit > integrated electronics unit > measurement and control unit > attitude sensor measurement and execution unit > electric propulsion unit; The attitude sensing measurement and execution unit includes: a micro star sensor, a micro fiber optic gyroscope and a magnetic torquer, which realizes the attitude determination and control of the lightweight spacecraft; The micro star sensor is used to realize high-precision solar safety control of the spacecraft using stars in emergency mode; The micro-fiber gyroscope is used to perform high-precision stabilization and alignment control on a spacecraft under continuous attitude changes; The magnetic torquer is used to perform attitude control and momentum management on the spacecraft using the geomagnetic field in an emergency mode, thereby achieving preliminary safety control over the sun.

10. The method for realizing the minimum safety system of a lightweight spacecraft according to claim 8, It is characterized in that The method further comprises: triggering a low-power emergency communication mode in the event of a spacecraft failure; and utilizing the integrated electronic stand-alone control to achieve information interaction with the ground in the low-power emergency communication mode; The electric propulsion unit assembly includes an electric propulsion cold gas propulsion unit and some pipelines. The cold gas propulsion unit serves as an emergency backup in the event of a spacecraft failure. Under the control of the integrated electronic motor, safe solar control is achieved; Allocate spacecraft telemetry key command resources, configure telemetry and commands separately, and enable the integrated electronic unit to obtain the key telemetry data of each unit in the minimum safety system in a fault mode without relying on an external bus interface and have the ability to drive external load commands.

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