A method and apparatus for constructing an integrated electronic system for spacecraft

Abstract

This invention discloses a method and apparatus for constructing an integrated electronic system for a spacecraft. The method includes: determining the functional modules included in the integrated electronic system of the spacecraft, and the number of functions or interfaces corresponding to each functional module; constructing interface specifications between the functional modules; determining the number of integrated electronic units included in the spacecraft and the functional range of each integrated electronic unit; iterating the functional range and functional modules of each integrated electronic unit; and selecting target functional modules to combine into integrated electronic units for each partition according to the current functional and interface quantity requirements, thereby constituting the integrated electronic system of the spacecraft. The spacecraft integrated electronic system construction method disclosed in this invention results in a spacecraft integrated electronic system with high functional integration, high resource reuse, good generalization, and strong adaptability to demand.

Description

Technical Field

[0001] This invention belongs to the field of aerospace integrated electronic technology, and in particular relates to a method and apparatus for constructing an integrated electronic system for spacecraft. Background Technology

[0002] Traditional spacecraft platform electronic system design typically employs a clearly defined subsystem approach, distributing electronic systems or equipment across multiple subsystems such as power supply, data management, telemetry, tracking, attitude and orbit control, propulsion, and thermal control based on major functional categories. This low functional density results in large overall weight, power consumption, and size, leading to high launch costs. The independent operation of subsystems results in a wide variety of products, high development costs, and difficulty in coordinating hardware resources. Furthermore, the equipment has extremely low standardization, being largely custom-made products; even slight changes in requirements necessitate design modifications.

[0003] With the development of my country's space program, spacecraft are becoming increasingly larger and more complex in function. Constrained by factors such as carrying capacity and launch costs, the functional density requirements for electronic systems are becoming increasingly stringent. At the same time, my country's space program is gradually entering the field of deep space exploration. Due to constraints such as long exploration distances, complex orbital control, and limited light energy, the weight, size, and power consumption of spacecraft are subject to stringent limitations.

[0004] Adopting an integrated electronic solution breaks down traditional subsystem boundaries, allowing for the integrated design of the entire platform's electronic systems. This improves the integration and sharing of hardware resources, reducing the overall weight, size, and power consumption, thus meeting relevant constraints. However, due to the increased functional integration of the integrated electronic system, its coupling with other subsystems is also quite high, making its design iteration process extremely complex and lengthy. This often results in a disadvantageous development cycle, as changes in one aspect can have far-reaching consequences, potentially leading to repeated iterations of the entire system's development. Therefore, simply increasing functional integration without addressing customization issues cannot meet the requirements of rapid development and the needs of my country's rapidly evolving aerospace industry. Summary of the Invention

[0005] The technical problem to be solved by the present invention is to provide a method and apparatus for constructing an integrated electronic system for spacecraft. The integrated electronic system for spacecraft constructed by the method has high functional integration, high resource reuse, good generalization and strong adaptability to demand.

[0006] To address the aforementioned technical problems, this invention discloses a method for constructing an integrated electronic system for spacecraft, wherein the method includes:

[0007] Determine the functional modules included in the spacecraft's integrated electronic system, and the number of functions or interfaces corresponding to each functional module;

[0008] Establish interface specifications between the various functional modules;

[0009] Determine the number of integrated electronic units contained in the spacecraft and the functional scope of each integrated electronic unit;

[0010] The functional scope and functional modules of each integrated electronic unit are iterated;

[0011] Based on the current functional and interface requirements, target functional modules are selected and combined into integrated electronic units for each zone, thus forming the spacecraft's integrated electronic system.

[0012] Optionally, the step of determining the functional modules included in the spacecraft's integrated electronic system includes:

[0013] The spacecraft's integrated electronic system is categorized into functional modules based on three main categories: power supply, central processing, and interface processing. The power supply category includes high-power power supply modules and low-power power supply modules. The central processing category includes high-performance processor modules, medium- and low-performance processor modules, and bus communication functional modules. The interface processing category is further divided into functional modules based on interface type.

[0014] Optionally, the interface specifications include: mechanical interface specifications, electrical interface specifications, and software interface specifications.

[0015] Optionally, the mechanical interface specifications include uniformly defined dimensions of the printed circuit board, dimensions of the printed circuit board metal frame, dimensions of the front and back envelopes of the assembled printed circuit board, model and layout of the internal bus connectors used for inter-module insertion, dimensions of the machine cover plate, and dimensions of the machine base plate.

[0016] Optionally, the electrical interface specification includes: uniformly defined bus connector contact signals between modules, electrical characteristics such as signal voltage and current, and signal timing.

[0017] Optionally, the step of determining the number of integrated electronic units contained in the spacecraft and the functional range of each integrated electronic unit includes: determining the number of integrated electronic units contained in the spacecraft and the functional range of each integrated electronic unit according to the functional requirements of different partitions of the spacecraft.

[0018] Optionally, the integrated electronic unit includes a main control unit and an integrated electronic remote control unit; the main control unit uses a high-performance central processing module for overall control, and selects a high-power or medium-low power supply module according to the overall power consumption requirements; the main control unit forwards remote control commands and sends program control commands to each of the integrated electronic remote control units through a bus, and collects telemetry information, health information and backup information of each of the integrated electronic remote control units.

[0019] Optionally, the integrated electronic remote control unit uses a low-to-medium power central processing module for overall control, and selects a high-power or low-to-medium power supply module according to the overall power consumption requirements.

[0020] Optionally, the iterative steps for the functional scope and functional modules of the integrated electronic unit include: stacking and interlocking the functional modules, transmitting the internal bus signals through interlocking functional modules, and keeping the signals corresponding to the same contact number consistent, so as to facilitate arbitrary interlocking and compatibility between functional modules and rapid combination and change of the integrated electronic unit.

[0021] The spacecraft integrated electronic system construction method disclosed in this invention has the following advantages: on the one hand, the constructed spacecraft integrated electronic system has high functional integration, high resource reuse, good generalization, and strong adaptability to demand; on the other hand, the method has high efficiency in constructing the spacecraft integrated electronic system. Attached Figure Description

[0022] Figure 1 This is a flowchart illustrating the steps of the spacecraft integrated electronic system construction method provided in an embodiment of the present invention;

[0023] Figure 2 This is a schematic diagram illustrating the mechanical interface requirements for printed circuit boards as specified in this embodiment of the invention;

[0024] Figure 3 This is a schematic diagram of the mechanical interface adaptation specified in the embodiments of the present invention;

[0025] Figure 4 This is a schematic diagram illustrating the mechanical interface envelope size requirements specified in the embodiments of the present invention;

[0026] Figure 5 This is a schematic diagram of the functional module combination in an embodiment of the present invention;

[0027] Figure 6 This is a structural block diagram of a spacecraft integrated electronic system construction device provided in an embodiment of the present invention. Detailed Implementation

[0028] The present invention will now be described in detail with reference to specific embodiments and accompanying drawings. It should be understood that the specific embodiments described herein are for illustrative purposes only and are not intended to limit the scope of the invention.

[0029] This invention addresses the problems of decentralized development, difficulty in resource coordination, poor adaptability to changing requirements, long development cycles, and high development costs of electronic systems for large or multi-module spacecraft in existing technologies. It provides a modular distributed integrated electronic system design scheme for spacecraft, and the specific implementation method is as follows.

[0030] Figure 1A flowchart illustrating the steps of a method for constructing an integrated electronic system for a spacecraft, as provided in an embodiment of the present invention.

[0031] like Figure 1 As shown, the spacecraft integrated electronic system construction method of this invention includes the following steps:

[0032] Step 101: Determine the functional modules included in the spacecraft's integrated electronic system, and the number of functions or interfaces corresponding to each functional module.

[0033] An optional method for determining the functional modules included in a spacecraft's integrated electronic system is as follows: The integrated electronic system is categorized into three main types: power supply, central processing, and interface processing. The power supply category includes high-power power supply modules and low-power power supply modules; the central processing category includes high-performance processor modules, medium-to-low-performance processor modules, and bus communication functional modules; and the interface processing category is divided according to interface type.

[0034] In practical implementation, the functional scope of the spacecraft's integrated electronic system can be determined, and different functions can be categorized and summarized into different functional modules. The number of functions or interfaces included in each functional module is initially planned. In this embodiment, the functional scope of the spacecraft's integrated electronic system can be divided into: data management, remote control processing, command transmission, telemetry acquisition, telemetry downlink, thermal control, pyrotechnics drive, DC brushless motor drive, and stepper motor drive.

[0035] The power supply module should be able to provide 5V and ±15V power supplies, and can be divided into two types: a high-power module with a power of approximately 50W and a medium-low power module with a power of 10-20W, to meet different power supply needs. The central processing module can also be divided into two types: a high-performance processor module with a processing capacity of no less than 80MIPS and a medium-low performance processor module with a processing capacity of 10-20MIPS, to meet different needs for complex and simple processing. The central processing module should include bus communication functionality, configured according to the overall bus selection. Interface processing modules are divided into various types based on interface type, and interfaces with similar single interface types or circuits should be grouped into the same functional module as much as possible. The number of interfaces in each module should be estimated based on the preset circuit board size. In this embodiment, a 60W high-power power module and a 20W medium-low power module are planned. The high-performance processor is selected as BM3803, the medium-low performance processor is selected as 80C32 microcontroller, and the system bus is selected as 1553B high-speed serial bus. Based on the system's functional scope, the following interface processing modules are planned: a remote control processing module, capable of receiving and processing 4 uplink telemetry signals; a command sending module, capable of outputting 64 OC gate commands; an A / D acquisition module, capable of acquiring 100 signals (compatible with voltage and resistance); a telemetry downlink module, capable of outputting 4 telemetry downlink signals; a heater control module, capable of controlling 50 heaters; a pyrotechnic drive module, capable of driving 20 dual-bridge wire pyrotechnics; a DC brushless motor drive module, capable of controlling 5 motors; and a stepper motor drive module, capable of controlling 5 motors.

[0036] Step 102: Construct interface specifications between various functional modules.

[0037] Interface specifications include mechanical interface specifications, electrical interface specifications, and software interface specifications. These interface specifications between functional modules enable seamless physical integration, electrical interconnection, matching interface levels and timing, and prevent software access conflicts.

[0038] The mechanical interface specifications include standardized dimensions of the printed circuit board (PCB), the PCB metal frame, the front and back envelope dimensions of the assembled PCB, the model and layout of the internal bus connectors used for inter-module interconnection, the dimensions of the overall cover plate, and the dimensions of the overall base plate. The electrical interface specifications include standardized definitions of the internal bus connector contact signals between modules, definitions of electrical characteristics such as signal voltage and current, and signal timing definitions. All functional modules use standardized internal bus contact definitions. The contact signal definitions should maintain a safe distance between power and ground signals as much as possible. Data lines, address lines, and control lines defined in the contact signals that are related to CPU (central processing unit) access should be compatible with both high-performance and low-performance processors. The contact definitions may include some dedicated signals for point-to-point interconnection between modules, and may also reserve some dedicated signal points. Access timing should be defined separately for high-performance and low-performance processors. The timing design of the interface between each interface processing functional module and the CPU is implemented using a programmable logic array. By injecting different versions of FPGA (Field Programmable Gate Array) software, timing matching with two different processor modules can be achieved.

[0039] The software interface specification that allows for arbitrary combinations of functional modules mainly refers to the unified definition of the access address ranges for each interface processing functional module. The access address ranges of each functional module should be kept separate from those of other functional modules to avoid access conflicts. For a single integrated electronic unit that may contain more than one functional module, multiple address ranges should be specified according to requirements, and corresponding functional modules should be designed with board selection jumpers to quickly adapt to different address range access requirements.

[0040] In this embodiment of the invention, the dimensions of the printed circuit board are specified as follows: Figure 2 As shown, the thickness of the printed circuit board (PCB) is uniformly required to be 2mm. The corresponding PCB metal frame dimensions match this, with a wall thickness of 2mm. The dimensions of the metal cover plate and metal base plate also match the frame dimensions. Specific dimensional diagrams are not provided in this embodiment. The internal bus connector used for inter-module insertion is the IEH HMM-276, a through-type board-to-board connector. Its positioning layout on the PCB is already shown in [the diagram]. Figure 2 The standard specifies that the effect after the printed circuit board and frame are assembled is shown in the image. Figure 3 ,in Figure 3 The left image shows the front view of the printed circuit board (PCB) assembled with the metal frame, and the right image shows the back view. To accommodate the insertion positions of these connectors, detailed specifications have been made for the height of the metal frame and the height of the components mounted on both sides of the PCB, such as... Figure 4As shown, the frame height of all modules equipped with this internal bus connector is 20mm, the height of components on the front of the printed circuit board must not exceed 13mm, and the height of components on the back must not exceed 3mm.

[0041] The electrical interface specifications that allow for arbitrary combinations of functional modules include uniformly defined inter-module internal bus connector contact signals, electrical characteristics such as signal voltage and current, and signal timing definitions. All modules use a unified internal bus contact definition. The contact signal definition should maintain a safe distance between power and ground signals as much as possible. Data lines, address lines, and control lines defined in the contact signals that are related to CPU access should be compatible with both high-performance and low-to-medium performance processors. The contact definition may include some dedicated signals for point-to-point interconnection between modules, and may also reserve some dedicated signal points. Access timing should be defined separately for high-performance and low-to-medium performance processors. The interface timing design between each interface processing functional module and the CPU is implemented using a programmable logic array. By injecting different versions of FPGA software, timing matching with two different processor modules can be achieved. In this embodiment, there are 276 internal bus contacts, including a dual-machine internal bus. Each machine defines 20 5V power supply lines, 10 15V power supply lines, 30 ground lines, 16 address lines, 16 data lines, 4 interrupt request signal lines, 3 global clock lines, and 6 control lines for read / write signals. The integrated electronic unit using the BM3803 processor occupies all 16 data lines, while the integrated electronic unit using the 80C32 processor only occupies the first 8 address lines, with the last 8 considered unused. The CPU access timing for functional modules is defined according to the usage requirements of the BM3803 and 80C32 processors. Multiplexed modules such as A / D acquisition are configured with two versions of FPGA software adapted to both the BM3803 and 80C32 processors, and are programmed as needed.

[0042] The software interface specification that allows for arbitrary combinations of functional modules mainly refers to the unified definition of the access address ranges for each interface processing functional module. The access address ranges of each functional module should be staggered from those of other functional modules to avoid access conflicts. For functional modules that may be installed in more than one unit within an integrated electronic unit, multiple address ranges should be specified according to requirements. Corresponding functional modules should be designed with board selection jumpers to quickly adapt to different address range access requirements. In this embodiment, different address ranges are assigned to each functional module connected to the internal bus. For example, the heater control module's address range is 0x5000~0x5FFF, the telemetry downlink module's address range is 0x8000~0x8FFF, and the A / D acquisition module, since two of these modules may be installed in one integrated electronic unit, has an address range of 0xB000~0xCFFF. Board selection jumpers are also set on the module; when board selection is 0, the address range is 0xB000~0xBFFF, and when board selection is 1, the address range is 0xC000~0xCFFF.

[0043] Step 103: Determine the number of integrated electronic units contained in the spacecraft and the functional scope of each integrated electronic unit.

[0044] In practical implementation, the number of integrated electronic units (IEUs) and the functional scope of each IEU can be determined based on the functional requirements of different partitions within the spacecraft. A total of n IEUs are arranged across n partitions, with one IEU serving as the main control unit. The remaining IEUs communicate with the main control unit via a high-speed bus. The main control unit acts as the information hub and management center of the entire integrated electronic system. The IEUs other than the main control unit can also be referred to as remote IEUs. One exemplary partition management method involves dividing the spacecraft into four partitions based on equipment layout and signal acquisition and control requirements, with one main control unit and three remote IEUs. One remote IEU is located in a module that needs to be separated in orbit.

[0045] Spacecraft zone management should fully consider signal routing and the functional scale of a single integrated electronic unit. Zone division should minimize the length of signal cable paths leading to the integrated electronic unit in that zone, and ensure that the module combination of the integrated electronic unit in that zone does not exceed 12 modules and has a minimum of 5 modules. A single integrated electronic unit should not manage multiple modules.

[0046] The main control unit employs a high-performance central processing module for overall system control, selecting either a high-power or low-to-medium power supply module based on the overall power consumption requirements. The main control unit forwards remote control commands and sends programmable control commands to each of the integrated electronic remote control units via a bus, and collects telemetry, health, and backup information from each integrated electronic remote control unit. The main control unit also acts as the master controller of the entire system's bus, controlling all bus communications. The integrated electronic remote control units employ low-to-medium power central processing modules for overall system control, selecting either a high-power or low-to-medium power supply module based on the overall power consumption requirements.

[0047] In one feasible implementation, based on the estimated energy and performance requirements, the main control unit uses a 60W high-power power supply module and a BM3803 central processing module. It communicates with three remote units via a 1553B bus. The main control unit is the BC side of the 1553B bus, while the three remote units and other terminal devices are the RT side. Based on the estimated energy and performance requirements, each of the three remote units uses a 20W low-to-medium power supply module and an 80C32 central processing module.

[0048] Step 104: Iterate on the functional scope and functional modules of each integrated electronic unit.

[0049] Based on the current functional scope of each integrated electronic unit, corresponding functional modules are selected and combined to form integrated electronic units for each zone. The types of functional modules selected for different integrated electronic units can overlap. Taking into account the utilization rate of module resources and the composition scale of the integrated electronic units under the current combination, appropriate adjustments are made to the module types, the functions or interface scale of certain modules, or the number and functional composition of the integrated electronic units. The number of functions or interfaces possessed by a functional module can be slightly greater than the required number, provided that circuit board space resources are fully utilized, to facilitate functional or interface expansion. A utilization rate of circuit resources on the module should not be less than 80% to be considered reasonable. The maximum module combination scale of an integrated electronic unit is no more than 12, and the minimum is no less than 5.

[0050] In one feasible embodiment, the initial module combination scheme is as follows: The main control unit consists of 10 modules: 1 power supply module, 2 BM3803 central processing modules, 1 remote control processing module, 2 command output modules, 1 A / D acquisition module, 1 telemetry downlink module, and 2 heater control modules. Remote unit one consists of 10 modules: 1 power supply module, 2 80C32 central processing modules, 1 command output module, 1 A / D acquisition module, 1 heater control module, 2 pyrotechnic drive modules, and 2 stepper motor drive modules. Remote unit two consists of 6 modules: 1 power supply module, 2 80C32 central processing modules, 1 command output module, 1 A / D acquisition module, and 1 heater control module. Remote unit three consists of 9 modules: 1 power supply module, 2 80C32 central processing modules, 1 command output module, 1 A / D acquisition module, 2 heater control modules, and 2 DC brushless motor control modules. In subsequent iterations, due to the increased number of regional commands and regional data acquisition requirements, the main control unit added one command output module, and the remote unit 2 added one A / D acquisition module. Except for a few command output modules, A / D acquisition modules, and heater drive modules whose actual usage did not reach 100%, the resource utilization rate of all other modules reached 100%.

[0051] This step should be carried out at an appropriate time, after the functional requirements of each integrated electronic unit have been basically determined, in order to avoid major changes in the future and reduce the number of iterations and costs.

[0052] An optional method for iterating the functional scope and functional modules of an integrated electronic unit is as follows: the functional modules are stacked and inter-insulated, and the internal bus signals are transmitted layer by layer through the inter-plug-in of the functional modules. The signals corresponding to the same contact number are kept consistent, so as to facilitate arbitrary inter-insulation compatibility between functional modules and rapid combination and change of the integrated electronic unit.

[0053] The integrated electronic unit modules are constructed using a stacked, interlocking module design, without a motherboard. Internal bus signals are transmitted layer by layer through interlocking modules, with signals corresponding to the same contact number remaining consistent. This facilitates seamless interlocking and compatibility between modules and rapid combination and modification of the integrated electronic unit. The contact definitions of the through-type connectors between functional modules are completely identical, allowing functional modules to be interlocked and combined in any order. The combined effect of the functional modules is as follows: Figure 5 As shown.

[0054] Step 105: Based on the current functional and interface requirements, select target functional modules to combine into integrated electronic units for each partition, thus forming the spacecraft integrated electronic system.

[0055] After the integrated electronic system of the spacecraft is constructed, when the functional or interface quantity requirements change, if the change is small, the first consideration is to utilize the existing circuit margins of the integrated electronic unit modules or reduce the number of similar circuits assembled on the modules to adapt to the change. If the change is large, then module-level additions or subtractions are made to adapt to the change. In this embodiment, after the composition of the four integrated electronic unit modules is determined, redundant circuits are required not to be assembled in order to reduce weight. Subsequently, other individual subsystems proposed to increase the demand for telemetry acquisition. Since the original module design still has a margin in the number of A / D acquisitions, it can be achieved by simply supplementing the assembly requirements, and the design can remain unchanged.

[0056] The spacecraft integrated electronic system construction method provided by the embodiments of the present invention has the following advantages: on the one hand, the constructed spacecraft integrated electronic system has high functional integration, high resource reuse, good generalization and strong adaptability to demand; on the other hand, the scheme has high efficiency in constructing the spacecraft integrated electronic system.

[0057] Figure 6 This is a structural block diagram of a spacecraft integrated electronic system construction device provided in an embodiment of the present invention.

[0058] like Figure 6 As shown, the spacecraft integrated electronic system construction device of this embodiment includes: a first determining module 601, used to determine the functional modules included in the spacecraft integrated electronic system, and the number of functions or interfaces corresponding to each functional module; a construction module 602, used to construct the interface specifications between the functional modules; a second determining module 603, used to determine the number of integrated electronic units included in the spacecraft and the functional range of each integrated electronic unit; an iteration module 604, used to iterate the functional range and functional modules of each integrated electronic unit; and a construction module 605, used to select target functional modules to combine into integrated electronic units of each partition according to the current functional and interface quantity requirements, thereby constituting the spacecraft integrated electronic system.

[0059] The spacecraft integrated electronic system construction device provided in this embodiment of the invention has the following advantages: on the one hand, the constructed spacecraft integrated electronic system has high functional integration, high resource reuse, good generalization and strong adaptability to demand; on the other hand, the device has high efficiency in constructing the spacecraft integrated electronic system.

[0060] It should be noted that the above description is only a preferred embodiment of the present invention. It should be understood that those skilled in the art can make several changes and improvements without departing from the technical concept of the present invention, and these are all included within the protection scope of the present invention.

[0061] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for constructing an integrated electronic system for a spacecraft, characterized in that, The method includes: Determine the functional modules included in the spacecraft's integrated electronic system, and the number of functions or interfaces corresponding to each functional module; Establish interface specifications between the various functional modules; Determine the number of integrated electronic units contained in the spacecraft and the functional scope of each integrated electronic unit; The functional scope and functional modules of each integrated electronic unit are iterated; Based on the current functional and interface requirements, target functional modules are selected and combined into integrated electronic units for each zone to form the spacecraft integrated electronic system. The steps for iterating the functional scope and functional modules of the integrated electronic unit include: The functional modules are stacked and interlocked, and the internal bus signals are transmitted layer by layer through the interlocking of functional modules. The signals corresponding to the same contact number are kept consistent, so as to facilitate the arbitrary interlocking and compatibility between functional modules and the rapid combination and change of integrated electronic units. The steps for determining the functional modules included in the spacecraft's integrated electronic system include: The spacecraft's integrated electronic system is categorized into functional modules based on three main categories: power supply, central processing, and interface processing. The power supply category includes high-power power supply modules and low-power power supply modules; the central processing category includes high-performance processor modules, medium- and low-performance processor modules, and bus communication functional modules; and the interface processing category is divided according to interface type. The timing design of the interface between each interface processing functional module and the CPU is implemented using a programmable logic array (FPGA). By injecting different versions of FPGA software, timing matching with two different processor modules is achieved. The maximum number of modules in a single integrated electronic unit shall not exceed 12 and the minimum number shall not be less than 5, and the management scope shall not span multiple compartments.

2. The method according to claim 1, characterized in that, The interface specifications include: mechanical interface specifications, electrical interface specifications, and software interface specifications.

3. The method according to claim 2, characterized in that: The mechanical interface specifications include uniformly defined dimensions of the printed circuit board, dimensions of the printed circuit board metal frame, dimensions of the front and back envelopes of the assembled printed circuit board, model and layout of the internal bus connectors used for inter-module insertion, dimensions of the machine cover plate, and dimensions of the machine base plate.

4. The method according to claim 2, characterized in that, The electrical interface specifications include: The definitions of bus connector contact signals, electrical characteristics such as signal voltage and current, and signal timing are uniformly defined between modules.

5. The method according to claim 1, characterized in that, The step of determining the number of integrated electronic units contained in the spacecraft and the functional range of each integrated electronic unit includes: Based on the functional requirements of different zones of the spacecraft, determine the number of integrated electronic units included in the spacecraft and the functional scope of each integrated electronic unit.

6. The method according to claim 1, characterized in that, The integrated electronic unit includes a main control unit and an integrated electronic remote control unit; The main control unit uses a high-performance central processing module for overall machine control, and selects a high-power or medium-low power supply module according to the overall power consumption requirements of the machine. The main control unit forwards remote control commands and sends program control commands to each of the integrated electronic remote control units via the bus, and collects telemetry information, health information, and backup information from each of the integrated electronic remote control units.

7. The method according to claim 6, characterized in that, The integrated electronic remote control unit uses a low-to-medium power central processing module for overall control, and selects a high-power or low-to-medium power supply module according to the overall power consumption requirements.

8. A spacecraft integrated electronic system construction device, characterized in that, The device includes: The first determining module is used to determine the functional modules included in the spacecraft integrated electronic system, and the number of functions or interfaces corresponding to each functional module; The building module is used to build the interface specifications between the various functional modules; The second determining module is used to determine the number of integrated electronic units contained in the spacecraft and the functional range of each integrated electronic unit; An iteration module is used to iterate the functional scope and functional modules of each integrated electronic unit. The construction module is used to select target functional modules and combine them into integrated electronic units for each partition according to the current functional and interface requirements, thereby forming the integrated electronic system of the spacecraft. The functional scope and functional modules of the integrated electronic unit are iterated, specifically including: The functional modules are stacked and interlocked, and the internal bus signals are transmitted layer by layer through the interlocking of functional modules. The signals corresponding to the same contact number are kept consistent, so as to facilitate the arbitrary interlocking and compatibility between functional modules and the rapid combination and change of integrated electronic units. The functional modules included in the determined spacecraft integrated electronic system specifically include: The spacecraft's integrated electronic system is categorized into functional modules based on three main categories: power supply, central processing, and interface processing. The power supply category includes high-power power supply modules and low-power power supply modules; the central processing category includes high-performance processor modules, medium- and low-performance processor modules, and bus communication functional modules; and the interface processing category is divided according to interface type. The timing design of the interface between each interface processing functional module and the CPU is implemented using a programmable logic array (FPGA). By injecting different versions of FPGA software, timing matching with two different processor modules is achieved. The maximum number of modules in a single integrated electronic unit shall not exceed 12 and the minimum number shall not be less than 5, and the management scope shall not span multiple compartments.

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