Manned spacecraft assembly integrated system design method

Abstract

The invention relates to a design method for an integrated system of a manned spacecraft assembly. The design method comprises the steps that a system and cabin section two-stage assembly is set up; the system generally unifies construction specifications, development processes and production and insurance requirements, and puts forward universal design requirements of each cabin section; according to the system, the top layer design and the system function design of a manned spacecraft assembly are integrally completed, task functions and an index system are decomposed into cabin sections, and interfaces between the cabin sections are defined; overall decomposition of the cabin sections is carried out, functional and technical requirements on subsystems are issued, and interfaces between the subsystems are defined; the subsystem completes subsystem-level design, decomposes and issues function and technical requirements for single machines, and defines interfaces between the single machines; a single machine completes design and development of products according to unified construction specifications and technical requirements. The method can be popularized and applied to design of complex system spacecraft assemblies in various fields, and it is ensured that functions of all cabin sections are fused and coordinated, the product generalization rate is high, the technical state convergence speed is high, and the product maturity is high.

Description

Technical Field

[0001] The invention belongs to the technical field of overall design of manned spacecraft, and relates to an integrated system design method for a manned spacecraft assembly. Background Art

[0002] Spacecraft are undertaking more and more tasks, their functions are becoming more and more complex, and their scale is getting larger and larger. Considering the volume of the launch vehicle fairing and the launch capacity constraints, large and complex spacecraft assemblies represented by manned space stations have gradually become the direction of development of aerospace engineering. Large and complex spacecraft assemblies are generally composed of multiple modules, which are launched into orbit separately, and the construction of the spacecraft assembly is completed through on-orbit assembly and construction.

[0003] Unlike single-body spacecraft launched into orbit, large spacecraft assemblies consisting of multiple compartments have higher requirements for the matching, compatibility and overall efficiency of the compartments. For example, in the International Space Station, which is mainly built by the United States and Russia and 16 countries and regions, the thermal control system of the Russian compartment uses ethylene glycol as the loop working fluid, while the American compartment uses water. The fluid circuits between the compartments cannot be connected, and heat dissipation cannot be uniformly allocated. The Russian compartment is responsible for the unified management of propellants, but the propellant type is different from that of the European cargo spacecraft. As a result, in addition to configuring its own propulsion system, the European automatic transfer vehicle cargo ship also needs to be specially equipped with additional tanks for loading the fuel used by the Russian side, which increases the complexity and cost of the system. In addition, the spacecraft assembly is large in scale, with many supporting products and many participating research units. Without unified design rules, verification planning and test plans, it is bound to lead to slow determination of technical status, low development efficiency and high management costs.

[0004] Therefore, it is very necessary to design an integrated system design method for manned spacecraft assemblies. Summary of the invention

[0005] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art and propose an integrated system design method for a manned spacecraft assembly, so as to solve the problems existing in the prior art such as low matching and compatibility between compartments of a large spacecraft assembly composed of multiple compartments, slow determination of technical status, low development efficiency, and high management cost.

[0006] The technical solution of the present invention is: a method for designing an integrated system of a manned spacecraft assembly, comprising the following steps:

[0007] Establish two levels of system and compartment, and determine the four-level technical status control architecture of system, compartment, subsystem and single machine;

[0008] The system uniformly formulates construction specifications, development processes and production and maintenance requirements, and proposes universal design requirements for each compartment;

[0009] The system overall completes the top-level design of the manned spacecraft assembly and the system function design, decomposes the mission functions and indicator systems into each module, and clarifies the interfaces between modules;

[0010] The overall design of the cabin is completed according to the construction specifications, development process, production and maintenance requirements, universal design requirements of each cabin, and the task functions and indicator system of each cabin decomposed by the overall system. The functional and technical requirements of the subsystems are decomposed and issued, and the interfaces between subsystems are clarified.

[0011] The subsystem is designed according to the construction specifications, development process, production and maintenance requirements uniformly formulated by the overall system, as well as the functional and technical requirements of the overall compartment decomposed into subsystems, decomposes and issues functional and technical requirements for individual units, and clarifies the interfaces between individual units;

[0012] The design and development of the single machine product is completed based on the overall unified construction specifications, development process, production and maintenance requirements of the system, as well as the functional and technical requirements of the subsystem decomposed into the single machine.

[0013] Furthermore, the construction specifications are a complete set of research and development standardization system, which is transmitted step by step from system to compartment to subsystem to single machine, including: mechanical, power supply and distribution, signal interface, heat, electromagnetic compatibility, ergonomics, optics, testability, maintainability, reliability, safety, software, space environment, components and product environmental testing.

[0014] Furthermore, the system function design includes sealing and bearing, attitude and orbit control, energy, information, thermal management, manned environment, assembly and construction, extravehicular operations, astronaut residence support, and experimental application guarantee.

[0015] Furthermore, the interfaces between modules include rendezvous and docking interfaces, assembly management interfaces, assembly and construction interfaces, and ground joint test project interfaces.

[0016] Furthermore, the rendezvous and docking interface stipulates the rendezvous and docking equipment configuration, technical indicators, layout and field of view requirements, docking initial conditions, docking mechanical environment, launch window, orbit characteristics, docking mode, flight attitude, ground-to-space measurement and control communication frequency band, air-to-air communication interface, target optical characteristics and surface potential control;

[0017] The assembly management interface stipulates assembly configuration, quality characteristics, leakage rate requirements, assembly environment control, assembly residence support, assembly attitude and orbit control, assembly energy and power supply and distribution management, information system management, on-orbit complex operation task interface, maintainability and expansion interface;

[0018] The assembly and construction interface stipulates the construction of key parking positions, attitude orbits and solar wing rotation control during construction, assembly actuator movement timing, and assembly state settings after assembly and construction;

[0019] The interface of the ground joint test project is divided into two modes according to the test status of the cabin: if the entire aircraft participates in the test, the matching of the electrical interface and information interface between the cabins and the correctness of the integrated control of the energy system and information system are verified, the integrated use of sensors and actuators and the matching of the control interface of the main control cabin for the attitude of the assembly are verified, the matching of the electrical interface of the crawling robot between the cabins is verified, and the correctness of the flight program in the dynamic mission is verified; if the entire aircraft is not used to participate in the test, but a special test is carried out through the cooperation of key equipment on the cabin and ground tooling, then the accuracy of the dynamic parameter setting of the key equipment under the simulation of real on-orbit action, the timing correctness of the key equipment under the simulation of real on-orbit action, and the coordination of on-orbit mission information interaction are verified.

[0020] Furthermore, the interface between subsystems includes an electrical interface with the power supply and distribution subsystem, a signal interface with the control subsystem, a remote control and telemetry interface with the measurement, control and communication subsystem, a command and parameter interface with the digital control subsystem, and a thermal interface with the thermal control subsystem.

[0021] Furthermore, the overall design of the module also includes: according to the unique tasks and functional characteristics of the module, the module overall organization carries out targeted special design work. Among them, the subsequent modules mainly used for in-orbit experiments focus on special projects such as assembly and construction, combination control and management, and experimental payload support.

[0022] Furthermore, the assembly and construction project includes the transfer scheme of the rear launch module, the transfer attitude control scheme, the transfer mode of different actuators, the transfer measurement and control coverage, and the transfer energy balance;

[0023] The said complex control and management project includes complex attitude and orbit control management, complex energy system management, complex information system management, complex thermal management, and complex manned environment management;

[0024] The experimental payload support project includes in-cabin payload support interface, out-cabin standard payload support interface, out-cabin expandable payload support interface, and microgravity level.

[0025] Furthermore, the attitude and orbit control management of the combination is mainly based on the attitude and orbit control computer of the main control cabin, and the attitude and orbit control computer of the backup function cabin is used as a supplement to realize the integrated control of the attitude and orbit of the combination and the rotation control of the solar wing, and the integrated use of sensors and actuators of each cabin is realized through the cabin bus;

[0026] The energy system management of the combination is to connect the energy system between cabins to the grid, and multiple independent busbars run through the cabins to realize the establishment of the space station energy system, with the main control cabin as the main part and the backup function cabin as the backup to perform energy management of the combination;

[0027] The information system management of the complex is that the information systems of each cabin are connected to the network, information is transmitted between cabins, and the main control cabin manages the complex information system in a unified manner; when the ground-to-ground link of the main control cabin fails, the complex realizes ground-to-ground data communication of the complex through the ground-to-ground links of other cabins; when the core processing unit of the data control subsystem of the main control cabin fails, the complex enters the minimum safety mode of the entire station, and the core processing unit of the data control subsystem of the backup function cabin manages the complex;

[0028] The combined thermal management is to allocate heat to other compartments through the inter-compartment coupling circuit system when a circuit failure occurs in a compartment or the heat load level exceeds the limit, so as to realize inter-compartment joint heat dissipation;

[0029] The manned environment management of the combination is to enable the backup function cabin manned environment system to control the manned environment of the combination when the entire station is adjusting the working mode, or the manned environment system of the main control cabin fails and cannot be repaired.

[0030] The beneficial effects of the present invention compared with the prior art are:

[0031] (1) The present invention develops an integrated system design method for manned spacecraft assemblies, which can be extended and applied to complex system spacecraft assemblies in various fields;

[0032] (2) The integrated system design method of the present invention ensures the integration and coordination of the functions of each compartment, high product commonality, fast technical status convergence, and high product maturity;

[0033] (3) The present invention realizes the ordering of the development process, the mass production of products, and the maximization of quality, benefit and efficiency, which greatly promotes the product quality, development progress and cost control of complex system spacecraft assemblies. BRIEF DESCRIPTION OF THE DRAWINGS

[0034] Figure 1 A flow chart of the integrated system design method for a manned spacecraft assembly of the present invention;

[0035] Figure 2 It is an application case of the manned space station integrated system design method involved in the specific implementation mode of the present invention. DETAILED DESCRIPTION

[0036] In order to solve the problems existing in the prior art, such as the low compatibility between the cabins of a large spacecraft assembly composed of multiple cabins, slow determination of technical status, low development efficiency, and high management costs, the present invention proposes a manned spacecraft assembly integrated system design method, which is used for the integrated design and development of a large spacecraft assembly composed of multiple cabins. Through the integration of resources and functions of each cabin, a large spacecraft assembly with complete functions and powerful performance is jointly formed, realizing the three-level integration of equipment level, function level, and cabin level. This method realizes the unification of product interfaces, orderly development process, batch production of products, and maximization of quality, benefit and efficiency. It has a great promoting effect on accelerating the convergence of the technical status of complex spacecraft systems, improving development efficiency and product maturity, ensuring development quality, guaranteeing progress, and reducing costs.

[0037] The present invention proposes a manned spacecraft assembly integrated system design method, which is used for the integrated design and development of a large spacecraft assembly consisting of multiple cabins, such as Figure 1 As shown, the following steps are included:

[0038] Step 1: Establish the system and compartment levels, and determine the four-level technical status control architecture of system, compartment, subsystem, and single machine;

[0039] Step 2: The overall system uniformly formulates construction specifications, development processes and production and maintenance requirements, and proposes universal design requirements for each compartment;

[0040] Step 3: The system overall completes the top-level design of the manned spacecraft assembly and the system function design, decomposes the mission functions and indicator system into each module, and clarifies the interface between modules;

[0041] Step 4: The overall design of the cabin is completed according to the construction specifications, development process, production and maintenance requirements, universal design requirements of each cabin, and the task functions and indicator system of each cabin that are decomposed by the overall system. The functional and technical requirements of the subsystems are decomposed and issued, and the interfaces between the subsystems are clarified.

[0042] Step 5: The subsystem is designed according to the unified construction specifications, development process, production and maintenance requirements of the system as a whole, and the functional and technical requirements of the cabin section as a whole. The functional and technical requirements of the single machine are decomposed and issued, and the interface between the single machines is clarified;

[0043] Step 6: Complete the design and development of the product based on the overall unified construction specifications, development process, production and maintenance requirements of the system, and the functional and technical requirements of the subsystem decomposed into the single machine. Among them, the universal product is uniformly tested and verified; the deliverables include physical products and digital models.

[0044] In step one: according to the idea of ​​combining top-level design with engineering management, promoting management with technology and innovation with management, the manned spacecraft assembly is set up as a system and a cabin, and a four-level technical status control architecture is implemented according to the system, cabin, subsystem, and single machine. Figure 2 shown.

[0045] In step two: the system will uniformly formulate overall construction specifications, development processes and production assurance requirements applicable to manned spacecraft assemblies, and propose universal design requirements for each module.

[0046] Standardize the design, testing and functional verification of universal products to avoid repeated identification, accelerate the convergence of technical status, and improve development efficiency and product maturity. The construction specification is a complete development standardization system, which is passed down from system to compartment to subsystem to single machine, including: machinery, power supply and distribution, signal interface, heat, electromagnetic compatibility, ergonomics, optics, testability, maintainability, reliability, safety, software, space environment, components and product environmental testing.

[0047] Products with the same mission and functional requirements in each compartment are designed in accordance with universal requirements, and their test verification is evaluated according to the most stringent test conditions in each compartment, so that all aircraft can be covered in one test.

[0048] In step three: the system is responsible for decomposing the user's technical and management requirements, completing the top-level design of the manned spacecraft assembly, and completing the system function design in terms of sealing load, attitude and orbit control, energy, information, thermal management, manned environment, assembly and construction, extravehicular operations, astronaut residence support, experimental application guarantee, etc. On this basis, the mission function indicator system of the manned spacecraft assembly is constructed, and the above system functions and performance indicators are clarified to each module, and the interfaces between modules are clarified.

[0049] The interfaces between modules mainly include rendezvous and docking interfaces, assembly management interfaces, assembly and construction interfaces, and ground joint test project interfaces. For the rendezvous and docking interface, the key points include the configuration of rendezvous and docking equipment, technical indicators, layout and field of view requirements, initial docking conditions, docking mechanical environment, launch window, orbit characteristics, docking mode, flight attitude, ground-to-space measurement and control communication frequency band, air-to-air communication interface, target optical characteristics, and surface potential control. For the assembly management interface, the key points include assembly configuration, quality characteristics, leakage rate requirements, assembly environment control, assembly residence support, assembly attitude and orbit control, assembly energy and power supply and distribution management, information system management, on-orbit complex operation task interface, maintainability, and expansion interface. For the assembly and construction interface, the key points include the construction of key parking positions, attitude orbit and solar wing rotation control during construction, assembly actuator motion timing, and assembly state setting after assembly and construction. For the ground joint test project interface, it is divided into two modes according to the test status of the module. If the entire aircraft is used in the test, the focus will be on verifying the matching of the electrical and information interfaces between the cabins, the correctness of the integrated control of the energy system and information system, the integrated use of sensors and actuators, and the matching of the control interface of the main control cabin for the attitude of the assembly, the matching of the electrical interface of the crawling robot arm between the cabins, and the correctness of the flight program in the dynamic mission. If the entire aircraft is not used in the test, but special tests are carried out through the cooperation of key equipment on the cabin and ground tooling, such as the special test of the robot arm transfer mission, the focus will be on verifying the accuracy of the dynamic parameter settings of key equipment under the simulation of real on-orbit actions, the timing correctness of key equipment under the simulation of real on-orbit actions, and the coordination of on-orbit mission information interaction.

[0050] In step 4: the overall cabin design is completed according to the technical requirements issued by the overall system, the functional and technical requirements of the subsystems are decomposed and issued, and the interfaces between the subsystems are clarified.

[0051] The interfaces between subsystems mainly include: electrical interface with power supply and distribution subsystem, signal interface with control subsystem, remote control and telemetry interface with measurement and control communication subsystem, command and parameter interface with digital control subsystem, thermal interface with thermal control subsystem, etc. On the other hand, the overall organization of the cabin carries out targeted special design work in view of the unique tasks and functional characteristics of the cabin. For example, the rear cabin, which is mainly based on on-orbit experiments, should focus on special designs such as assembly construction, assembly control and management, and experimental payload support. For the assembly construction project, it mainly includes the transfer scheme of the rear cabin, the transfer attitude control scheme, the transfer mode of different actuators, the transfer measurement and control coverage, and the transfer energy balance. For the experimental payload support project, it mainly includes the cabin load support interface, the extravehicular standard load support interface, the extravehicular expandable load support interface, and the microgravity level. For the assembly control and management project, it mainly includes the assembly attitude orbit control management, the assembly energy system management, the assembly information system management, the assembly thermal management, and the assembly manned environment management.

[0052] Taking the manned space station as an example, the attitude and orbit control management of the complex is mainly based on the attitude and orbit control computer of the main control cabin, and the attitude and orbit control computer of the backup function cabin is used as a supplement to realize the integrated control of the complex attitude and orbit and the rotation control of the solar wing. The integrated use of sensors and actuators in each cabin is realized through the cabin bus. The energy system management of the complex is the grid connection of the energy system between cabins. Multiple independent buses run through the cabins to realize the establishment of the space station energy system. The main control cabin is the main part and the backup function cabin is the backup for the energy management of the complex. The information system management of the complex is the grid connection and interconnection of the information systems of each cabin. Information is transmitted between cabins, and the main control cabin manages the information system of the complex in a unified manner. When the ground-to-ground link of the main control cabin fails, the complex realizes the ground-to-ground data communication of the complex through the ground-to-ground links of other cabins. When the core processing unit of the data control subsystem of the main control cabin fails, the complex enters the minimum safety mode of the entire station, and the core processing unit of the data control subsystem of the backup function cabin manages the complex. The combined thermal management is to allocate heat to other compartments through the inter-cabin coupling circuit system when a circuit failure occurs in a compartment or the heat load level exceeds the limit, so as to achieve inter-cabin joint heat dissipation. The combined manned environment management is to enable the backup function compartment manned environment system to control the combined manned environment when the whole station is adjusting the working mode or the main control compartment manned environment system fails and cannot be repaired.

[0053] In step five: each subsystem forms a subsystem-level design plan based on the overall technical requirements of the cabin, completes the interface design between the units under its jurisdiction, and decomposes and issues the functional and technical requirements of each unit. In the entire combination, the subsystems of each cabin maintain unity, and uniformly carry out professional design across aircraft to ensure the integration of each cabin and the universalization of cross-cabin products.

[0054] In step six: the single machine completes the design and development of the product according to the unified construction specifications and technical requirements issued by the subsystems, which are passed down from top to bottom, and the universal products are uniformly tested and verified. The products of each compartment are delivered to the subsystems and compartments step by step. In addition to delivering physical products, the single machine formulates the digital model of each device according to the unified digital model interface and unified modeling requirements, and submits the overall integration into the digital model of the whole cabin after testing and verification.

[0055] The present invention will be further described below in conjunction with the embodiments.

[0056] Example 1

[0057] Take the manned space station as an example to power the payload equipment:

[0058] 1) In the design of the space station system, the overall power supply design for the payload equipment of the space station system is: during the three-cabin combination, the steady-state power provided by the Wentian Experimental Module for the payload is 5kW, of which the power consumption of the in-cabin payload does not exceed 5kW; the power consumption of the out-cabin payload does not exceed 3kW. During the three-cabin combination, under the condition that the platform energy balance allows, through the inter-cabin payload power allocation between the Wentian Experimental Module and the Mengtian Experimental Module, the peak power consumption of the Wentian Experimental Module for the applied payload is up to 7kW, of which the power consumption of the in-cabin payload does not exceed 5kW; the power consumption of the out-cabin payload does not exceed 3kW.

[0059] 2) This design is transformed into the overall technical requirements of the space station system for the Wentian laboratory module: after the Wentian laboratory module is docked with the core module to form a combination, under nominal conditions, the Wentian laboratory module has the ability to supply power to the loads inside and outside the cabin of no less than 5kW; after forming a T-shaped three-cabin combination configuration, with the cooperation of the Mengtian laboratory module, through inter-cabin power supply allocation, the Wentian laboratory module should have the ability to provide a maximum power supply of 7kW for the payload, of which the maximum power supply capacity inside the cabin is 5kW, and the maximum power supply capacity outside the cabin is 3kW.

[0060] 3) The overall requirements of the Wentian Experimental Module for the space station system are implemented in the "Payload Power Distribution Support" section of the design report, which is described as follows: The power distribution system provides 5kW of steady-state total power support for the Wentian Experimental Module payload, of which the maximum power of the in-cabin payload is 5KW and the maximum power of the out-cabin payload is 3KW. Under the condition that the total power limit of the three-cabin payload remains unchanged, it is allowed to temporarily deploy the 2kW power supply of the Mengtian Experimental Module, so that the maximum total power supported by the Wentian Experimental Module is 7kW.

[0061] 4) Since this design is applicable to both experimental cabins and is a universal design, both the Wentian experimental cabin and the Mengtian experimental cabin have issued this requirement to the power distribution subsystem. The description is: provide 100V primary bus power supply, the steady-state power supply capacity of a single cabin is not less than 5kW, and can support dynamic allocation of load power to a peak power of 7kW.

[0062] 5) After receiving the overall technical requirements issued by the two experimental cabins, the power distribution subsystem completed the subsystem design and put forward relevant technical requirements to the single-machine units. The above content is described as follows: The power distribution subsystem is equipped with X X-type power control units (single unit rated at 5KW) and X XX-type power control units (rated power 3kW) to provide power for the load equipment.

[0063] 6) The single-machine unit completes the design of the X-type power control unit and the XX-type power control unit according to the overall unified construction specifications issued by the space station system and the technical requirements issued by the power distribution subsystem. For universal products, unified test verification is carried out.

[0064] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

[0065] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A method for designing an integrated system of a manned spacecraft assembly, characterized in that: The following steps are involved: Establish two levels of system and compartment, and determine the four-level technical status control architecture of system, compartment, subsystem and single machine; The system uniformly formulates construction specifications, development processes and production and maintenance requirements, and proposes universal design requirements for each compartment; The system overall completes the top-level design of the manned spacecraft assembly and the system function design, decomposes the mission functions and indicator systems into each module, and clarifies the interfaces between modules; The overall design of the cabin is completed according to the construction specifications, development process, production and maintenance requirements, universal design requirements of each cabin, and the task functions and indicator system of each cabin decomposed by the overall system. The functional and technical requirements of the subsystems are decomposed and issued, and the interfaces between subsystems are clarified. The subsystem is designed according to the construction specifications, development process, production and maintenance requirements uniformly formulated by the overall system, as well as the functional and technical requirements of the overall compartment decomposed into subsystems, decomposes and issues functional and technical requirements for individual units, and clarifies the interfaces between individual units; The design and development of the single machine product is completed based on the overall unified construction specifications, development process, production and maintenance requirements of the system, as well as the functional and technical requirements of the subsystem decomposed into the single machine.

2. A method for designing an integrated system of a manned spacecraft assembly according to claim 1, characterized in that: The construction specifications are a complete set of research and development standardization systems, which are passed down step by step from system to compartment to subsystem to single machine, including: mechanical, power supply and distribution, signal interface, heat, electromagnetic compatibility, ergonomics, optics, testability, maintainability, reliability, safety, software, space environment, components and product environmental testing.

3. The method for designing an integrated system of a manned spacecraft assembly according to claim 1, characterized in that: The system functional design includes sealing and bearing, attitude and orbit control, energy, information, thermal management, manned environment, assembly and construction, extravehicular operations, astronaut residence support, and experimental application guarantee.

4. The method for designing an integrated system of a manned spacecraft assembly according to claim 1, characterized in that: The interfaces between the modules include rendezvous and docking interfaces, combination management interfaces, assembly and construction interfaces, and ground joint test project interfaces.

5. The method for designing an integrated system of a manned spacecraft assembly according to claim 4, characterized in that: The rendezvous and docking interface stipulates the rendezvous and docking equipment configuration, technical indicators, layout and field of view requirements, docking initial conditions, docking mechanical environment, launch window, orbit characteristics, docking mode, flight attitude, ground-to-space measurement and control communication frequency band, air-to-air communication interface, target optical characteristics and surface potential control; The assembly management interface stipulates assembly configuration, quality characteristics, leakage rate requirements, assembly environment control, assembly residence support, assembly attitude and orbit control, assembly energy and power supply and distribution management, information system management, on-orbit complex operation task interface, maintainability and expansion interface; The assembly and construction interface stipulates the construction of key parking positions, attitude orbits and solar wing rotation control during construction, assembly actuator movement timing, and assembly state settings after assembly and construction; The interface of the ground joint test project is divided into two modes according to the test status of the cabin: if the entire aircraft participates in the test, the matching of the electrical interface and information interface between the cabins and the correctness of the integrated control of the energy system and information system are verified, the integrated use of sensors and actuators and the matching of the control interface of the main control cabin for the attitude of the assembly are verified, the matching of the crawling electrical interface of the robotic arm between the cabins is verified, and the correctness of the flight program in the dynamic mission is verified; if the entire aircraft is not used to participate in the test, but a special test is carried out through the cooperation of key equipment on the cabin and ground tooling, then the accuracy of the dynamic parameter setting of the key equipment under the simulation of real on-orbit action, the timing correctness of the key equipment under the simulation of real on-orbit action, and the coordination of on-orbit mission information interaction are verified.

6. The method for designing an integrated system of a manned spacecraft assembly according to claim 1, characterized in that: The interfaces between subsystems include an electrical interface with the power supply and distribution subsystem, a signal interface with the control subsystem, a remote control and telemetry interface with the measurement, control and communication subsystem, a command and parameter interface with the digital control subsystem, and a thermal interface with the thermal control subsystem.

7. The method for designing an integrated system of a manned spacecraft assembly according to claim 1, characterized in that: The overall design of the module also includes: according to the unique tasks and functional characteristics of the module, the module overall organization carries out targeted special design work. Among them, the subsequent modules mainly used for in-orbit experiments focus on special projects such as assembly and construction, combination control and management, and experimental payload support.

8. The method for designing an integrated system of a manned spacecraft assembly according to claim 7, characterized in that: The assembly and construction project includes the transfer scheme of the rear launch module, the transfer attitude control scheme, the transfer modes of different actuators, the transfer measurement and control coverage, and the transfer energy balance; The said complex control and management project includes complex attitude and orbit control management, complex energy system management, complex information system management, complex thermal management, and complex manned environment management; The experimental payload support project includes in-cabin payload support interface, out-cabin standard payload support interface, out-cabin expandable payload support interface, and microgravity level.

9. The method for designing an integrated system of a manned spacecraft assembly according to claim 8, characterized in that: The attitude and orbit control management of the combination is mainly based on the attitude and orbit control computer of the main control cabin, and the attitude and orbit control computer of the backup function cabin is used as a supplement to realize the integrated control of the attitude and orbit of the combination and the rotation control of the solar wing, and the integrated use of sensors and actuators of each cabin is realized through the cabin bus; The energy system management of the combination is to connect the energy system between cabins to the grid, and multiple independent busbars run through the cabins to realize the establishment of the space station energy system, with the main control cabin as the main part and the backup function cabin as the backup to perform energy management of the combination; The information system management of the complex is that the information systems of each cabin are connected to the network, information is transmitted between cabins, and the main control cabin manages the complex information system in a unified manner; when the ground-to-ground link of the main control cabin fails, the complex realizes ground-to-ground data communication of the complex through the ground-to-ground links of other cabins; when the core processing unit of the data control subsystem of the main control cabin fails, the complex enters the minimum safety mode of the entire station, and the core processing unit of the data control subsystem of the backup function cabin manages the complex; The combined thermal management is to allocate heat to other compartments through the inter-compartment coupling circuit system when a circuit failure occurs in a compartment or the heat load level exceeds the limit, so as to realize inter-compartment joint heat dissipation; The manned environment management of the combination is to enable the backup function cabin manned environment system to control the manned environment of the combination when the entire station is adjusting the working mode, or the manned environment system of the main control cabin fails and cannot be repaired.

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