A general test platform for GNC system of multi-cabin spacecraft

Abstract

The application provides a universal test platform of a multi-cabin spacecraft GNC system, relates to spacecraft GNC data processing, and comprises the following modules: an application software module suitable for a multi-cabin spacecraft, a dynamics software module, a component interface simulation module, an injection management module and a running monitoring module, wherein the application software module is used for universal modification of on-board application software so that the on-board application software communicates data with the remaining modules; the dynamics software module is used for updating state information according to control instructions; the component interface simulation module is used for data communication between the application software module and the dynamics software module; the injection management module is used for controlling the universal test platform to complete simulation tests according to test instructions; and the running monitoring module is used for monitoring telemetry data output by the application software module and the dynamics software module. The scheme can solve the problems of long simulation verification time and low efficiency of a traditional hardware platform for a multi-cabin spacecraft.

Description

Technical Field

[0001] This invention relates to the field of spacecraft GNC data processing technology, and in particular to a universal test platform for a multi-module spacecraft GNC system. Background Technology

[0002] GNC is an abbreviation for Guidance, Navigation and Control. It represents a flight system. Unlike traditional spacecraft models, large on-orbit assembled spacecraft have many modules, complex configurations, high control difficulty, and complex redundancy and reconfiguration strategies. Therefore, the design and experimental verification of multi-module GNC subsystems of large spacecraft need to be carried out in parallel.

[0003] In related technologies, the commonly used single-module system test equipment cannot support the concurrent testing and verification of multiple configurations, and it is also difficult to achieve accelerated testing and verification, which greatly reduces the efficiency of software development and verification for multi-module spacecraft.

[0004] Therefore, there is an urgent need for a universal testing platform for multi-module spacecraft GNC systems to solve the above-mentioned technical problems. Summary of the Invention

[0005] This invention provides a universal test platform for GNC systems of multi-module spacecraft, which can solve the problems of long simulation verification time and low efficiency of traditional hardware platforms for multi-module spacecraft.

[0006] In a first aspect, embodiments of the present invention provide a universal test platform for a multi-module spacecraft GNC system, including an application software module, a dynamics software module, a component interface simulation module, an injection management module, and an operation monitoring module suitable for multi-module spacecraft, wherein:

[0007] The application software module is determined based on the on-board application software of the multi-module spacecraft being tested. The application software module is used to make generalized modifications to the input and output interfaces of the on-board application software so that the on-board application software can communicate with the other modules of the general test platform.

[0008] The dynamics software module is equipped with a standard dynamics software interface and an integrated configuration parameter management submodule. The dynamics software module is used to update the current state of the multi-module spacecraft actuators according to the control commands given by the application software module, thereby updating the current dynamic state of the multi-module spacecraft. Based on the updated dynamic state, the module updates the state of the ground simulation sensors and transmits the updated state information to the application software module through the component interface simulation module. The state information includes the spacecraft's attitude, position, velocity, and other operational information, as well as its combined configuration information.

[0009] The component interface simulation module encapsulates multiple general component simulation docking functions, and the component interface simulation module is used to enable data communication between the application software module and the dynamics software module.

[0010] The injection management module is communicatively connected to the application software module, the dynamics software module, and the component interface simulation module, respectively. The injection management module is used to control the general test platform to complete the simulation test according to the test command.

[0011] The operation monitoring module is used to monitor the telemetry data output by the application software module and the dynamics software module, and to switch the configuration of the multi-module spacecraft according to control commands.

[0012] This invention provides a universal test platform for a multi-module spacecraft GNC system. This platform integrates onboard computer and ground dynamics software, enabling direct embedded verification of onboard application software and allowing for synchronous updates and parallel verification on both a system test semi-physical simulation platform and a rapid simulation platform. Furthermore, it provides integrated management of the combined configuration parameters, enabling free switching and concurrent testing and verification of multiple configurations. Simultaneously, it utilizes a universal test platform to implement a multi-level fault injection method, ensuring complete verification of various fault modes. This method shortens the overall development cycle of the multi-module software test platform and solves the problems of long simulation verification times and low efficiency associated with traditional hardware platforms. Attached Figure Description

[0013] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0014] Figure 1 This is a schematic diagram of a general test platform for a multi-module spacecraft GNC system provided in an embodiment of the present invention;

[0015] Figure 2 This is an initialization scenario flowchart provided by an embodiment of the present invention. Detailed Implementation

[0016] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are some embodiments of the present invention, but not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0017] As mentioned earlier, existing hardware and software platforms generally only simulate single-module spacecraft configuration testing. This method cannot support concurrent testing and verification of multiple configurations, nor can it achieve multi-module spacecraft rate acceleration testing and verification.

[0018] Based on this, the concept of the present invention is to enable the hardware platform and simulation platform to carry out multi-section testing in parallel through the design ideas of software module generalization and parameter integration.

[0019] The specific implementation of the above concept is described below.

[0020] Please refer to Figure 1 This invention provides a universal test platform for a multi-module spacecraft GNC system, the platform comprising:

[0021] Application software modules, dynamics software modules, component interface simulation modules, injection management modules, and operation monitoring modules suitable for multi-module spacecraft, including:

[0022] The application software module is determined based on the on-board application software of the multi-module spacecraft being tested. The application software module is used to make generalized modifications to the input and output interfaces of the on-board application software so that the on-board application software can communicate with the other modules of the general test platform.

[0023] The dynamics software module is equipped with a standard dynamics software interface and an integrated configuration parameter management submodule. The dynamics software module is used to update the current state of the multi-module spacecraft actuators according to the control commands given by the application software module, thereby updating the current dynamic state of the multi-module spacecraft. Based on the updated dynamic state, the module updates the state of the ground simulation sensor and transmits the updated state information to the application software module through the component interface simulation module. The state information includes operational information and combined configuration information.

[0024] The component interface simulation module encapsulates multiple general component simulation docking functions, and the component interface simulation module is used to enable data communication between the application software module and the dynamics software module.

[0025] The injection management module is communicatively connected to the application software module, the dynamics software module, and the component interface simulation module, respectively. The injection management module is used to control the general test platform to complete the simulation test according to the test instructions under different working conditions.

[0026] The operation monitoring module is used to monitor the telemetry data output by the application software module and the dynamics software module, and to switch the configuration of the multi-module spacecraft according to control commands.

[0027] In this embodiment of the invention, the onboard computer and ground dynamics software are integrated, enabling direct embedded verification of onboard application software and allowing for synchronous updates and parallel verification on both the system test semi-physical simulation platform and the rapid simulation platform. Furthermore, the configuration parameters of the combined system are managed in an integrated manner, allowing for free switching between multiple configurations and concurrent testing and verification. Simultaneously, a multi-level fault injection method is implemented using a generalized testing platform to ensure complete verification of various fault modes. This method shortens the overall development cycle of the multi-module software testing platform and solves the problems of long development time and low efficiency in traditional hardware platform simulation verification.

[0028] The following description Figure 1 The modules shown.

[0029] First, the application software modules are designed.

[0030] In the prior art, hardware simulation platforms use on-board control computers and test equipment to conduct semi-physical simulations. The on-board application software algorithms used are consistent with those used in actual orbit. However, the general test platform provided in this embodiment of the invention runs the on-board application software and dynamics software together in a ground test machine. This requires additional processing of the input and output interfaces of the application software modules to achieve communication between platform modules. The interface design of this part is consistent with the general design in the application software modules, and a general modification process is carried out.

[0031] In this embodiment of the invention, the application software module includes an interface input / output logic submodule, a sensor acquisition and information processing submodule, an actuator instruction and information processing submodule, and a validity judgment and fault diagnosis submodule. The application software module has a multi-layered architecture, with each layer independently encapsulated and communicating via standard interfaces. The interface input / output logic submodule, the sensor acquisition and information processing submodule, and the actuator instruction and information processing submodule are all configured with the maximum interface envelope to ensure universality of internal and external interfaces. Parameters and parameter call timing for each submodule are configured according to the functional requirements of each module. The validity judgment and fault diagnosis submodule is used to determine the validity of communication status, self-test status, outliers, and theoretical value ranges, and performs fault diagnosis on assemblies of different configurations and modules according to preset fault diagnosis principles. Faults include component-level faults and system-level faults. Component-level faults are determined based on multiple component measurement data or in conjunction with consistency checks, while system-level faults are determined based on control performance.

[0032] Specifically, the application software module architecture adopts a unified design. The configuration in the application software can meet the requirements of integrated use of multi-module spacecraft in orbit. It adopts a modular, universal, and standardized design, and designs the multi-module platform software architecture into a multi-layer architecture of "general, core, and application". Each layer is independently encapsulated, and inter-layer communication uses standard interfaces. The functional components within each layer can be selected and replaced according to the mission characteristics of each module spacecraft, which improves the maintainability, adaptability, and scalability of the software, minimizes development costs, and improves quality and efficiency.

[0033] Furthermore, based on the interface input / output logic, sensor acquisition and information processing, and actuator command and information processing functional submodules implemented in traditional single-module spacecraft, the "general" functions in the application software platform-based software architecture design of multi-module spacecraft need to adopt the maximum interface envelope to achieve universality of internal and external interfaces. Secondly, configuration parameters should be designed and implemented according to the requirements of each module, such as using compilation parameters, parameter lists, parameter files, etc., to identify the environment and system requirements on which the parameters depend. Finally, the timing of using configuration parameters should be designed to determine when different configuration files or parameters can achieve variability, such as before compilation or before operation.

[0034] Furthermore, for the validity assessment and fault diagnosis submodules, a generalized design can be implemented through the following steps: Commonly used data validity assessments include communication status, self-test status, outliers, and theoretical value range assessments; fault diagnosis is divided into component-level (judging component fault status based on multiple component measurement data or combining consistency checks) and system-level (judging system-level faults based on control performance and ensuring control system safety through fault handling). The fault diagnosis and handling modules for single-cabin flight and combined flight phases are designed uniformly. Component-level fault diagnosis principles remain consistent across all cabins. It is necessary to distinguish that the installation matrices of various sensors and actuators in different cabin sections under different configurations differ in the combined coordinate system. After configuration switching, the installation matrix and other parameters need to be initialized.

[0035] In summary, the generalization of application software modules mainly addresses the standardization of variable type modifications and hardware code masking required when transferring application software code to a software testing platform. This includes standardizing application software modifications for porting to the software platform and standardizing the export of application software global variables used for input / output. Standardization of application software modifications for porting to a general simulation platform can be achieved by comparing the differences between the on-board software and hardware operating environment and the software testing platform operating environment, summarizing standard steps, and establishing a generalized modification process.

[0036] Next, the dynamics software module in the general test platform was designed in a standardized manner.

[0037] In this embodiment of the invention, the dynamics software module is equipped with a multi-module spacecraft attitude and orbit dynamics characteristic model and an environmental model, and establishes simulation models of multi-module sensors and multi-module actuators. Based on the initial state or the current dynamic state, the new state at the next moment is calculated by numerically solving the dynamic equations, including the attitude, position, and velocity states of the multi-module spacecraft, sensor state information, actuator state information, multi-module spacecraft body characteristics, and other state characteristics.

[0038] The dynamics software module can also simulate the dynamic state of the actuators after receiving control commands from the control computer (hardware simulation platform) or application software module, update the position, velocity, sensors, actuators, and other information of the multi-module spacecraft in the ground dynamics, and then transmit the sensor and actuator information back to the control computer (hardware simulation platform) or application software module. The dynamics software sets the initial state of the integrated configuration parameter module or updates the configuration parameter information online based on the multi-module injection management software interface, enabling mode switching between different configurations of the multi-module spacecraft.

[0039] To achieve the above functions, the interfaces of the dynamics software modules need to be standardized. That is, the standardization of control input interfaces, excitation output interfaces, dynamics initialization, and dynamics operation needs to be addressed.

[0040] The standard versions of the control input interface and excitation output interface can be implemented by defining a generic dynamics interface file, which contains array objects that cover the input / output requirements of all configurations of complex spacecraft, thereby covering the input / output requirements of multiple modules; the standardization of dynamics initialization and dynamics operation is achieved by establishing standardized functions.

[0041] Specifically, the dynamics interface file includes a standardized input structure and a standardized output structure. The standardized input structure is defined based on the control commands received by each actuator of the general test platform; the standardized output structure is defined based on the calculated state information of the multi-module spacecraft.

[0042] In other words, the standardized input structure is defined separately according to the characteristics of the instructions received by the specific actuators, including, for example, the start-up duration of each thruster, the angular velocity command of the control torque gyroscope, the frame angular position, the rotor speed, and the command words for the solar panels. The standardized output structure defines the simulation cycle, simulation time system, multi-module mode operating status, orbital position information, attitude information, velocity information, central gravitational field mode, solar model, Earth model, aerodynamic model, mode control status information, center of mass, mass, inertia, and external system simulator data information.

[0043] Furthermore, the standardized functions include dynamic initialization standard functions and dynamic operation standard functions. The dynamic initialization standard functions are determined based on the initialization scenario, module initialization state, and spacecraft initial operation mode; the dynamic operation standard functions are determined based on the operational parameters for data interaction within the multi-module spacecraft.

[0044] In other words, such as Figure 2 As shown, the standard function for dynamic initialization should first select the dynamic initialization scenario, hardware platform, or simulation platform based on the test scenario's operating mode; then, select the initialization module state, whether multi-module or single-module (selecting specific module information); finally, initialize the spacecraft's initial operating mode. After initialization, different initialization functions are generated. The data interaction required by the standard function for dynamic operation includes spacecraft initialization mode, operating mode, actuator command input, dynamic output parameters, sensor output data, injection parameter interface (orbit, attitude, open-loop component data injection, etc.), and dynamic simulation duration.

[0045] It is worth noting that, considering the different operating environments of traditional hardware platforms and the general test platform in this embodiment, the above definition process needs to be defined separately according to the characteristics of each operating system.

[0046] Since this embodiment is applied to the testing scenario of multi-module spacecraft, there are a large number of multi-module spacecraft and configurations. It is difficult to achieve integrated multi-module application by using the traditional method of defining configuration parameters one by one, and it will lead to a lot of redundancy in the configuration switching related code. In order to solve the need to switch configuration parameters from single-module mode to support multi-module testing, this embodiment also designs an integrated configuration parameter management sub-module in the dynamics software module to manage the configuration. This module can be quickly expanded according to the future on-orbit assembly mission requirements of complex spacecraft.

[0047] In this embodiment of the invention, the configuration parameter integrated management submodule performs the following operations when executing configuration parameter conversion: processing the combined state of the spacecraft according to a preset configuration parameter conversion algorithm, and outputting the configuration parameters of each segment of the spacecraft; processing all configuration parameters using a multi-dimensional array to obtain a parameter binding table for storage and retrieval; and indexing the parameter binding table according to the application scenario number, the current segment number, the current configuration number, and the current mode sub-configuration number to conduct simulation experiments on the multi-module spacecraft.

[0048] Specifically, considering multi-module spacecraft, the position / attitude of each module relative to the geometric coordinate system of the single module / combined module can be transformed under any single-module / combined module configuration. Therefore, based on typical configurations of multi-module spacecraft in combined state and visiting spacecraft in parked state, an automatic generation module for the transformation of configuration parameters was designed. This module realizes the automatic transformation and processing of configuration parameters in single-module and combined modes, and optimizes the dynamic initialization parameter calling module. During the development phase, parameters such as the system's center of mass, inertia, sensors, and actuators change significantly with the upgrade of the software under test. Integrated management of configuration parameters can greatly reduce parameter change time and manpower costs.

[0049] In this embodiment of the invention, the automatic generation module for configuration parameter conversion is designed through the following steps:

[0050] S1. Initialize the "Current Aircraft Application Scenario" and "Current Aircraft Segment Number" using the define method. Considering that the application scenario during testing is related to the platform type used, when using a hardware simulation platform for testing, the segment needs to be bound to the currently used hardware platform, and modification is not supported during the testing process; only code changes are allowed. When using a fast simulation platform for testing, the segment supports modification of initialization input.

[0051] S2. Use the #IF conditional compilation method to initialize the selection of binding parameters using "current aircraft application scenario" and "current aircraft module number". The meanings of the module numbers and configuration types in the fast simulation platform scenario and system test scenario are agreed to be consistent with the numbering written in the application software. Furthermore, the names of the binding parameter tables for various sensors / actuators are kept consistent across different application scenarios / aircraft to facilitate the main dynamics program's access to the parameter tables. Different modules are set to different files for easy access.

[0052] S3, the mounting parameter tables for each application scenario / aircraft scenario all use the "module main configuration" to index parameters such as the installation matrix of sensors / actuators; for the body characteristics (mass / inertia / center of mass), the "module sub-mode" indexes the differences in characteristics at different stages of the aircraft.

[0053] S4. For test scenarios of different configurations of the current spacecraft segment, the "Current Spacecraft Dynamics Mode" is set for invocation. The "Current Spacecraft Dynamics Mode" arranges the parameters of dynamic sensors / actuators according to the characteristic locations such as docking interfaces and parking points of the multi-segment spacecraft, creating index numbers for the parameter setting sub-tables of each segment. All parameters are then filled with the physical characteristics of each individual segment. An example list of the meanings of the "Current Spacecraft Dynamics Mode" in each segment's sub-table is as follows:

[0054] Test Module I (Current Spacecraft Segment Number 1):

[0055] 0: Single-cabin parameter 1: Parameter 2: Parameter corresponding to berthing port 1 when docking with interface 1

[0056] Test Module II (current aircraft segment number 2):

[0057] 0: Single-cabin parameter 1: Parameter 2: Parameter corresponding to berthing port 2 when docking with interface 1

[0058] Nth compartment (current aircraft compartment number N):

[0059] 0: Single-cabin parameter 1: Parameter 2: Parameter corresponding to berthing port X when docking with interface X

[0060] Visiting aircraft 1:

[0061] 0: Single-module parameter 1: Parameter 2 when docking with interface 1; ...

[0062] S5. Before the test, determine and set the application scenario (quick simulation platform or hardware platform). Then, through the interface of the dynamics injection software, input the "spacecraft segment number", "segment main configuration" and "segment sub-mode". The corresponding "current spacecraft segment number", "segment main configuration" and "current spacecraft sub-mode" in the dynamics software will be updated. The dynamics software will judge the "current spacecraft dynamics mode" according to the "segment main configuration" and extract the small table information such as the installation position under the corresponding configuration to realize the switching of configuration and mode of multiple segments.

[0063] After obtaining the configuration parameter conversion automatic generation module through the above process, the various parameters, including cabin mass inertia characteristics, sensor installation, actuator installation and output characteristics, and solar panel flexibility, are uniformly processed according to the output results of the configuration parameter automatic generation module. The parameters are stored and retrieved in a multi-dimensional array manner to generate a parameter binding table.

[0064] Next, the parameters of each stage of the multi-module spacecraft dynamics are indexed and accessed using the application scenario number, current module number, current configuration number, and current mode sub-configuration number. Under different application scenarios / spacecraft, the names of the parameter tables for various sensors / actuators remain consistent to facilitate the main dynamics program's access to the parameter tables. The parameter tables for each application scenario / spacecraft scenario are designed, using the main configuration to index the parameters of sensors / actuators, and the mode sub-configuration to index the differences in characteristics at different stages of the spacecraft.

[0065] The following section describes the design of the component interface simulation module in the general test platform.

[0066] The component interface simulation module can also be called the multi-module sensor / actuator general simulation module. During in-orbit flight, the onboard multi-module control computer collects the status information of the onboard multi-module sensors and actuators through the control bus, calculates the position, attitude, and velocity information of the multi-module spacecraft based on the collected status, and issues corresponding control commands. The onboard multi-module actuators control the attitude, position, and velocity of the multi-module spacecraft according to the control commands. In this implementation, the component interface simulation module is mainly used to simulate the communication connection between the multi-module dynamics sensors / actuators and the onboard control computer, that is, the data communication between the aforementioned application software module and the dynamics software module. It also simulates the redundancy layer, communication layer, and protocol layer, and can perform corresponding fault settings and simulations.

[0067] This module can be applied to the intermediate docking platform of the hardware test platform. Based on the integration and sharing of control resources between modules, it can use the functions of dynamic calculation, component excitation and information acquisition, use the equivalent functions of external system interfaces (or through matching between subsystems), and use other types of module equivalents. It can also be applied to the general test platform of this embodiment. By encapsulating multiple general component simulation docking functions, it solves the connection problem between the standardized dynamic interface and the generalized application software interface, as well as the problem of realizing the output channel of application software solar panels, control moment gyroscopes, engine control commands to dynamics, and the excitation channel of sensor measurements obtained from dynamic trajectory and attitude calculations to application software.

[0068] In this embodiment of the invention, the general component simulation docking function includes many types, listed as follows: inertial measurement unit excitation function, solar sensor excitation function, star sensor excitation function, infrared sensor excitation function, and GPS excitation function. Specifically: the inertial measurement unit excitation function accumulates the angular rate and acceleration output by the dynamic inertial measurement unit and performs equivalent conversion to obtain the change within a single control cycle; the star sensor excitation function converts, packages, and frames the raw star sensor data output by the dynamic software according to the excitation requirements of the star sensor to obtain the transmission data required for excitation; the infrared sensor excitation function performs unit conversion on the chord width and angle of incidence data; and the GPS excitation function converts and packages the UTC and GPS position / velocity data output by the dynamic software according to the GPS excitation requirements to obtain the transmission data required for excitation and sends it to the on-board interface device.

[0069] Specifically, 1) Excitation function of the inertial measurement unit:

[0070] The angular rate and acceleration output from the dynamic inertial measurement unit are accumulated and converted to equivalent values ​​to obtain the changes within a single control cycle. Based on the onboard software segment labels, an interface conversion is performed between the dynamics and onboard systems to ensure that the inertial measurement unit data in the dynamics correctly excites the onboard software.

[0071] 2) Excitation function of the solar sensor:

[0072] Based on the onboard software segment labels, an interface conversion is performed between dynamics and the onboard system to ensure that the solar sensor correctly excites the onboard software in the dynamics. For example, based on the projection of the solar vector onto this system, the azimuth information of the 0-1 solar sensor in different directions is calculated. The output value of the 0-1 solar sensor is obtained according to the range of the included angles in different directions, and then the onboard software is excited.

[0073] 3) Star sensor excitation function

[0074] Based on the excitation requirements of the star sensor, the raw star sensor data output from the dynamics system is converted, packaged, and framed to obtain the transmission data required for excitation. According to the onboard software segment labels, interface conversion between the dynamics system and the onboard system is performed to ensure that the star sensor data in the dynamics system correctly excites the onboard software.

[0075] 4) Infrared sensor excitation function

[0076] The units of the chord width and angle of incidence data were converted. Then, according to the onboard software segment labels, the interface conversion between dynamics and onboard was performed to ensure that the infrared data in the dynamics correctly excited the onboard software.

[0077] 5) GPS excitation function

[0078] According to the GPS excitation requirements, the UTC and GPS position / velocity data output by the dynamics are converted and packaged to obtain the transmission data required for excitation, which is then sent to the on-board interface device.

[0079] The following section describes the design of the injection management module in the general testing platform.

[0080] The injection management module is equipped with a dynamic injection data submodule, a general fault injection submodule, and an injection data and command submodule. It sequentially sets the parameters and states of the dynamic software module, application software module, and component interface simulation module, and simultaneously performs multi-module, multi-mode flight simulation switching. It can directly control ground equipment or receive commands from the remote control receiving module of the onboard control computer to generate command control for other products on the satellite or other simulation equipment on the ground, thereby realizing test conditions such as simulated flight testing and simulated testing of multi-level faults and handling.

[0081] In this embodiment of the invention, the general fault injection submodule is used to perform fault simulation injection to verify whether the fault diagnosis and handling strategy of the on-board application software meets the software requirements. The simulation process includes: superimposing the state information output by the dynamics software module to simulate numerical faults; adding an interface layer and a protocol layer to the component interface simulation module to simulate interface faults and protocol faults.

[0082] Specifically, faults can generally be divided into three categories: numerical faults, such as zero drift, noise, constant values, and jumps in sensors and actuators; interface faults, such as communication interruptions and data not being updated; and protocol faults, such as self-test status alarms, checksum errors, and data validity errors. The injection simulation function for the first type of fault can be simulated by superimposing signals on the output values ​​of the sensors / actuators in the dynamic model. For the second and third types of faults, an interface layer and a protocol layer are added to the intermediate simulation docking platform to simulate the interface.

[0083] Finally, the operation monitoring module in the general testing platform is designed.

[0084] The operation monitoring module communicates with the dynamics software module and the application software module through the multi-module display module. On the one hand, this module can receive telemetry information output from the downlink of the on-board application software and decode and display the telemetry data according to the protocol. On the other hand, it can also decode and display the telemetry information calculated and output by the dynamics software module. The display cycle is related to the calculation and operation cycle of the on-board application software and the ground dynamics, thereby realizing data monitoring during the experiment.

[0085] Furthermore, unlike traditional single-compartment dynamics injection, multi-compartment integrated universal test platforms need to have the ability to freely switch configurations. Therefore, the design incorporates an online injection switching function for the platform's dynamics modules through an operational monitoring module, mainly including the following functions: orbit injection, time injection, attitude injection, and configuration switching injection.

[0086] In this embodiment of the invention, configuration switching includes the following steps: determining the primary configuration number and sub-configuration number to be switched based on the parameter data in the parameter binding table, and transmitting the primary configuration number and the sub-configuration number to the dynamics software module; wherein, the primary configuration number is used to determine the spacecraft structure, and the sub-configuration number is used to determine the different on-orbit phase characteristics of a specific configuration; the dynamics software queries and modifies the corresponding injected interface variables according to the number information and enters the configuration switching program; and reassigns the primary configuration number and the sub-configuration number by re-calling the initialization function to update the spacecraft's cabin characteristic information.

[0087] Specifically, the configuration numbering for the configuration switching injection function includes the main configuration and the mode sub-configuration. The main configuration number is consistent with the configuration definition of the application software used on the satellite, including various single modules and various assembly situations, which are used to determine the spacecraft structure. The sub-configuration number is used to determine the different on-orbit stage characteristics of a specific single module / assembly configuration (such as different mass / inertia characteristics before the solar panels are deployed, after the solar panels are deployed once, and after fuel consumption).

[0088] The main program of the dynamics software module modifies the corresponding injected interface variables by querying them, enters the configuration switching program, calls the re-initialization function, enters the main configuration and mode sub-configuration to reassign values, and updates the characteristics of the spacecraft cabin (mass, inertia, center of mass), various sensors, various actuators, solar panel flexibility parameters and aerodynamic parameters, without affecting information such as orbit, time, and attitude.

[0089] In summary, the universal test and verification platform described in this embodiment of the invention offers improved simulation test and verification speed, enabling rapid and flexible problem localization and analysis. It truly realizes the parallel development concept of "designing and testing simultaneously," significantly reducing development costs, shortening the R&D cycle, and improving system reliability. Furthermore, it provides software and hardware support platforms and corresponding technical solutions for model development, holding significant practical importance and application value for the development of multi-module spacecraft attitude control systems. Its coverage of various configurations and operating conditions provides testers with a flexible and scalable software testing platform.

[0090] Furthermore, the aforementioned universal test and verification platform boasts advantages such as versatility and high efficiency, enabling it to adapt to control subsystem configurations of single and multi-module configurations. Moreover, it can rapidly fulfill the requirements for expanded module verification and testing when facing the need for new docking modules in spacecraft, demonstrating broad application prospects. If this platform can be marketed, its convenience, speed, and efficiency are expected to give it strong market competitiveness.

[0091] It should be noted that, in this document, relational terms such as "first" and "second" are used only to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0092] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A generic test platform for multi-cabin spacecraft GNC system, characterized in that, This includes application software modules, dynamics software modules, component interface simulation modules, injection management modules, and operation monitoring modules suitable for multi-module spacecraft, among which: The application software module is determined based on the on-board application software of the multi-module spacecraft being tested. The application software module is used to make generalized modifications to the input and output interfaces of the on-board application software so that the on-board application software can communicate with the other modules of the general test platform. The dynamics software module is equipped with a standard dynamics software interface and an integrated configuration parameter management submodule. The dynamics software module is used to update the current state of the multi-module spacecraft actuators according to the control commands given by the application software module, thereby updating the current dynamic state of the multi-module spacecraft. Based on the updated dynamic state, the module updates the state of the ground simulation sensor and transmits the updated state information to the application software module through the component interface simulation module. The state information includes operational information and combined configuration information. The component interface simulation module encapsulates multiple general component simulation docking functions, and the component interface simulation module is used to enable data communication between the application software module and the dynamics software module. The injection management module is communicatively connected to the application software module, the dynamics software module, and the component interface simulation module, respectively. The injection management module is used to control the general test platform to complete the simulation test according to the test command. The injection management module is provided with a dynamics injection data submodule, a general fault injection submodule, and an injection data and command submodule. The general fault injection submodule is used to perform fault simulation injection to verify whether the fault diagnosis and handling strategies of the on-board application software meet the software requirements. When performing fault simulation injection, the general fault injection submodule is specifically used to perform the following operations: The state information output by the dynamics software module is superimposed to simulate numerical faults; An interface layer and a protocol layer are added to the component interface simulation module to simulate interface failures and protocol failures; The operation monitoring module is used to monitor the telemetry data output by the application software module and the dynamics software module, and to switch the configuration of the multi-module spacecraft according to control commands, including: Based on the parameter data in the parameter binding table, the main configuration number and sub-configuration number that need to be switched are determined, and the main configuration number and the sub-configuration number are transmitted to the dynamics software module; wherein, the main configuration number is used to determine the spacecraft structure, and the sub-configuration number is used to determine the different on-orbit phase characteristics of a specific configuration; The dynamics software queries and modifies the corresponding injected interface variables based on the number information to enter the configuration switching program; The spacecraft's cabin characteristic information is updated by re-assigning the primary configuration number and the sub-configuration number through a re-initiation function call.

2. The platform of claim 1, wherein, The application software module includes an interface input / output logic submodule, a sensor acquisition and information processing submodule, an actuator instruction and information processing submodule, and an validity judgment and fault diagnosis submodule, wherein: The multi-module spacecraft application software module has a multi-layer architecture, with each layer being independently encapsulated and communicating with each other using standard interfaces; The interface input / output logic submodule, the sensor acquisition and information processing submodule, and the actuator instruction and information processing submodule are all set to the maximum interface envelope to make the internal and external interfaces universal, and the parameters and parameter calling timing of the corresponding submodules are configured according to the functional requirements of each compartment. The validity judgment and fault diagnosis submodule is used to judge the validity of communication status, self-test status, outlier and theoretical value range, and to perform fault diagnosis on the assembly of different configurations and different compartments according to the preset fault diagnosis principles. The faults include component-level faults and system-level faults. Component-level faults are judged based on the measurement data of multiple components or in combination with consistency checks. System-level faults are judged based on control performance.

3. The platform of claim 1, wherein, The dynamics software standard interface includes dynamics interface files and standardized software, wherein: The dynamics interface file contains an array of objects that cover the configuration requirements of all complex spacecraft, in order to standardize the control input interface and excitation output interface of the general test platform. The dynamics interface file includes a standardized input structure and a standardized output structure. The standardized software includes standard functions for dynamic initialization and standard functions for dynamic operation.

4. The platform of claim 3, wherein, include: The standardized input structure is defined according to the control instructions received by each actuator of the general test platform; The standardized output structure is defined based on the calculated state information of the multi-module spacecraft; The dynamic initialization standard function is determined based on the initialization scenario, the initialization state of the module, and the initial operating mode of the spacecraft; wherein, the initialization scenario is determined based on the operating mode of the test scenario, and the initialization state and the initial operating mode are determined based on the module information; The dynamic operation standard function is determined based on the operating parameters for data interaction in multi-module spacecraft.

5. The platform of claim 1, wherein, The integrated configuration parameter management submodule is used to convert the spacecraft's configuration parameters from a single-module mode to a multi-module test mode. Specifically, the integrated configuration parameter management submodule performs the following operations during the configuration parameter conversion: The combined state of the spacecraft is processed according to the preset configuration parameter conversion algorithm, and the configuration parameters of each module of the spacecraft are output. All configuration parameters are processed using multidimensional arrays to obtain a parameter binding table for storage and retrieval; The parameter binding table is indexed and accessed based on the application scenario number, current module number, current configuration number, and current mode subconfiguration number to conduct simulation tests on multi-module spacecraft.

6. The platform of claim 1, wherein, The general component simulation docking function is used to perform interface conversion between the dynamics software and the on-board software according to the on-board software segment markers. The general component simulation docking function includes the inertial measurement unit excitation function, the sun sensor excitation function, the star sensor excitation function, the infrared sensor excitation function, and the GPS excitation function, wherein: The excitation function of the inertial measurement unit is used to accumulate the angular rate and acceleration output by the dynamic inertial measurement unit, and after equivalent conversion, obtain the change in a single control cycle. The star sensor excitation function is used to convert, package, and frame the raw star sensor data output by the dynamics software according to the excitation requirements of the star sensor, so as to obtain the transmission data required for excitation. The infrared sensor excitation function is used to perform unit conversion on the chord width and angle of incidence data; The GPS excitation function is used to convert and package the UTC and GPS position / velocity data output by dynamics according to the GPS excitation requirements, so as to obtain the transmission data required for excitation and send it to the on-board interface device.

7. The platform of claim 1, wherein, The parameters and states of the dynamics software module, application software module, and component interface simulation module are set sequentially, and the multi-cabin, multi-mode flight simulation is switched. The general test platform is then controlled to conduct simulation tests according to the control commands.

8. The platform of claim 1, wherein, The operation monitoring module communicates with the dynamics software module and the application software module through the multi-compartment display module, and decodes and displays the telemetry data according to a preset protocol to monitor the data during the simulation test.

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