Simulation Verification Method for Low-Thrust Orbit Change Strategy of All-Electric Propulsion Satellite Platform

Abstract

This invention discloses a simulation verification method for a low-thrust orbit change strategy for an all-electric propulsion satellite platform, relating to the field of spacecraft orbit control and simulation technology. To address the shortcomings of existing technologies, which lack high-precision simulation and verification techniques for low-thrust orbit change strategies of all-electric propulsion satellites while comprehensively considering the effects of multiple perturbations, the technical solution provided by this invention includes: steps of acquiring a preset input model and inputting the satellite's orbital six elements, thruster parameters, and orbit change strategy; steps of determining the orbit change stage and generating the orbit change strategy based on the preset input model; steps of acquiring a preset orbital dynamics model, calculating the thrust vector according to the orbit change strategy, and updating the satellite's orbital parameters and mass consumption; and steps of acquiring a preset attitude control model, solving the current satellite's attitude quaternions, and determining whether the satellite's attitude is stable based on the solution results. This method is applicable to the design and verification of low-thrust, long-period orbit change strategies for all-electric propulsion satellite platforms.

Description

Technical Field

[0001] This involves the field of spacecraft orbit control and simulation technology, specifically the simulation verification of low-thrust orbit change strategies for all-electric propulsion satellite platforms. Background Art

[0002] All-electric propulsion technology has gained widespread attention in spacecraft propulsion systems in recent years, particularly in the field of geostationary orbit (GEO) satellites. Compared to traditional chemical propulsion systems, all-electric propulsion systems significantly improve propulsion efficiency through high specific impulse propulsion technologies such as ion electric propulsion or Hall electric propulsion, thereby reducing the total launch mass and development cost of satellites. The core advantage of this technology lies in its ability to significantly reduce propellant load, resulting in a significantly improved payload ratio. For example, Boeing Satellite Systems, Orbital Sciences Corporation, and Loral Space Systems in the United States have all successfully developed all-electric propulsion satellite platforms, which are widely used in communication and navigation satellites. The German company OHB has also conducted in-depth research in this field and launched a commercial satellite platform based on all-electric propulsion technology.

[0003] Compared to international advanced technologies, my country is also accelerating the application and development of all-electric propulsion technology. Related research focuses on ion electric propulsion systems, optimized design of Hall thrusters, and efficient orbit transfer strategies, with significant progress already achieved. However, satellite platforms based on all-electric propulsion technology typically suffer from low thrust, long orbit change times, and a large number of orbits, making their orbit change strategies more complex than those of chemical propulsion systems. In practical applications, satellites undergo long-period orbit change processes to move from geostationary transfer orbit (GTO) to GEO orbit, while being affected by various factors such as atmospheric drag, Earth's non-spherical gravitational field, solar radiation pressure, and gravitational perturbations from the Moon and the Sun. Therefore, the precise design and verification of orbit change strategies has become a key technical challenge.

[0004] Existing research has accumulated some methods for orbit change strategies of finite thrust satellites, such as classical orbital dynamics modeling and thruster design, but these mainly focus on short-period chemical propulsion orbit change processes. For all-electric propulsion satellite platforms, existing methods cannot fully adapt to their complex orbit change processes involving long periods, multiple stages, and multiple perturbations. For example:

[0005] Strategy design limitations: Existing methods are difficult to systematically optimize the directional control and magnitude distribution of thrust vectors in multiple orbit change stages, resulting in low orbit change efficiency.

[0006] Insufficient simulation accuracy: There is still considerable room for improvement in high-precision long-period simulation technology, especially in the comprehensive modeling and processing of multiple perturbation factors.

[0007] Attitude control coupling problem: There is no perfect solution for the attitude control coupling between the thruster's operating state and orbital dynamics, which may affect the stability of orbit change.

[0008] The aforementioned problems not only limit the further application of all-electric propulsion technology, but also increase the complexity and uncertainty of orbit transfer.

[0009] Therefore, how to achieve high-precision simulation and verification of low-thrust orbit change strategy for all-electric propulsion satellites under the premise of comprehensively considering the influence of multiple perturbations has become a technical problem that urgently needs to be solved in this field. Summary of the Invention

[0010] To address the shortcomings of existing technologies, which lack high-precision simulation and verification techniques for low-thrust orbit-changing strategies of all-electric propulsion satellites while comprehensively considering the effects of multiple perturbations, the technical solution provided by this invention is as follows:

[0011] A simulation method for a low-thrust orbital maneuvering strategy on an all-electric propulsion satellite platform includes:

[0012] The steps for collecting the preset input model and inputting the satellite's orbital root numbers, thruster parameters, and orbit change strategy are as follows:

[0013] Based on the preset input model, the steps of determining the trajectory change stage and generating the trajectory change strategy are as follows:

[0014] The steps include: acquiring a preset orbital dynamics model, calculating the thrust vector based on the orbit change strategy, and updating the satellite's orbital parameters and mass consumption;

[0015] The steps include: collecting a preset attitude control model, calculating the attitude quaternions of the current satellite, and determining whether the satellite attitude is stable based on the calculation results.

[0016] Furthermore, a preferred embodiment is provided in which the orbital dynamics model updates the satellite's orbital parameters by combining the thrust vector, the satellite's current position, and velocity through an orbital recursion model.

[0017] Furthermore, a preferred embodiment is provided in which the orbital dynamics model calculates the propellant mass consumption based on the satellite's thrust and specific impulse.

[0018] Furthermore, a preferred embodiment is provided in which the attitude control model calculates the attitude quaternions in the orbital coordinate system and converts them into attitude quaternions in the inertial coordinate system.

[0019] Based on the same inventive concept, this invention also provides a simulation verification device for a low-thrust orbit-changing strategy for an all-electric propulsion satellite platform, comprising:

[0020] The module collects the preset input model and inputs the satellite's orbital root numbers, thruster parameters, and orbit change strategy;

[0021] Based on the preset input model, a module is used to determine the trajectory change stage and generate a trajectory change strategy;

[0022] A module that acquires a preset orbital dynamics model, calculates the thrust vector based on the orbit change strategy, and updates the satellite's orbital parameters and mass consumption;

[0023] This module collects a preset attitude control model, calculates the attitude quaternions of the current satellite, and determines whether the satellite's attitude is stable based on the calculation results.

[0024] Based on the same inventive concept, this invention also provides a simulation verification method for a low-thrust orbit-changing strategy on an all-electric propulsion satellite platform, comprising:

[0025] The steps include outputting the simulation results of the described simulation method and verifying the simulation accuracy through multi-perturbation factor modeling.

[0026] Based on the same inventive concept, this invention also provides a simulation verification device for a low-thrust orbit-changing strategy for an all-electric propulsion satellite platform, comprising:

[0027] This module outputs the simulation results of the described simulation method and verifies the simulation accuracy through multi-perturbation factor modeling.

[0028] Based on the same inventive concept, the present invention also provides a computer storage medium for storing a computing program, wherein when the computer program is read by a computer, the computer executes the method described thereon.

[0029] Based on the same inventive concept, the present invention also provides a computer, including a processor and a storage medium, wherein when the processor reads a computer program stored in the storage medium, the computer executes the method described thereon.

[0030] Based on the same inventive concept, the present invention also provides a computer program product, which, when executed, implements the method described.

[0031] Compared with the prior art, the advantages of the technical solution provided by the present invention are as follows:

[0032] The low-thrust orbit change input model for an all-electric propulsion satellite platform enables comprehensive input of orbital parameters, thruster parameters, and orbit change strategies, allowing the simulation process to accurately reflect actual mission requirements. In existing methods, the input model often fails to refine down to the thruster's on / off states and operating parameters, while this model provides more precise strategy input, significantly improving the reliability and adaptability of the simulation model.

[0033] The low-thrust orbit change strategy interpretation model can automatically determine the current orbit change stage and send corresponding commands based on the number of orbital elements and the propulsion stage determination conditions. Compared with existing fixed strategy models, this dynamic interpretation model optimizes orbit change efficiency by adjusting the thrust direction and magnitude in real time, showing significant advantages, especially in long-period, multi-stage orbit changes.

[0034] The orbital dynamics model employs a high-precision thrust vector calculation method, incorporating orbital perturbation factors into the model to achieve accurate calculations of thrust direction, orbital parameters, and mass changes. Compared to existing studies that lack sufficient modeling of multiple perturbation factors, this significantly improves the accuracy of long-term orbit change simulation results, with a simulation error better than 1‰, making it more valuable for practical satellite orbit change missions.

[0035] The attitude control model effectively avoids orbital errors caused by attitude instability by accurately calculating the stable attitude quaternions of the satellite platform under thrust ignition and shutdown conditions. In existing methods, attitude control and thrust calculation are usually performed independently, while this method achieves coupled optimization of the two, ensuring a high degree of consistency between thrust direction and attitude control.

[0036] The strategy interpretation model, orbital dynamics model, and attitude control model are integrated into a complete simulation framework and embedded in the SpaceSim simulation software. Through a visualization module, the orbit change process is not only presented intuitively, but also high-precision orbital six-factor comparison data are output, showing a high degree of consistency with real satellite telemetry data. Compared to fragmented, single simulation tools, this integrated design improves user convenience and system efficiency while reducing mission planning complexity.

[0037] During the simulation, the thruster status was determined by combining historical thruster on / off states with real-time sun exposure conditions, thus optimizing the thruster's energy utilization efficiency. Existing research typically has low time resolution for thruster status control, but this method achieves more refined energy management, thereby further improving the economy and feasibility of the trajectory change process.

[0038] It is applicable to the design and verification of low-thrust, long-period orbit change strategies for all-electric propulsion satellite platforms. Attached Figure Description

[0039] Figure 1 A flowchart illustrating a simulation verification method for a low-thrust orbit-changing strategy on an all-electric propulsion satellite platform;

[0040] Figure 2 The results of the comparison of the six elements of the orbit during the orbit change process are as follows: (a) comparison of the semi-major axis, (b) comparison of the eccentricity, (c) comparison of the orbit inclination, (d) comparison of the right ascension of the ascending node, (e) comparison of the argument of perihelion, and (f) comparison of the mean perihelion.

[0041] Figure 3 A schematic diagram for setting satellite platform orbital parameters;

[0042] Figure 4 A schematic diagram of thruster parameter settings;

[0043] Figure 5 A schematic diagram illustrating the trajectory change strategy;

[0044] Figure 6 for Figure 1 A schematic diagram of the implementation in SpaceSim software;

[0045] Figure 7 This is a visual representation of the simulation process;

[0046] Figure 8 This is a schematic diagram of the simulation results. Detailed Implementation

[0047] To make the advantages and benefits of the technical solution provided by the present invention clearer, the technical solution provided by the present invention will now be described in further detail with reference to the accompanying drawings, specifically:

[0048] Implementation Method 1: This implementation method provides a simulation method for a low-thrust orbital maneuvering strategy on an all-electric propulsion satellite platform, including:

[0049] The steps for collecting the preset input model and inputting the satellite's orbital root numbers, thruster parameters, and orbit change strategy are as follows:

[0050] Based on the preset input model, the steps of determining the trajectory change stage and generating the trajectory change strategy are as follows:

[0051] The steps include: acquiring a preset orbital dynamics model, calculating the thrust vector based on the orbit change strategy, and updating the satellite's orbital parameters and mass consumption;

[0052] The steps include: collecting a preset attitude control model, calculating the attitude quaternions of the current satellite, and determining whether the satellite attitude is stable based on the calculation results.

[0053] Implementation Method 2: This implementation method further defines the simulation method for a low-thrust orbit change strategy of an all-electric propulsion satellite platform provided in Implementation Method 1. The orbital dynamics model updates the satellite's orbital parameters by combining the thrust vector, the satellite's current position, and velocity with an orbital recursion model.

[0054] Implementation Method 3: This implementation method further defines the simulation method for a low-thrust orbit change strategy of an all-electric propulsion satellite platform provided in Implementation Method 1. The orbital dynamics model calculates the propellant mass consumption based on the satellite's thrust and specific impulse.

[0055] Implementation Method 4: This implementation method further defines the simulation method for a low-thrust orbit change strategy of an all-electric propulsion satellite platform provided in Implementation Method 1. The attitude control model calculates the attitude quaternions in the orbital coordinate system and converts them into attitude quaternions in the inertial coordinate system.

[0056] Implementation Method 5: This implementation method provides a simulation verification device for a low-thrust orbit change strategy on an all-electric propulsion satellite platform, comprising:

[0057] The module collects the preset input model and inputs the satellite's orbital root numbers, thruster parameters, and orbit change strategy;

[0058] Based on the preset input model, a module is used to determine the trajectory change stage and generate a trajectory change strategy;

[0059] A module that acquires a preset orbital dynamics model, calculates the thrust vector based on the orbit change strategy, and updates the satellite's orbital parameters and mass consumption;

[0060] This module collects a preset attitude control model, calculates the attitude quaternions of the current satellite, and determines whether the satellite's attitude is stable based on the calculation results.

[0061] Implementation Method Six: This implementation method provides a simulation verification method for a low-thrust orbit-changing strategy on an all-electric propulsion satellite platform, including:

[0062] The steps include outputting the simulation results of the simulation method provided in Implementation Method 1, and verifying the simulation accuracy through multi-perturbation factor modeling.

[0063] Specifically, including:

[0064] Step 1: Construction of the input model for low-thrust orbit change on an all-electric propulsion satellite platform

[0065] By inputting the model to set the satellite platform's orbital parameters, thruster parameters, and orbit change strategy, complete initial conditions are provided for simulation verification.

[0066] Detailed Description: Based on the visualization-based human-computer interface of the simulation software, users input the six orbital elements of the all-electric propulsion satellite platform (including semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, argument of perigee, and mean perigee), the total mass of the satellite platform, and the dynamic model parameters. Simultaneously, key parameters such as the number of thrusters, thrust, and specific impulse are set, and the propulsion phase judgment conditions and participating thrusters in the orbit change strategy are clarified. In this step, all input data is verified by software to ensure its completeness and rationality. After completing this step, a complete input model file is generated, serving as the initial data source for subsequent simulation processes.

[0067] Step 2: Implementation of the interpretation model for low-thrust trajectory change strategy

[0068] Based on the input model and the six orbital elements, the current orbital change stage is interpreted in real time, and corresponding task execution instructions are generated.

[0069] Detailed Description: The system analyzes the six fundamental numbers of the satellite's orbit using an interpretation model to determine the current orbital maneuver phase, including the in-orbit propulsion phase, inclination adjustment phase, and eccentricity adjustment phase. Based on the phase characteristics, mission execution commands are generated and pushed to the orbital dynamics model and attitude control model. Simultaneously, it continuously monitors whether the satellite is in the Earth's shadow region and the thruster's on / off status to ensure thruster energy utilization efficiency. In the second and third phases of the orbital maneuver, the system calculates the latitude argument to determine whether a thrust direction change is necessary and pushes the results to the orbital dynamics model. The output of this step is the current mission command, thruster status, and thrust direction adjustment result, which serve as input to the orbital dynamics model.

[0070] Step 3: Construction and Operation of the Orbital Dynamics Model

[0071] Based on mission instructions and thruster status, thrust vector calculation, trajectory recursion, and mass consumption calculation are performed.

[0072] Detailed Description: The system receives task commands and thruster on / off status from the interpretation model and calculates the thrust vector according to different orbital maneuvering phases. During the orbital propulsion phase, the thrust vector is along the velocity direction; during the inclination adjustment phase, the thrust vector forms a fixed angle with the orbital plane, while also including a component along the velocity direction; during the eccentricity adjustment phase, the thrust vector is perpendicular to the eccentricity vector and forms a fixed angle with the orbital plane. Using a high-precision orbital recursive model (HPOP), combined with the thrust vector, the satellite's current position, and velocity, the system calculates the satellite's orbital parameters and mass loss at the next moment and updates them to the simulation system for use in subsequent steps.

[0073] Step 4: Establishment and Application of Attitude Control Model

[0074] Based on mission commands and thrust vectors, the satellite's attitude quaternions are calculated to ensure the accuracy of the thrust direction.

[0075] Detailed Description: This model receives attitude control commands from the interpretation model, calculates the satellite's attitude quaternions in the orbital coordinate system according to different attitude control requirements, and further converts them to attitude quaternions in the inertial coordinate system. During thrust engine firing, it ensures the satellite's attitude aligns with the thrust direction; during thrust engine shutdown, it maintains the satellite's attitude stability. Through this model, thrust engine firing and attitude adjustment form a closed-loop control, ensuring the satellite's dynamic balance during orbit changes.

[0076] Step 5: Output and Visualization of Simulation Results

[0077] The simulation results are integrated into the visualization module to enable dynamic display and storage of the data.

[0078] Detailed Description: Based on the outputs of the orbital dynamics model and attitude control model, simulation data is generated, including changes in the orbital six roots, thruster status, thrust direction, and satellite attitude information. Through a visualization module, users can dynamically observe the real-time changes in orbital parameters and satellite status during orbit changes and compare them with actual telemetry data for verification. The simulation data is stored in a specified format, supporting subsequent mission planning and data analysis.

[0079] Step Six: Full Process Verification and Optimization

[0080] The entire process was verified based on simulation data, and the trajectory change strategy was adjusted to optimize the results.

[0081] Detailed Description: This study analyzes the changes in the six orbital elements in the simulation results and compares the errors with actual telemetry data, focusing on evaluating the accuracy of thrust vector calculation and the stability of attitude control. For stages with large errors, the orbital change efficiency and simulation accuracy are optimized by adjusting the mission command generation logic and thrust direction control strategy in the interpretation model. The optimized scheme is verified to be repeatable until the expected results are achieved.

[0082] Implementation Method Seven: This implementation method provides a simulation verification device for a low-thrust orbit-changing strategy on an all-electric propulsion satellite platform, comprising:

[0083] This module outputs the simulation results of the simulation method provided in Implementation Method 1 and verifies the simulation accuracy through multi-perturbation factor modeling.

[0084] Implementation Method 8: This implementation method provides a computer storage medium for storing a computing program. When the computer program is read by the computer, the computer executes the method provided in Implementation Method 1.

[0085] Implementation Method Nine: This implementation method provides a computer, including a processor and a storage medium. When the processor reads a computer program stored in the storage medium, the computer executes the method provided in Implementation Method One.

[0086] Implementation Method 10: This implementation method provides a computer program product. As a computer program, when the computer program is executed, it implements the method provided in Implementation Method 1.

[0087] Implementation Method Eleven: Combination Figure 1-6 This embodiment describes the technical solution provided above in further detail through specific examples. Specifically:

[0088] This implementation method includes a low-thrust orbit change input model for an all-electric propulsion satellite platform, a strategy interpretation model, an orbital dynamics model, and an attitude control model. The main implementation process is as follows:

[0089] 1. The input model for low-thrust orbit change of the all-electric propulsion satellite platform is responsible for the input of platform parameters and strategies, including:

[0090] 1) Input the number of orbital support points for the all-electric propulsion satellite platform, the total mass of the platform, and the parameters of the platform's dynamic model;

[0091] 2) Input the thruster parameters of the all-electric propulsion satellite platform, including the number of thrusters, the thrust and specific impulse of each thruster;

[0092] 3) Input the low-thrust orbit change strategy for the all-electric propulsion satellite platform, including the propulsion phase judgment conditions and the participation of the thruster.

[0093] 2. The all-electric propulsion satellite platform's low-thrust orbit change strategy interpretation model is responsible for strategy interpretation and command delivery, including:

[0094] 1) Based on the six orbital elements of the all-electric propulsion satellite platform and the propulsion stage determination conditions, the current orbit change stage of the all-electric propulsion satellite is determined, which is divided into three stages: the first stage is propulsion within the orbital plane, raising the perigee; the second stage is propulsion at a certain angle to the orbital plane, raising the orbital altitude while reducing the orbital inclination; the third stage is thrust within the orbital plane perpendicular to the eccentricity vector, while at a certain angle to the orbital plane, reducing the eccentricity and reducing the orbital inclination.

[0095] 2) Based on different mission stages, construct mission execution instructions and push them to the orbital dynamics model and attitude control model.

[0096] 3) Determine the extent of sunlight exposure on the satellite platform and whether it is in the Earth's shadow.

[0097] 4) Based on the platform thruster’s historical on / off status and current exposure to sunlight, determine the thruster’s on / off status at the current moment and push the thruster’s on / off status to the orbital dynamics model.

[0098] 5) During the second and third stages of orbit change, determine whether the thrust needs to change direction, provided that the amplitude angle is 90° or 270°, and push the judgment result to the orbit dynamics model.

[0099] 3. The low-thrust orbital dynamics model for the all-electric propulsion satellite platform is responsible for thrust calculation and orbital recursion, including:

[0100] 1) Receive the task execution instructions, thruster on / off status and thrust direction conversion results from the interpretation model.

[0101] 2) Based on the input parameters and strategy, calculate the thrust vector in the J2000 inertial frame (hereinafter referred to as the inertial frame). The calculation process varies depending on the stage, and is specifically divided into:

[0102] A. First stage: Thrust vector is along the velocity direction.

[0103] The velocity vector of the satellite inertial system is obtained by solving the six roots of the satellite platform orbit. After normalization, the thrust direction is obtained. The thrust vector is obtained by determining the magnitude of the thruster thrust.

[0104] B. Second stage: The thrust vector forms a certain angle α with the track surface and moves along the velocity direction within the track surface.

[0105] Based on the six orbital roots of the satellite platform, the velocity vector in the satellite's inertial frame is calculated and transformed into the VVLH orbital coordinate system (referred to as the orbital frame). The flight path angle β is then calculated. Based on the flight path angle and the thrust vector angle, the thrust vector in the orbital frame can be calculated.

[0106]

[0107] In the formula, F represents the magnitude of the thrust of the thruster. After switching to an inertial frame, the thrust vector in the inertial frame can be obtained.

[0108] C. Third stage: The thrust vector forms a certain angle α with the track surface and is perpendicular to the eccentricity vector within the track surface.

[0109] The unit vector of eccentricity in the satellite's inertial frame is calculated based on the six fundamental parameters of the satellite platform's orbit. Then, it is transformed to the orbital coordinate system, thereby obtaining the unit thrust component in the orbital plane.

[0110]

[0111] In the formula, y is the unit vector of the orbital system.

[0112] Therefore, the thrust vector in the orbital system can be obtained:

[0113]

[0114] In the formula, F represents the magnitude of the thrust of the thruster. Let F be the in-plane thrust component of the orbital system, and it should satisfy F. orbit_y =0. (This likely refers to a specific value or quantity.) After switching to an inertial frame, the thrust vector in the inertial frame can be obtained.

[0115] 3) Calculate the mass loss Δm based on the magnitude of the thrust:

[0116]

[0117] 4) Based on the thrust vector in the inertial frame, the current position and velocity of the satellite platform, and the mass loss, the HPOP high-precision orbital recursive model is used to update the position, velocity, and mass of the satellite platform.

[0118] 4. The low-thrust attitude control model for the all-electric propulsion satellite platform is responsible for attitude control during the orbit change process, solving for the stable attitude quaternions of the satellite platform under ignition or shutdown conditions. This mainly includes:

[0119] 1) Receive satellite attitude control commands from the interpretation model.

[0120] 2) Based on different attitude control commands, calculate the attitude quaternions of the satellite platform in the orbital coordinate system.

[0121] 3) Based on the satellite platform attitude quaternion in the orbital frame, solve for the satellite platform attitude quaternion in the inertial frame.

[0122] Based on the above process, the simulation process of the low-thrust orbit change strategy for an all-electric propulsion satellite platform is as follows: Figure 1 As shown.

[0123] In practical implementation, this implementation method can achieve high-precision simulation of the low-thrust orbit change process of an all-electric propulsion platform by inputting a specified orbit change strategy, which involves long cycles, multiple perturbations, and complex strategies. Figure 2 This is a comparison of the orbit root numbers between simulation data and telemetry data from a real satellite platform during a low-thrust orbit change on an all-electric propulsion satellite platform under the same orbit change strategy.

[0124] As can be seen from the figure, the low-thrust orbit change process of the all-electric propulsion satellite platform implemented by this embodiment has extremely high accuracy. Compared with the actual telemetry data, the simulation results are better than 1‰ in accuracy during the orbit change process over one day. It can realize the low-thrust orbit change process of the all-electric propulsion satellite platform with high precision.

[0125] according to Figure 3-8 In the software simulation work, the SpaceSim spacecraft system simulation software independently developed by Harbin Institute of Technology was used for implementation. The implementation method is as follows:

[0126] Step 1: Based on the visual human-computer interaction interface of SpaceSim software, implement the low-thrust orbit change input model of the satellite platform. Set the satellite platform orbit parameters, thruster parameters, and orbit change strategy through the visual human-computer interaction interface, such as... Figure 3-5 As shown.

[0127] Step Two: According to Figure 1The process shown embeds the strategy interpretation model, orbital dynamics model, and attitude control model from the simulation method of low-thrust orbit change strategy for all-electric propulsion satellite platforms into the SpaceSim software, as follows: Figure 6 As shown.

[0128] Step 3: Connect the simulation results to the software's visualization and data storage modules to achieve visualized display and output of the simulation results, such as... Figure 7 and 8 As shown.

[0129] The above description of several specific embodiments further details the technical solution provided by the present invention in order to highlight the advantages and benefits of the technical solution provided by the present invention. However, the above-described specific embodiments are not intended to limit the present invention. Any reasonable modifications and improvements to the present invention, combinations of embodiments, and equivalent substitutions based on the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A simulation method for a low-thrust orbital maneuvering strategy on an all-electric propulsion satellite platform, characterized in that, include: The steps for collecting the preset input model and inputting the satellite's orbital root numbers, thruster parameters, and orbit change strategy are as follows: Based on the preset input model, the steps of determining the trajectory change stage and generating the trajectory change strategy are as follows: The steps include: acquiring a preset orbital dynamics model, calculating the thrust vector based on the orbit change strategy, and updating the satellite's orbital parameters and mass consumption; The steps include: collecting a preset attitude control model, calculating the attitude quaternion of the current satellite, and determining whether the satellite attitude is stable based on the calculation results; in, The system uses an interpretation model to analyze the six fundamental numbers of the satellite's orbit, determining the current orbital maneuvering phase, including the in-orbit propulsion phase, inclination adjustment phase, and eccentricity adjustment phase. Based on the phase characteristics, it generates mission execution commands and pushes them to the orbital dynamics model and attitude control model. It also determines in real time whether the satellite is in the Earth's shadow region and the on / off status of the thrusters. During the inclination and eccentricity adjustment phases of the orbital maneuver, it calculates the latitude argument to determine whether the thrust direction needs to be changed and pushes the thrust results to the orbital dynamics model. The system receives task commands and thruster on / off status pushed by the interpretation model, and calculates the thrust vector according to different orbit change stages. During the orbital propulsion stage, the thrust vector is along the velocity direction. During the inclination adjustment stage, the thrust vector is at a fixed angle with the orbital surface, and also includes a component along the velocity direction. During the eccentricity adjustment stage, the thrust vector is perpendicular to the eccentricity vector and at a fixed angle with the orbital surface. Using a high-precision orbital recursive model, combined with the thrust vector, the satellite's current position and velocity, the system calculates the satellite's orbital parameters and mass consumption at the next moment, and updates them to the simulation system. It receives attitude control commands from the interpretation model, calculates the satellite's attitude quaternions in the orbital coordinate system according to different attitude control requirements, and further converts them into attitude quaternions in the inertial coordinate system; when the thrusters are firing, it ensures that the satellite's attitude is consistent with the thrust direction; when the thrusters are off, it maintains the stability of the satellite's attitude.

2. The simulation method for a low-thrust orbit change strategy of an all-electric propulsion satellite platform according to claim 1, characterized in that, The orbital dynamics model updates the satellite's orbital parameters by combining the thrust vector, the satellite's current position, and velocity with an orbital recursion model.

3. The simulation method for a low-thrust orbit change strategy of an all-electric propulsion satellite platform according to claim 1, characterized in that, The orbital dynamics model calculates propellant mass consumption based on the satellite's thrust and specific impulse.

4. The simulation method for a low-thrust orbit change strategy of an all-electric propulsion satellite platform according to claim 1, characterized in that, The attitude control model calculates the attitude quaternions in the orbital coordinate system and converts them into attitude quaternions in the inertial coordinate system.

5. A simulation verification device for a low-thrust orbit change strategy on an all-electric propulsion satellite platform, characterized in that, include: The module collects the preset input model and inputs the satellite's orbital root numbers, thruster parameters, and orbit change strategy; Based on the preset input model, a module is used to determine the trajectory change stage and generate a trajectory change strategy; A module that acquires a preset orbital dynamics model, calculates the thrust vector based on the orbit change strategy, and updates the satellite's orbital parameters and mass consumption; A module that collects a preset attitude control model, calculates the attitude quaternion of the current satellite, and determines whether the satellite attitude is stable based on the calculation results; in, The system uses an interpretation model to analyze the six fundamental numbers of the satellite's orbit, determining the current orbital maneuvering phase, including the in-orbit propulsion phase, inclination adjustment phase, and eccentricity adjustment phase. Based on the phase characteristics, it generates mission execution commands and pushes them to the orbital dynamics model and attitude control model. It also determines in real time whether the satellite is in the Earth's shadow region and the on / off status of the thrusters. During the inclination and eccentricity adjustment phases of the orbital maneuver, it calculates the latitude argument to determine whether the thrust direction needs to be changed and pushes the thrust results to the orbital dynamics model. The system receives task commands and thruster on / off status pushed by the interpretation model, and calculates the thrust vector according to different orbit change stages. During the orbital propulsion stage, the thrust vector is along the velocity direction. During the inclination adjustment stage, the thrust vector is at a fixed angle with the orbital surface, and also includes a component along the velocity direction. During the eccentricity adjustment stage, the thrust vector is perpendicular to the eccentricity vector and at a fixed angle with the orbital surface. Using a high-precision orbital recursive model, combined with the thrust vector, the satellite's current position and velocity, the system calculates the satellite's orbital parameters and mass consumption at the next moment, and updates them to the simulation system. It receives attitude control commands from the interpretation model, calculates the satellite's attitude quaternions in the orbital coordinate system according to different attitude control requirements, and further converts them into attitude quaternions in the inertial coordinate system; when the thrusters are firing, it ensures that the satellite's attitude is consistent with the thrust direction; when the thrusters are off, it maintains the stability of the satellite's attitude.

6. A simulation verification method for a low-thrust orbit-changing strategy on an all-electric propulsion satellite platform, characterized in that, include: The steps include outputting the simulation results of the simulation method described in claim 1 and verifying the simulation accuracy through multi-perturbation factor modeling.

7. A simulation verification device for a low-thrust orbit change strategy on an all-electric propulsion satellite platform, characterized in that, include: The module outputs the simulation results of the simulation method described in claim 1 and verifies the simulation accuracy through multi-perturbation factor modeling.

8. A computer storage medium for storing computing programs, characterized in that, When the computer program is read by the computer, the computer executes the method of claim 1.

9. A computer, comprising a processor and a storage medium, characterized in that, When the processor reads the computer program stored in the storage medium, the computer executes the method of claim 1.

10. A computer program product, as a computer program, is characterized by: When the computer program is executed, it implements the method of claim 1.

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