Earth-moon transfer orbit flight strategy closed-loop verification method and system

Abstract

The invention provides an earth-moon transfer orbit flight strategy closed-loop verification method and system. The method comprises the steps that satellites are connected to form a closed-loop verification system; powering up and setting the working state of each subsystem; powering off and pressing the travel switch; connecting spatial dynamics simulation equipment and setting the spatial dynamics simulation equipment to be in an initial open-loop state; performing satellite-rocket separation power-on program control; interpreting an initial orbit injection working condition until the flywheel is autonomously switched into a flywheel rough sun alignment working condition; verifying to enter a precise daily working condition; performing earth-moon cruise flight phase verification, moon orbit brake capture phase verification and moon orbit double-satellite separation phase verification; and replacing flight parameters and selecting main and standby single machines to carry out performance research verification. According to the method, deduction verification can be carried out aiming at flight missions with more constraints, restrictions and high complexity, the correctness of control strategies in all stages is assessed, the work collaboration of all subsystems and the rationality of flight program setting are assessed, and reliable operation of the moonlet after launching and orbit injection and smooth execution of task processes in all flight stages are guaranteed.

Description

Technical Field

[0001] The present invention belongs to the field of spacecraft testing technology, and in particular, relates to a closed-loop verification method and system for an Earth-Moon transfer orbit flight strategy, and in particular to a closed-loop verification method for an Earth-Moon transfer orbit flight strategy of a lunar exploration small satellite. Background Art

[0002] In the field of lunar exploration, in addition to the Chang'e-1 to Chang'e-5 lunar probes, and the Queqiao-1 and Queqiao-2 lunar relay satellites, small satellites such as Tiandu-1 have also been launched. Compared with the Chang'e series probes and Queqiao relay satellites, small lunar exploration satellites weighing less than 100 kilograms have more constraints and restrictions in performing lunar exploration missions due to the shortage of resources in terms of single-machine configuration, fuel weight, and measurement and control support. However, in the process of transferring orbits from the Earth to the Moon, the flight is highly complex and any detail may affect the success or failure of the entire mission.

[0003] After searching the prior art, the patent document "A method for rapid joint simulation and verification of the entire system of a microsatellite overall plan" (CN104865841A) discloses a method for parallel simulation of various subsystems of a microsatellite design plan using a general simulator architecture. The simulation goal is to verify the satellite design rather than the flight strategy. At the same time, as a simulation method, it does not test the functions and performance of the actual satellite product.

[0004] The verification method disclosed in the patent document "A closed-loop simulation verification system and method for agile satellite mission interpretation" (CN105334756A) effectively solves the key links of verification simulation, modification inversion, etc. of agile satellite mission planning, and can ensure the correct implementation of agile satellite mission planning. However, it is also only a simulation rather than a test, and only uses virtual means, and cannot constitute a closed-loop verification system.

[0005] The patent document "A Semi-Physical Simulation Verification System and Method for Heterogeneous Interconnected Satellite Broadband Communication" (CN106211220A) adopts the idea of ​​combining software and hardware for system design and construction, which has the characteristics of high flexibility and strong scalability, but only verifies the satellite broadband communication subsystem and does not involve simulation and testing of other satellite subsystems. It is necessary to verify the information interaction and coordination of work between the subsystems. Verifying only the communication subsystem cannot meet the verification requirements of the flight strategy.

[0006] In addition, the verification methods disclosed in the above patent documents are only applicable to low-Earth satellites, lack the ability to verify key lunar exploration strategies such as weak-light power-on, trip switch operation, and separation of binary satellite assemblies, and do not have the function of verifying the flight strategies of small lunar exploration satellites.

[0007] At present, no other similar technical descriptions or reports have been found for the verification of the Earth-Moon transfer orbit flight strategy for small satellites, and no other similar information has been collected at home and abroad. Therefore, it is necessary to conduct a comprehensive evaluation and verification of the various states, modes, and polarities of the lunar exploration small satellites in the development stage to meet the verification needs of the lunar exploration small satellites or small satellite assemblies, so as to ensure the reliable operation of the small satellites after they are launched into orbit, and the smooth execution of the mission processes in each flight phase. Summary of the invention

[0008] In view of the defects in the prior art, the purpose of the present invention is to provide a closed-loop verification method and system for Earth-Moon transfer orbit flight strategy.

[0009] A closed-loop verification method for an Earth-Moon transfer orbit flight strategy provided by the present invention comprises:

[0010] Step S1: The satellite is connected to the ground test system to form a closed-loop verification system;

[0011] Step S2: Power on the satellite and set the working state of each subsystem to the weak light power-on quasi state before the satellite-rocket separation;

[0012] Step S3: The satellite is connected to a space dynamics simulation device, and the initial value of the simulation is set to an initial open-loop state;

[0013] Step S4: performing power-on program control for satellite-rocket separation;

[0014] Step S5: Determine the initial orbit insertion condition until the satellite autonomously switches to the flywheel coarse alignment condition;

[0015] Step S6: Enable the star sensor data to refresh the gyro data function to verify that the satellite has entered the precise sun-pointing condition;

[0016] Step S7: Conducting Earth-Moon cruise flight phase verification;

[0017] Step S8: Perform lunar orbit braking capture phase verification;

[0018] Step S9: Perform lunar orbit binary satellite separation phase verification;

[0019] Step S10: Change the flight parameters and select the primary and standby machines to conduct performance verification.

[0020] Preferably, the step S4 includes:

[0021] Step S4.1: Set the solar simulation array weak light curve to judge the execution effect of the satellite weak light power-on program control process.

[0022] and\or

[0023] Step S4.2: Manually release the travel switch, inject the initial orbit parameter package, and simultaneously set the space dynamics simulation device to a closed-loop state.

[0024] Preferably, in step S1, a single-star or dual-star connection method is set according to specific task requirements;

[0025] When conducting dual-satellite combination flight mission verification, both satellites are placed on their respective parking racks, the inter-satellite cables are not connected, and the satellites are connected to their respective ground test systems;

[0026] Connect the satellite measurement link, control software link, execution output link and ground simulation link in series to form a ground simulation flight closed-loop verification system;

[0027] The inertia, center of mass and mass of another satellite after assembly are simulated equivalently through space dynamics simulation equipment.

[0028] In step S2, the first power-on setting is performed;

[0029] Through the satellite MRAM, the satellite-rocket separation self-start enable, satellite-rocket separation program control, low-light power-on and other flags are set to determine whether the corresponding telemetry data meets the requirements. If and only if the requirements are met, the satellite is powered off.

[0030] In step S3, the space dynamics simulation equipment is operated to provide external electrical excitation for the satellite gyroscope and star sensor, and feedback signals from the flywheel and thruster components are collected, and a satellite on-orbit closed-loop flight simulation environment is constructed in combination with the dynamics model.

[0031] Preferably, in step S7, the satellite and the space dynamics simulation device jump time, record the cruise flight orbit parameters, enter the Earth-Moon cruise flight phase, and interpret the telemetry parameters to verify the correctness of the satellite system polarity.

[0032] In step S8, the satellite and the space dynamics simulation device jump time, record the braking and capturing orbit parameters, enter the lunar orbit braking and capturing phase, and interpret the telemetry parameters to verify the correctness of the satellite system polarity.

[0033] In the step S9, a two-star separation unlocking pyrotechnic device instruction is received through the pyrotechnic device signal detection device;

[0034] The space dynamics simulation equipment is used to set the changes in inertia, center of mass and mass of the binary satellite assembly before and after separation, and the binary satellite separation is carried out under the lunar orbit. The thruster jet is injected to verify the satellite attitude before and after separation and whether the disturbance can be offset in the closed-loop system.

[0035] If the satellite attitude and orbit control program indicates that the telemetry always maintains or restores the precise solar pointing mode, it is determined that the disturbance can be offset in the closed-loop system. If the satellite attitude and orbit control program indicates that the telemetry cannot maintain or restore the precise solar pointing mode, it is determined that the disturbance cannot be offset and this working condition is maintained for troubleshooting.

[0036] The telemetry parameters to be judged are the orbit type, satellite position and satellite speed parameters. If the parameters are consistent with the above-mentioned parameters, it is determined that the satellite control is correct. If the parameters are inconsistent with the above-mentioned parameters, it is determined that the satellite control is wrong, and this working condition is maintained for problem troubleshooting.

[0037] Preferably, in step S4.1, the dynamic illumination and changes of sunlight after the launch of the carrier rocket and before the separation of the satellite and the rocket are simulated by using the weak light curve of the solar simulation array, and the weak light power-on function of the satellite is triggered by the changes in voltage and current.

[0038] In the step S5, the signal is received by the pyrotechnic device signal detection device.

[0039] During the simulation, interpretation and verification of the satellite Earth-Moon transfer orbit flight process in steps S4 to S8, the entire Earth-Moon transfer flight process is accelerated by time hopping.

[0040] A closed-loop verification system for Earth-Moon transfer orbit flight strategy provided by the present invention comprises:

[0041] Module M1: The satellite is connected to the ground test system to form a closed-loop verification system;

[0042] Module M2: Power on the satellite and set the working state of each subsystem to the weak light power-on state before the separation of the satellite and rocket;

[0043] Module M3: Satellite connection space dynamics simulation equipment, the simulation initial value is set to the initial open loop state;

[0044] Module M4: Perform power-on program control for satellite-rocket separation;

[0045] Module M5: Determine the initial orbit insertion conditions until the satellite autonomously switches to the flywheel coarse alignment condition;

[0046] Module M6: Enable the star sensor data to refresh the gyro data function to verify that the satellite has entered the precise sun-pointing condition;

[0047] Module M7: Conduct verification of the Earth-Moon cruise flight phase;

[0048] Module M8: Conduct lunar orbit braking and capture phase verification;

[0049] Module M9: Conduct lunar orbit binary separation phase verification;

[0050] Module M10: Change flight parameters and select the primary and standby units to conduct performance verification.

[0051] Preferably, the module M4 includes:

[0052] Module M4.1: Set the solar simulation array weak light curve to judge the execution effect of the satellite weak light power-on program control process.

[0053] and\or

[0054] Module M4.2: Manually release the travel switch, inject the initial orbit parameter package, and simultaneously set the space dynamics simulation equipment to a closed-loop state.

[0055] Preferably, in the module M1, a single-star or dual-star connection method is set according to specific task requirements;

[0056] When conducting dual-satellite combination flight mission verification, both satellites are placed on their respective parking racks, the inter-satellite cables are not connected, and the satellites are connected to their respective ground test systems;

[0057] Connect the satellite measurement link, control software link, execution output link and ground simulation link in series to form a ground simulation flight closed-loop verification system;

[0058] The inertia, center of mass and mass of another satellite after assembly are simulated equivalently through space dynamics simulation equipment.

[0059] The module M2 performs the first power-on setting;

[0060] Through the satellite MRAM, the satellite-rocket separation self-start enable, satellite-rocket separation program control, low-light power-on and other flags are set to determine whether the corresponding telemetry data meets the requirements. If and only if the requirements are met, the satellite is powered off.

[0061] The module M3 operates and connects the space dynamics simulation equipment to provide external electrical excitation for the satellite gyroscope and star sensor, collects feedback signals from the flywheel and thruster components, and builds a satellite on-orbit closed-loop flight simulation environment in combination with the dynamics model.

[0062] Preferably, the satellite and space dynamics simulation equipment in the module M7 jumps time, notes the cruise flight orbit parameters, enters the Earth-Moon cruise flight phase, and interprets the telemetry parameters to verify the correctness of the satellite system polarity.

[0063] The satellite and space dynamics simulation equipment in the module M8 jumps time, injects braking and capturing orbit parameters, enters the lunar orbit braking and capturing phase, and interprets telemetry parameters to verify the correctness of the satellite system polarity.

[0064] The module M9 receives the instruction of two-star separation and unlocking of the pyrotechnic device through the pyrotechnic device signal detection device;

[0065] The space dynamics simulation equipment is used to set the changes in inertia, center of mass and mass of the binary satellite assembly before and after separation, and the binary satellite separation is carried out under the lunar orbit. The thruster jet is injected to verify the satellite attitude before and after separation and whether the disturbance can be offset in the closed-loop system.

[0066] If the satellite attitude and orbit control program indicates that the telemetry always maintains or restores the precise solar pointing mode, it is determined that the disturbance can be offset in the closed-loop system. If the satellite attitude and orbit control program indicates that the telemetry cannot maintain or restore the precise solar pointing mode, it is determined that the disturbance cannot be offset and this working condition is maintained for troubleshooting.

[0067] The telemetry parameters to be judged are the orbit type, satellite position and satellite speed parameters. If the parameters are consistent with the above-mentioned parameters, it is determined that the satellite control is correct. If the parameters are inconsistent with the above-mentioned parameters, it is determined that the satellite control is wrong, and this working condition is maintained for problem troubleshooting.

[0068] Preferably, in the module M4.1, the dynamic illumination and changes of sunlight after the launch of the carrier rocket and before the separation of the satellite and the rocket are simulated by using the weak light curve of the solar simulation array, and the weak light power-on function of the satellite is triggered by the changes in voltage and current.

[0069] The module M5 receives the signal through an initiating device signal detection device.

[0070] During the simulation, interpretation and verification of the satellite Earth-Moon transfer orbit flight process in modules M4 to M8, the entire Earth-Moon transfer flight process is accelerated by time jumping.

[0071] Compared with the prior art, the present invention has the following beneficial effects:

[0072] 1. The present invention can simulate and verify the Earth-Moon transfer orbit flight process of lunar exploration small satellites in view of more constraints and restrictions, examine the correctness of the control strategies at each stage of Earth-Moon transfer, and evaluate the coordination of the work of various subsystems of the small satellite and the rationality of the flight program settings.

[0073] 2. The present invention can detect possible fault events and design defects for highly complex flight missions, and ensure that the small lunar exploration satellite can smoothly enter the lunar orbit according to the set procedures after being launched into orbit.

[0074] 3. During the development phase, the present invention conducts a comprehensive evaluation and verification of the various states, modes, and polarities of the lunar exploration small satellite to ensure the reliable operation of the small satellite after it is launched into orbit, as well as the smooth execution of the mission processes in each flight phase. BRIEF DESCRIPTION OF THE DRAWINGS

[0075] Other features, objects and advantages of the present invention will become more apparent from the detailed description of non-limiting embodiments made with reference to the following drawings:

[0076] Figure 1 This is a flow chart of the closed-loop verification method for the Earth-Moon transfer orbit flight strategy;

[0077] Figure 2 This is a connection diagram of the small satellite assembly in test state;

[0078] Figure 3 A schematic block diagram of an embodiment of satellite polarity verification is shown. DETAILED DESCRIPTION

[0079] The present invention is described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those of ordinary skill in the art, several changes and improvements can also be made without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0080] The present invention is applicable to the on-orbit flight strategy verification of a small lunar orbit exploration satellite or a small satellite assembly. It can simulate the entire process of flight strategy from launch to entry into lunar orbit during the development phase of a small lunar exploration satellite for closed-loop verification, assess the correctness of the flight strategy at each stage of Earth-Moon transfer, evaluate the coordination of the work of various systems of the small lunar exploration satellite and the rationality of the flight program design, troubleshoot possible hidden dangers and design defects, and ensure that the small lunar exploration satellite can smoothly enter the lunar orbit according to the designed program after being launched into orbit. According to a closed-loop verification method for Earth-Moon transfer orbit flight strategy provided by the present invention, Figure 1 For example, the following steps are included:

[0081] Step S1: The satellite or satellite assembly is connected to the ground test system, and a closed-loop verification system is formed using a space dynamics simulation device and a satellite high-precision sensor and actuator;

[0082] Specifically, in step S1, the flight strategy verification of the Earth-Moon transfer orbit is supported in both single-satellite and double-satellite combination states. Figure 2 For example, technicians set up single-satellite or dual-satellite connection methods according to specific mission requirements. When verifying the flight mission of a dual-satellite combination, both satellites are placed on their respective parking racks, the intersatellite cables are not connected, and the satellites are connected to their respective ground test systems.

[0083] The above two states are relatively common in lunar exploration small satellite missions. Because some small satellites with extremely small weight do not have the propulsion control capabilities to autonomously transfer from Earth orbit to lunar orbit, they need to serve as auxiliary satellites and carry other spacecraft into lunar orbit. When such missions are launched, they fly as a binary satellite combination to complete the Earth-Moon transfer. After the combination enters the lunar orbit, the two satellites separate and continue to perform their respective scientific exploration missions.

[0084] Using space dynamics simulation equipment and satellite high-precision sensors and actuators, a closed-loop verification system is constructed through physical and virtual means to test product functions and performance. Figure 3 Taking the satellite measurement link, control software link, execution output link and ground simulation link as an example, the ground simulation flight closed-loop verification system is formed. This system can be used to comprehensively verify the satellite's design polarity, production polarity, assembly polarity, software polarity and interface polarity between large systems.

[0085] During the Earth-Moon transfer flight of small satellites, measurement and control resources are limited, telemetry and remote control windows are few, and many on-orbit flight actions rely more on satellite autonomous program control, with low degree of human error correction intervention. Once the polarity error occurs, it will cause catastrophic effects. Therefore, the use of this system is of great significance for the polarity verification of small satellite Earth-Moon transfer orbital flight missions. After the closed-loop system is formed, the technicians use the space dynamics simulation equipment to set up the space dynamics flight simulation environment, and equivalently simulate the inertia, center of mass and mass of another satellite after the combination, so as to support the simulation of the real flight parameters on orbit.

[0086] Step S2: Power on the satellite and set the working state of each subsystem to the weak light power-on quasi state before the satellite-rocket separation;

[0087] Specifically, in step S2, after completing the satellite-ground connection and ground equipment debugging, the technician performs the first power-on setting. The satellite MRAM is used to set the satellite-rocket separation self-start enable, satellite-rocket separation program control, and weak light power-on, etc. to ensure that the above important signs are not lost after power failure. After the setting is completed, the corresponding telemetry data is judged to see whether it meets the requirements. If the requirements are met, the satellite is powered off and the travel switch is pressed.

[0088] The process simulates the actual satellite power-off launch state, that is, due to resource limitations, the lunar exploration small satellite is not powered when the carrier rocket is launched. The satellite needs to be triggered by the satellite-rocket separation signal or weak light to turn on the discharge switch and power on.

[0089] Step S3: The simulation initial value of the space dynamics simulation device is set to an initial open-loop state;

[0090] Specifically, in step S3, the space dynamics simulation device is connected by operation to provide external electrical excitation for satellite components such as gyroscopes and star sensors, and feedback signals of flywheels, propulsion and other components are collected. Combined with the dynamic model, a satellite on-orbit closed-loop flight simulation environment is jointly constructed. By setting the device parameters, the simulation of on-orbit flight parameters such as satellite orbital parameters, inertia, center of mass and mass is completed.

[0091] Step S4: Perform power-on program control for satellite-rocket separation.

[0092] Step S4.1: Set the solar simulation array weak light curve, simulate the satellite automatically powering on after encountering weak light in the active section, and judge the execution effect of the satellite weak light power-on program control process;

[0093] Specifically, in step S4.1, the weak light curve of the solar simulation array is used to simulate the dynamic exposure and changes of sunlight after the launch of the carrier rocket and before the separation of the satellite and the rocket (i.e., the active stage), and the satellite weak light power-on function is triggered through changes in voltage and current.

[0094] Step S4.2: manually release the travel switch, inject the initial orbit parameter package, simulate the separation of the satellite and the rocket, and simultaneously set the space dynamics simulation device to a closed-loop state;

[0095] Weak-light power-on, travel switch operation, and binary-satellite assembly separation are verification capabilities for key lunar exploration strategies.

[0096] Step S5: Determine the initial orbit entry condition until the satellite autonomously switches to the flywheel coarse alignment condition;

[0097] Specifically, the signal is received by a pyrotechnic signal detection device.

[0098] Step S6: Enable the star sensor data to refresh the gyro data function to verify that the satellite has entered the "precise alignment" working condition;

[0099] The precise tracking or “precise tracking” both indicate that the satellite is in the “precise tracking” state, and the coarse tracking indicates that the satellite is in the “coarse tracking” state.

[0100] Step S7: Conducting Earth-Moon cruise flight phase verification;

[0101] The satellite and space dynamics simulation equipment jumps time, notes the cruise flight orbit parameters, enters the Earth-Moon cruise flight phase, interprets the telemetry parameters, and verifies the correctness of the satellite system polarity.

[0102] Specifically, the telemetry parameters such as orbit type, satellite position, and satellite speed are judged. If the above parameters are consistent with the cruise flight orbit parameters noted above, it is considered that the satellite's Earth-Moon cruise flight control is correct. If the telemetry parameters are inconsistent with the cruise flight orbit parameters noted above, it is judged that the Earth-Moon cruise flight control is wrong, and it is necessary to maintain this working condition and conduct problem troubleshooting.

[0103] Step S8: Perform lunar orbit braking capture phase verification;

[0104] The satellite and space dynamics simulation equipment jumps time, injects braking and capture orbit parameters, enters the lunar orbit braking and capture phase, interprets telemetry parameters, and verifies the correctness of the satellite system polarity.

[0105] Specifically, the telemetry parameters such as orbit type, satellite position, and satellite speed are judged. If the above parameters are consistent with the braking and capture orbit parameters noted above, it is considered that the satellite lunar orbit braking and capture control is correct. If the telemetry parameters are inconsistent with the braking and capture orbit parameters noted above, it is judged that the control of the lunar orbit braking and capture phase is wrong, and it is necessary to maintain this operating condition for troubleshooting.

[0106] Step S9: Perform lunar orbit binary satellite separation phase verification;

[0107] Specifically, in step S9, the pyrotechnic device signal detection device receives a two-star separation and unlocking pyrotechnic device instruction.

[0108] Verify the separation of the two satellites after the satellite reaches the lunar orbit. Use the space dynamics simulation equipment to set the changes in inertia, center of mass and mass of the two satellite combination before and after separation, and perform the separation of the two satellites under the lunar orbit. Use the upper injection thruster jet and send the thruster point jet at the separation time to simulate the disturbance force and attitude deviation caused by the separation of the two satellites. Verify whether the satellite can maintain the correct attitude before and after separation, and whether the satellite can quickly offset the disturbance in the closed-loop system. At this point, the closed-loop verification of the Earth-Moon transfer orbit flight strategy of the typical two-star combination satellite is completed.

[0109] Specifically, if the satellite attitude and orbit control program mark telemetry can always maintain the "precise pointing to the Sun" mode, or can recover to the "precise pointing to the Sun" mode after the disturbance ends, it is considered that under the closed-loop system, the satellite can quickly offset the disturbance. If the satellite attitude and orbit control program mark telemetry cannot maintain or recover the "precise pointing to the Sun" mode after the disturbance occurs, it is considered that the satellite closed-loop control system cannot overcome the disturbance caused by separation, and it is necessary to maintain this working condition and conduct problem troubleshooting.

[0110] Step S10: Change the flight parameters and select the primary and standby machines to conduct performance verification.

[0111] In more preferred examples, during the simulation, interpretation and design verification of the satellite Earth-Moon transfer orbit flight process in steps S4 to S8, the entire process of the simulated Earth-Moon transfer flight can be accelerated by time jumping without the need for a long 1:1 waiting time.

[0112] Compared with simple simulation, real-time satellite status collection through test equipment is more realistic and accurate. For key lunar exploration strategies such as low-light power-on, travel switch operation, and separation of binary satellites, equivalent verification methods and steps are proposed. From the process point of view, the verification system can fully verify all subsystems of the lunar exploration satellite, including power supply, attitude and orbit control, integrated electronics, payload, and communication. It can simulate the entire process of the lunar exploration satellite from pre-launch status setting to satellite-rocket separation and then entering the lunar orbit.

[0113] According to the needs, change the flight parameters and the main and standby single machine selection to perform performance verification. Step S4 is optional because it is difficult to determine whether sunlight can definitely pass through the carrier rocket fairing wave-transmitting hole to illuminate the satellite solar panels and trigger weak light power-on during the active phase of the actual launch of the carrier rocket. If step S4.1 is skipped, step S4.2 can be used to perform satellite-rocket separation power-on program control.

[0114] A closed-loop verification system for Earth-Moon transfer orbit flight strategy provided by the present invention comprises:

[0115] Module M1: The satellite is connected to the ground test system to form a closed-loop verification system;

[0116] Module M2: Power on the satellite and set the working state of each subsystem to the weak light power-on state before the separation of the satellite and rocket;

[0117] Module M3: Satellite connection space dynamics simulation equipment, the simulation initial value is set to the initial open loop state;

[0118] Module M4: Perform power-on program control for satellite-rocket separation;

[0119] Module M5: Determine the initial orbit insertion conditions until the satellite autonomously switches to the flywheel coarse alignment condition;

[0120] Module M6: Enable the star sensor data to refresh the gyro data function to verify that the satellite has entered the "precision pointing to the sun" condition;

[0121] Module M7: Conduct verification of the Earth-Moon cruise flight phase;

[0122] Module M8: Conduct lunar orbit braking and capture phase verification;

[0123] Module M9: Conduct lunar orbit binary separation phase verification;

[0124] Module M10: Change flight parameters and select the primary and standby units to conduct performance verification.

[0125] In more preferred embodiments, the module M4 includes:

[0126] Module M4.1: Set the solar simulation array weak light curve to judge the execution effect of the satellite weak light power-on program control process.

[0127] and\or

[0128] Module M4.2: Manually release the travel switch, inject the initial orbit parameter package, and simultaneously set the space dynamics simulation equipment to a closed-loop state.

[0129] Furthermore, in the module M1, a single-star or dual-star connection method is set according to specific mission requirements;

[0130] When conducting dual-satellite combination flight mission verification, both satellites are placed on their respective parking racks, the inter-satellite cables are not connected, and the satellites are connected to their respective ground test systems;

[0131] Connect the satellite measurement link, control software link, execution output link and ground simulation link in series to form a ground simulation flight closed-loop verification system;

[0132] The inertia, center of mass and mass of another satellite after assembly are simulated equivalently through space dynamics simulation equipment.

[0133] The module M2 performs the first power-on setting;

[0134] Through the satellite MRAM, the satellite-rocket separation self-start enable, satellite-rocket separation program control, low-light power-on and other flags are set to determine whether the corresponding telemetry data meets the requirements. If and only if the requirements are met, the satellite is powered off.

[0135] The module M3 operates and connects the space dynamics simulation equipment to provide external electrical excitation for the satellite gyroscope and star sensor, collects feedback signals from the flywheel and thruster components, and builds a satellite on-orbit closed-loop flight simulation environment in combination with the dynamics model.

[0136] Furthermore, the satellite and space dynamics simulation equipment in the module M7 jumps time, injects cruise flight orbit parameters, enters the Earth-Moon cruise flight phase, and interprets telemetry parameters to verify the correctness of the satellite system polarity.

[0137] The satellite and space dynamics simulation equipment in the module M8 jumps time, injects braking and capturing orbit parameters, enters the lunar orbit braking and capturing phase, and interprets telemetry parameters to verify the correctness of the satellite system polarity.

[0138] The module M9 receives the instruction of two-star separation and unlocking of the pyrotechnic device through the pyrotechnic device signal detection device;

[0139] The space dynamics simulation equipment is used to set the changes in inertia, center of mass and mass of the binary satellite assembly before and after separation, and the binary satellite separation is carried out under the lunar orbit. The thruster jet is injected to verify the satellite attitude before and after separation and whether the disturbance can be offset in the closed-loop system.

[0140] If the satellite attitude and orbit control program marks that the telemetry always maintains or restores the "precise sun pointing" mode, it is determined that the disturbance can be offset in the closed-loop system. If the satellite attitude and orbit control program marks that the telemetry cannot maintain or restore the "precise sun pointing" mode, it is determined that the disturbance cannot be offset, and this working condition is maintained for problem troubleshooting;

[0141] The telemetry parameters to be judged are the orbit type, satellite position and satellite speed parameters. If the parameters are consistent with the above-mentioned parameters, it is determined that the satellite control is correct. If the parameters are inconsistent with the above-mentioned parameters, it is determined that the satellite control is wrong, and this working condition is maintained for problem troubleshooting.

[0142] Furthermore, in the module M4.1, the weak light curve of the solar simulation array is used to simulate the dynamic illumination and changes of sunlight after the launch of the carrier rocket and before the separation of the satellite and the rocket, and the satellite weak light power-on function is triggered through the changes in voltage and current.

[0143] The module M5 receives the signal through an initiating device signal detection device.

[0144] During the simulation, interpretation and verification of the satellite Earth-Moon transfer orbit flight process in modules M4 to M8, the entire Earth-Moon transfer flight process is accelerated by time jumping.

[0145] The above describes the specific embodiments of the present invention. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essence of the present invention. In the absence of conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A closed-loop verification method for Earth-Moon transfer orbit flight strategy, characterized in that: include: Step S1: The satellite is connected to the ground test system to form a closed-loop verification system; Step S2: Power on the satellite and set the working state of each subsystem to the weak light power-on quasi state before the satellite-rocket separation; Step S3: The satellite is connected to a space dynamics simulation device, and the initial value of the simulation is set to an initial open-loop state; Step S4: performing power-on program control for satellite-rocket separation; Step S5: Determine the initial orbit insertion condition until the satellite autonomously switches to the flywheel coarse alignment condition; Step S6: Enable the star sensor data to refresh the gyro data function to verify that the satellite has entered the precise sun-pointing condition; Step S7: Conducting Earth-Moon cruise flight phase verification; Step S8: Perform lunar orbit braking capture phase verification; Step S9: Perform lunar orbit binary satellite separation phase verification; Step S10: Change the flight parameters and select the primary and standby machines to conduct performance verification.

2. The closed-loop verification method for Earth-Moon transfer orbit flight strategy according to claim 1, characterized in that: The step S4 includes: Step S4.1: Set the solar simulation array weak light curve to judge the execution effect of the satellite weak light power-on program control process; and\or Step S4.2: Manually release the travel switch, inject the initial orbit parameter package, and simultaneously set the space dynamics simulation device to a closed-loop state.

3. The closed-loop verification method for Earth-Moon transfer orbit flight strategy according to claim 1, characterized in that: In step S1, a single-satellite or dual-satellite connection method is set according to specific task requirements; When conducting dual-satellite combination flight mission verification, both satellites are placed on their respective parking racks, the inter-satellite cables are not connected, and the satellites are connected to their respective ground test systems; Connect the satellite measurement link, control software link, execution output link and ground simulation link in series to form a ground simulation flight closed-loop verification system; Use space dynamics simulation equipment to simulate the inertia, center of mass and mass of another satellite combination; In step S2, the first power-on setting is performed; Set the satellite-rocket separation self-start enable, satellite-rocket separation program control, low-light power-on and other flags through the satellite MRAM to determine whether the corresponding telemetry data meets the requirements. If and only if the requirements are met, power off the satellite. In step S3, the space dynamics simulation equipment is operated to provide external electrical excitation for the satellite gyroscope and star sensor, and feedback signals from the flywheel and thruster components are collected, and a satellite on-orbit closed-loop flight simulation environment is constructed in combination with the dynamics model.

4. The closed-loop verification method for Earth-Moon transfer orbit flight strategy according to claim 1, characterized in that: In step S7, the satellite and the space dynamics simulation device jump time, record the cruise flight orbit parameters, enter the Earth-Moon cruise flight phase, and interpret the telemetry parameters to verify the correctness of the satellite system polarity; In step S8, the satellite and the space dynamics simulation device jump time, add braking and capturing orbit parameters, enter the lunar orbit braking and capturing phase, and interpret telemetry parameters to verify the correctness of the satellite system polarity; In the step S9, a two-star separation unlocking pyrotechnic device instruction is received through the pyrotechnic device signal detection device; The space dynamics simulation equipment is used to set the changes in inertia, center of mass and mass of the binary satellite assembly before and after separation, and the binary satellite separation is carried out under the lunar orbit. The thruster jet is injected to verify the satellite attitude before and after separation and whether the disturbance can be offset in the closed-loop system. If the satellite attitude and orbit control program indicates that the telemetry always maintains or restores the precise solar pointing mode, it is determined that the disturbance can be offset in the closed-loop system. If the satellite attitude and orbit control program indicates that the telemetry cannot maintain or restore the precise solar pointing mode, it is determined that the disturbance cannot be offset and this working condition is maintained for troubleshooting. The telemetry parameters to be judged are the orbit type, satellite position and satellite speed parameters. If the parameters are consistent with the above-mentioned parameters, it is determined that the satellite control is correct. If the parameters are inconsistent with the above-mentioned parameters, it is determined that the satellite control is wrong, and this working condition is maintained for problem troubleshooting.

5. The closed-loop verification method for Earth-Moon transfer orbit flight strategy according to claim 2, characterized in that: In step S4.1, the dynamic irradiation and change of sunlight after the launch of the carrier rocket and before the separation of the satellite and the rocket are simulated by using the weak light curve of the solar simulation array, and the satellite weak light power-on function is triggered by the change of voltage and current; In the step S5, the signal is received by the pyrotechnic device signal detection device; During the simulation, interpretation and verification of the satellite Earth-Moon transfer orbit flight process in steps S4 to S8, the entire Earth-Moon transfer flight process is accelerated by time hopping.

6. A closed-loop verification system for Earth-Moon transfer orbit flight strategy, characterized in that: include: Module M1: The satellite is connected to the ground test system to form a closed-loop verification system; Module M2: Power on the satellite and set the working state of each subsystem to the weak light power-on state before the separation of the satellite and rocket; Module M3: Satellite connection space dynamics simulation equipment, the simulation initial value is set to the initial open loop state; Module M4: Perform power-on program control for satellite-rocket separation; Module M5: Determine the initial orbit insertion conditions until the satellite autonomously switches to the flywheel coarse alignment condition; Module M6: Enable the star sensor data to refresh the gyro data function to verify that the satellite has entered the precise sun-pointing condition; Module M7: Conduct verification of the Earth-Moon cruise flight phase; Module M8: Conduct lunar orbit braking and capture phase verification; Module M9: Conduct lunar orbit binary separation phase verification; Module M10: Change flight parameters and select the primary and standby units to conduct performance verification.

7. The closed-loop verification system for Earth-Moon transfer orbit flight strategy according to claim 6, characterized in that: The module M4 includes: Module M4.1: Set the solar simulation array weak light curve to judge the execution effect of the satellite weak light power-on program control process; and\or Module M4.2: Manually release the travel switch, inject the initial orbit parameter package, and simultaneously set the space dynamics simulation equipment to a closed-loop state.

8. The closed-loop verification system for Earth-Moon transfer orbit flight strategy according to claim 6, characterized in that: In the module M1, a single-satellite or dual-satellite connection method is set according to specific mission requirements; When conducting dual-satellite combination flight mission verification, both satellites are placed on their respective parking racks, the inter-satellite cables are not connected, and the satellites are connected to their respective ground test systems; Connect the satellite measurement link, control software link, execution output link and ground simulation link in series to form a ground simulation flight closed-loop verification system; Use space dynamics simulation equipment to simulate the inertia, center of mass and mass of another satellite combination; The module M2 performs the first power-on setting; Set the satellite-rocket separation self-start enable, satellite-rocket separation program control, low-light power-on and other flags through the satellite MRAM to determine whether the corresponding telemetry data meets the requirements. If and only if the requirements are met, power off the satellite. The module M3 operates and connects the space dynamics simulation equipment to provide external electrical excitation for the satellite gyroscope and star sensor, collects feedback signals from the flywheel and thruster components, and builds a satellite on-orbit closed-loop flight simulation environment in combination with the dynamics model.

9. The closed-loop verification system for Earth-Moon transfer orbit flight strategy according to claim 6, characterized in that: The satellite and space dynamics simulation equipment in the module M7 jumps time, notes the cruise flight orbit parameters, enters the Earth-Moon cruise flight phase, and interprets the telemetry parameters to verify the correctness of the satellite system polarity; In the module M8, the satellite and space dynamics simulation equipment jumps time, injects braking and capturing orbit parameters, enters the lunar orbit braking and capturing phase, and interprets telemetry parameters to verify the correctness of the satellite system polarity; The module M9 receives the instruction of two-star separation and unlocking of the pyrotechnic device through the pyrotechnic device signal detection device; The space dynamics simulation equipment is used to set the changes in inertia, center of mass and mass of the binary satellite assembly before and after separation, and the binary satellite separation is carried out under the lunar orbit. The thruster jet is injected to verify the satellite attitude before and after separation and whether the disturbance can be offset in the closed-loop system. If the satellite attitude and orbit control program indicates that the telemetry always maintains or restores the precise solar pointing mode, it is determined that the disturbance can be offset in the closed-loop system. If the satellite attitude and orbit control program indicates that the telemetry cannot maintain or restore the precise solar pointing mode, it is determined that the disturbance cannot be offset and this working condition is maintained for troubleshooting. The telemetry parameters to be judged are the orbit type, satellite position and satellite speed parameters. If the parameters are consistent with the above-mentioned parameters, it is determined that the satellite control is correct. If the parameters are inconsistent with the above-mentioned parameters, it is determined that the satellite control is wrong, and this working condition is maintained for problem troubleshooting.

10. The closed-loop verification system for Earth-Moon transfer orbit flight strategy according to claim 7, characterized in that: In the module M4.1, the dynamic irradiation and change of sunlight from the launch of the carrier rocket to the separation of the satellite and the rocket are simulated by using the weak light curve of the solar simulation array, and the satellite weak light power-on function is triggered by the change of voltage and current; The module M5 receives the signal through the pyrotechnic device signal detection device; During the simulation, interpretation and verification of the satellite Earth-Moon transfer orbit flight process in modules M4 to M8, the entire Earth-Moon transfer flight process is accelerated by time jumping.

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