Static Closed-loop Verification Method and System for Mars Exploration Separation Strategy

Abstract

The present invention provides a method and system for static closed-loop verification of the separation strategy for Mars exploration, including the following steps: Step S1: Verify the separation process of the orbiter and the lander; Step S2: Perform static equivalent separation of the orbiter and the lander according to the verified separation process. The present invention can deduce and verify the separation process between spacecraft during the development stage of Mars probes. The two spacecraft are statically placed on their respective parking racks. By operating the separation device between the two spacecraft and using the closed-loop verification system to collect separation data, the correctness of the separation control strategy between spacecraft can be examined, the coordination of the joint operation of the two spacecraft during the separation process and the rationality of the flight program design can be evaluated, possible fault events and design defects can be investigated, and it is ensured that the Mars probe can successfully execute the separation process between spacecraft according to the designed program after being launched into orbit.

Description

Technical Field

[0001] The present invention relates to the technical field of spacecraft testing, and in particular, to a static closed-loop verification method and system for a Mars exploration separation strategy. Background Art

[0002] The detector for China's first Mars exploration adopts the form of a combination of an orbiter and a lander-rover. The mission objective is to successfully send the lander-rover into a safe landing orbit after Mars capture, complete the relay communication service during the lifespan of the Mars rover and the global remote sensing mission of the orbiter itself. To achieve the above mission objectives, the Mars detector needs to go through multiple flight stages such as escaping the Earth, cruising around the sun, capturing Mars, adjusting the orbit around Mars, separating the spacecraft, and autonomously ascending the orbit. The flight process is complex and the technical difficulty is high. The success or failure of any event may affect the success or failure of the entire mission. Therefore, it is necessary to comprehensively evaluate and verify various states, modes, and characteristics of the Mars detector during the development stage to ensure the reliable operation after the Mars exploration is launched into orbit and the smooth execution of the mission processes in each flight stage.

[0003] After retrieving the prior art, the Chinese patent literature with the publication number CN104865841A discloses a method for rapid joint simulation verification of the overall scheme of a microsatellite. This method is based on parallel simulation of each subsystem of the microsatellite design scheme using the system hardware with a general simulator architecture, and realizes the joint closed-loop simulation of multiple different subsystems by uploading different models, such as simulating subsystems such as attitude and orbit control subsystem, power supply subsystem, thermal control subsystem, and TT&C subsystem.

[0004] The Chinese patent literature with the publication number CN105334756A discloses an agile satellite mission interpretation closed-loop simulation verification system and method. First, simulate the mission interpretation function of the on-board computer, replace the traditional on-board computer test bed, interpret the mission block into an instruction set executable by the on-board computer, and intuitively display information such as instruction name, parameters, and execution time in a tabular form, automatically judge the mission block interpretation result and output the abnormal information to avoid uploading incorrect mission blocks to the satellite. Secondly, the parameters and settings in the mission block can be modified through this system, the mission block can be regenerated and iteratively analyzed to ensure its correctness, avoiding the risks brought by manual arrangement of mission blocks, and improving the efficiency and reliability of mission block generation. Finally, compare the interpretation result of the finally generated correct mission block with the action sequence generated by the mission planning system to verify the correctness and rationality of the mission planning.

[0005] The Chinese invention patent document with the publication number CN106211220A discloses a heterogeneous interconnection satellite broadband communication semi-physical simulation verification system and method. The system includes a ground heterogeneous Internet gateway, a simulated satellite broadband communication subsystem, a ground-simulated satellite broadband communication subsystem, a real satellite broadband communication subsystem, and a ground network. The ground heterogeneous Internet gateway realizes the interconnection of different heterogeneous networks. The simulated satellite broadband communication subsystem realizes the network simulation of the broadband satellite communication subsystem. The ground-simulated satellite broadband communication subsystem realizes the ground simulation of the broadband satellite communication subsystem. The real satellite broadband communication subsystem accesses the services and applications of the real satellite broadband communication subsystem. The ground network introduces the services and applications of the ground network to solve the key technical problems in the heterogeneous interconnection of satellite broadband communication systems.

[0006] The Chinese invention patent document with the publication number CN108983797A discloses a method for on-orbit autonomous separation discrimination of a Mars probe, including the following steps: Step 1: Upload separation parameters at the ground TT&C station and set the state before separation; Step 2: Implement active thermal control on the swing bar unlocking pyrotechnics 30 minutes before the predetermined separation moment; Step 3: Turn on the telemetry monitoring probe 5 minutes before the predetermined separation moment; Step 4: After reaching the predetermined separation moment, according to the set separation parameters, the swing bar unlocking pyrotechnics detonates; Step 5: Judge whether the signal of the swing bar travel switch is triggered. If triggered, confirm that the swing bar is fully deployed and transfer to Step 7. If not triggered, transfer to Step 6.

[0007] Regarding the above related technologies, the inventor believes that the verification methods in the above invention patent documents are only applicable to low-earth orbit satellites and do not have the function of verifying the separation strategy of the Mars probe in the form of a combined body, and cannot deduce and verify the separation strategy between the orbiter and the lander. Currently, no other similar related technical descriptions or reports have been found, nor have other similar materials at home and abroad been collected. Summary of the Invention

[0008] Aiming at the defects in the prior art, the purpose of the present invention is to provide a static closed-loop verification method and system for the Mars exploration separation strategy.

[0009] A static closed-loop verification method for the Mars exploration separation strategy provided by the present invention includes the following steps:

[0010] Step S1: Verify the separation process of the orbiter and the lander;

[0011] Step S2: Perform static equivalent separation on the orbiter and the lander according to the verified separation process.

[0012] Preferably, the step S1 includes the following steps:

[0013] First predetermined time setting step: Power on the orbiter and the lander, simulate the state setting at the first predetermined time before the separation of the orbiter and the lander, and synchronize and set the on-board time of the orbiter and the lander to the first predetermined time moment before separation;

[0014] Uploading step: Simulate the workflow at the first predetermined time before the separation of the orbiter and the lander, upload a set of delay instructions scheduled to be executed during the separation process to the orbiter and the lander, and upload separation control strategy parameters to the orbiter;

[0015] Verification step: Activate the data download function of the orbiter and the lander, verify that the set of delay instructions and the separation control strategy parameters received by the orbiter are correct, and verify that the set of delay instructions received by the lander is correct;

[0016] Second predetermined time jump step: Jump the on-board time of the orbiter and the lander to the second predetermined time moment before separation;

[0017] Pre-separation monitoring step: Monitor the operating states of the orbiter and the lander according to the working time sequence of the on-orbit flight program before separation, confirm that the orbiter and the lander correctly execute the pre-separation delay instructions at the predetermined time, and confirm that the orbiter executes the pre-separation deorbiting attitude maneuver, deorbiting orbit control, and separation attitude maneuver according to the predetermined working time sequence and separation control strategy;

[0018] Separation state setting step: When the on-board time of the orbiter and the lander runs to the separation moment, send a separation signal;

[0019] Program-controlled task flow triggering step: Confirm that both the orbiter and the lander detect the separation signal and correctly trigger the program-controlled task flow after separation;

[0020] Post-separation monitoring step: Monitor the operating states of the orbiter and the lander according to the working time sequence of the on-orbit flight program after separation, confirm that the lander correctly executes the program-controlled task flow according to the predetermined working time sequence, and confirm that the orbiter correctly executes the post-separation delay instructions at the predetermined time, and executes the post-separation ascending orbit attitude maneuver, ascending orbit control, and post-ascending orbit attitude recovery according to the predetermined working time sequence and separation control strategy parameters;

[0021] Verification completion step: After the verification of the separation strategy process of the orbiter and the lander is completed, restore the states of the orbiter and the lander and implement power-off.

[0022] Preferably, the method further includes an internal power supply setting step;

[0023] In the first predetermined time setting step, after powering on the orbiter, the orbiter supplies power to the lander;

[0024] In the internal power supply setting step, according to the working sequence of the on-orbit flight program before separation, the lander battery discharge switch is turned on, and the orbiter's power supply switch to the lander is turned off, so that the lander is set to the internal battery power supply state.

[0025] Preferably, the method also includes a troubleshooting step: if there is an inconsistency between the operating process and the predetermined working sequence in the separation process of the orbiter and the lander, the status of the orbiter and the lander are restored and the power is cut off after troubleshooting, and the separation process verification of the orbiter and the lander is restarted from the first predetermined time setting step.

[0026] Preferably, in the injection step, the delay instruction group of the lander is forwarded to the lander through the orbiter.

[0027] A static closed-loop verification system for Mars exploration separation strategy provided by the present invention includes the following modules:

[0028] Module M1: Verify the separation process of the orbiter and lander;

[0029] Module M2: Perform static equivalent separation of the orbiter and lander according to the verified separation process.

[0030] Preferably, the module M1 includes the following modules:

[0031] The first scheduled time setting module: powers on the orbiter and the lander, simulates the first scheduled time state setting before the separation of the orbiter and the lander, and synchronizes the onboard time of the orbiter and the lander to the first scheduled time before the separation;

[0032] Injection module: simulates the workflow of the first scheduled time before the separation of the orbiter and the lander, injects the delay instruction group scheduled to be executed in the separation process into the orbiter and the lander, and injects the separation control strategy parameters into the orbiter;

[0033] Confirmation module: Start the data unloading function of the orbiter and the lander, confirm that the delay instruction group and separation control strategy parameters received by the orbiter are correct, and confirm that the delay instruction group received by the lander is correct;

[0034] Second scheduled time jump module: jump the orbiter and lander's onboard time to the second scheduled time before separation;

[0035] Pre-separation monitoring module: monitors the operating status of the orbiter and lander according to the working sequence of the on-orbit flight program before separation, confirms that the orbiter and lander correctly execute the delay instructions before separation at the scheduled time, and confirms that the orbiter performs the orbit lowering attitude maneuver, orbit lowering orbit control and separation attitude maneuver before separation according to the scheduled working sequence and separation control strategy;

[0036] Separation state setting module: When the on-board time of the orbiter and the lander runs to the separation moment, it issues a separation signal;

[0037] Program-controlled task process trigger module: It confirms that both the orbiter and the lander detect the separation signal and correctly trigger the program-controlled task process after separation;

[0038] Post-separation monitoring module: According to the working timing sequence of the on-orbit flight program after separation, it monitors the operating states of the orbiter and the lander, confirms that the lander correctly executes the program-controlled task process according to the predetermined working timing sequence, and confirms that the orbiter correctly executes the post-separation delay instruction at the predetermined moment, and performs the post-separation ascending orbit attitude maneuver, ascending orbit control, and post-ascending orbit attitude recovery according to the predetermined working timing sequence and separation control strategy parameters;

[0039] Verification completion module: After the separation strategy process of the orbiter and the lander is verified, it restores the states of the orbiter and the lander and implements power-off.

[0040] Preferably, the system further includes an internal power supply setting module;

[0041] In the first predetermined time setting module, after powering on the orbiter, the orbiter supplies power to the lander;

[0042] In the internal power supply setting module, according to the working timing sequence of the on-orbit flight program before separation, it turns on the discharge switch of the lander's battery and turns off the power supply switch of the orbiter to the lander, and sets the lander to the internal battery power supply state.

[0043] Preferably, the system further includes a troubleshooting module: If there is an inconsistency between the operation process and the predetermined working timing sequence in the separation process of the orbiter and the lander, after troubleshooting, it restores the states of the orbiter and the lander and implements power-off, and starts the verification of the separation process of the orbiter and the lander again from the first predetermined time setting module.

[0044] Preferably, in the injection module, the delay instruction group of the lander is forwarded to the lander through the orbiter.

[0045] Compared with the prior art, the present invention has the following beneficial effects:

[0046] 1. The present invention can deduce and verify the separation process of the spacecraft during the development stage of the Mars probe, and verify the correctness of the separation control strategy of the spacecraft;

[0047] 2. The present invention can evaluate the coordination of the joint work of the two spacecraft during the separation process and the rationality of the flight program design;

[0048] 3. The present invention can detect possible fault events and design defects. The orbiter and the lander are statically placed on their respective parking racks. By operating the separation device between the two vehicles and using the closed-loop verification system to collect separation data, it is ensured that after the Mars probe is launched into orbit, the separation process of the two vehicles can be successfully executed according to the designed procedure. BRIEF DESCRIPTION OF THE DRAWINGS

[0049] Other features, objects, and advantages of the present invention will become more apparent by reading the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0050] Figure 1 is a flowchart of the method for verifying the separation strategy deduction between the orbiter and the lander of the present invention;

[0051] Figure 2 is a connection diagram of the test states of the two vehicles corresponding to the method for verifying the separation strategy deduction between the orbiter and the lander of the present invention;

[0052] Figure 3 is a block diagram of an embodiment of the separation signal simulation device of the two vehicles corresponding to the method for verifying the separation strategy deduction between the orbiter and the lander of the present invention;

[0053] Figure 4 is a block diagram of an embodiment of the RF link attenuation adjustment device corresponding to the method for verifying the separation strategy deduction between the orbiter and the lander of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0054] The present invention will be described in detail below with reference to specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but do not limit the present invention in any form. It should be noted that those of ordinary skill in the art can make several changes and improvements without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0055] An embodiment of the present invention discloses a static closed-loop verification method for the Mars exploration separation strategy, as shown in Figure 1 and Figure 2 and includes the following steps: Step S1: Verify the separation process of the orbiter and the lander. Step S1 includes the following steps: First predetermined time setting step: Power on the orbiter and the lander, simulate the state setting at the first predetermined time before the separation of the orbiter and the lander, and synchronize and set the on-board time of the orbiter and the lander to the first predetermined time moment before the separation. After powering on the orbiter, the orbiter supplies power to the lander. That is, the orbiter and the lander are powered on, the state is set by simulating 11 hours before the separation, and the on-board time is synchronized and set to the moment 11 hours before the separation.

[0056] Specifically, before power-on, the orbiter and the lander are statically placed on their respective parking racks. Each of the two vehicles establishes a ground connection with the equipment of its respective test system. The low-frequency signal interfaces between the two vehicles are interconnected through a process cable and a separation signal simulation device for the two vehicles. The RF signal interfaces between the two vehicles are interconnected through an RF cable and an RF link attenuation adjustment device. The connection status block diagram of the two vehicles is as shown in Figure 2 the figure. Among them, the attitude and orbit control dynamics subsystem performs recursive simulation on the orbital environment during the separation process of the two vehicles, collects the operation signals of the actuators of the attitude and orbit control subsystem of the orbiter, and outputs simulation signals to each sensor of the attitude and orbit control subsystem according to the separation orbit model of the two vehicles; the TT&C communication subsystem includes a master control terminal for generating remote control commands, a telemetry terminal for receiving and analyzing telemetry data, a TT&C baseband device for modulating remote control signals and demodulating telemetry signals, and RF link devices such as an upconverter (such as the UX9000 upconverter of German WORK Company), a downconverter (such as the DX9000 downconverter of German WORK Company), and an adjustable attenuator (such as the 8496B and 8494B adjustable attenuators of KeightSight Company); the mechanism simulation subsystem is used to simulate the rotation of the solar array drive mechanism and the antenna drive mechanism during the separation process of the two vehicles; the solar array power supply subsystem consists of 18 E4361 DC power modules of KeightSight Company, which are used to simulate the power supply of the solar array to the orbiter; the bus monitoring device is connected to the on-board computer through a low-frequency cable to monitor the operation status of the on-board computer software during the separation process of the two vehicles; the pyrotechnic equivalent device is connected to the on-board power distribution and supply subsystem through a low-frequency cable to detect the correctness of the pyrotechnic unlocking command pulses for time delay and program control during the separation process of the two vehicles.

[0057] Before power-on of the orbiter and the lander, the separation signal simulation device for the two vehicles is set to the pre-separation state. An example of the separation signal simulation device for the two vehicles is as shown in Figure 3 the figure. The low-frequency signal interface of the orbiter and the low-frequency signal interface of the lander are respectively connected to the two low-frequency electrical connectors of the separation signal simulation device for the two vehicles through low-frequency cables. The separation signal simulation device for the two vehicles respectively transfers the positive and negative contacts of the separation signals in the low-frequency signal interfaces of the orbiter and the lander to the jacks. If the separation signal short-circuit plug is inserted into the separation signal jack to short-circuit the positive and negative lines of the separation signal, it is the pre-separation state of the two vehicles. If the separation signal short-circuit plug is removed to disconnect the positive and negative lines of the separation signal, it is the post-separation state of the two vehicles. In the first predetermined time setting step, the separation signal short-circuit plug is inserted in the separation signal simulation device for the two vehicles and set to the pre-separation state.

[0058] The orbiter and the lander are respectively powered on through the ground test system to simulate the on-orbit state, set the working states of each subsystem on the vehicles before the separation of the two vehicles, and simulate the on-orbit time 11 hours before the separation of the two vehicles to time the orbiter and set the time synchronization of the two vehicles.

[0059] The orbiter and lander are powered on in the following way: the orbiter is powered on by connecting the power of the solar simulation array electronic system. After the orbiter is powered on, it receives the orbiter power-on command sent by the orbiter test system. The orbiter's power supply and distribution subsystem turns on the power switch for the lander and powers the lander through the low-frequency signal interface between the two devices.

[0060] After the orbiter and the lander are powered on, the two devices are in the state before separation. In the state before separation, the telemetry data of the two devices are transmitted as follows: the telemetry data of the orbiter is transmitted to the orbiter's measurement, control and communication subsystem, and the measurement, control and communication subsystem parses and displays the telemetry status of the orbiter; while the telemetry data of the lander is sent to the orbiter through the low-frequency signal interface between the two devices, and the orbiter forwards it to the orbiter's measurement, control and communication subsystem. The orbiter's measurement, control and communication subsystem distributes the lander's telemetry to the lander's measurement, control and communication subsystem for telemetry analysis and display. The orbiter's measurement, control and communication subsystem and the lander's measurement, control and communication subsystem communicate data through the local area network. Specifically, the lander's telemetry terminal is connected to the orbiter's measurement and control baseband equipment through the local area network. The two communicate telemetry and remote control data according to the TCP / IP protocol (Transmission Control Protocol / Internet Protocol).

[0061] Injection step: simulate the workflow of the first scheduled time before the separation of the orbiter and the lander, inject the delay instruction group scheduled to be executed in the separation process into the orbiter and the lander, and inject the separation control strategy parameters into the orbiter. That is, inject the delay instruction group and separation control strategy parameters into the orbiter and the lander.

[0062] Specifically, the workflow of the two vehicles 11 hours before separation is simulated, and the delay instruction group scheduled to be executed in the separation control process is injected into the orbiter and the lander, and the separation control strategy parameters of the two vehicles are injected into the orbiter, including attitude maneuver parameters and orbit control parameters. The delay instruction group of the lander is forwarded to the lander through the orbiter. The orbiter and the lander are in the state before separation. The lander delay instruction is injected in the following way: the lander delay instruction group is sent to the orbiter through the orbiter's measurement, control and communication subsystem. The orbiter determines that the instruction is a lander instruction based on the spacecraft identifier in the delay instruction frame, and forwards it to the lander through the inter-vehicle low-frequency signal interface.

[0063] Confirmation steps: Activate the data download functions of the orbiter and the lander, confirm that the received delay command group and separation control strategy parameters of the orbiter are correct, and confirm that the received delay command group of the lander is correct. That is, activate the data download functions of the orbiter and the lander, and download the pre-injected delay command group and the separation control strategy parameters of the two vehicles to the ground through the downlink telemetry channel, and confirm that the received delay command content of the orbiter and the lander and the separation control strategy parameters are consistent with the predetermined content. Activate the data download functions of the orbiter and the lander, and confirm the correctness of the received delay commands and separation control strategy parameters of the orbiter and the lander.

[0064] Second predetermined time hopping step: Hopping the on-board time of the orbiter and the lander to the second predetermined time before separation. That is, set the simulation start time and simulation start parameters of the attitude control dynamics simulation system to 5 hours before the separation of the two vehicles, synchronously hop the on-board time of the orbiter and the lander to 5 hours before the separation of the two vehicles, and start the attitude control dynamics simulation. Specifically, set the simulation start time and simulation start parameters of the attitude and orbit control dynamics simulation system to the state 5 hours before the separation of the two vehicles. After the setting is completed, hop the on-board time of the orbiter and the lander to 5 hours before the separation of the two vehicles, and synchronously start the simulation process of the attitude and orbit control dynamics simulation system.

[0065] In the internal power supply setting step, according to the working timing of the on-orbit flight program before separation, turn on the battery discharge switch of the lander and turn off the power supply switch of the orbiter to the lander, and set the lander to the internal battery power supply state.

[0066] Pre-separation monitoring steps: According to the working timing of the on-orbit flight program before separation, monitor the operating states of the orbiter and the lander, confirm that the orbiter and the lander correctly execute the pre-separation delay commands at the predetermined time, and confirm that the orbiter gradually executes the pre-separation deorbiting attitude maneuver, deorbiting orbit control, and separation attitude maneuver strategies according to the predetermined working timing and separation control strategy parameters. That is, according to the working timing of the on-orbit flight program before the separation of the two vehicles, monitor the operating states of the orbiter and the lander, confirm that the orbiter and the lander correctly execute the pre-separation delay commands of the two vehicles at the predetermined time, and confirm that the orbiter gradually executes the pre-separation deorbiting attitude maneuver, deorbiting orbit control, and separation attitude maneuver strategies of the two vehicles according to the predetermined working timing.

[0067] Separation state setting steps: When the on-board time of the orbiter and the lander runs to the separation moment, a separation signal is sent. That is, when the on-board time of the orbiter and the lander runs to the separation moment of the two vehicles, manually operate the separation device between the orbiter and the lander, set the separation signal simulation device between the two vehicles from the state before separation to the state after separation, and send a separation signal. The setting of the separation signal simulation device between the two vehicles from the state before separation to the state after separation is achieved by disconnecting the contact connection corresponding to the separation signal in the low-frequency signal interface between vehicles, that is, setting the separation signal between the two vehicles from the short-circuit state to the open-circuit state. The operation method of setting the separation signal simulation device between the two vehicles from the state before separation to the state after separation is to remove the separation signal short-circuit plug and disconnect the positive and negative lines of the two-way separation signal of the orbiter and the lander. The operation time of step 7 needs to be strictly carried out according to the time set by the flight program, and the tolerable deviation is 1 minute. Otherwise, the orbiter and the lander will fail to collect the separation signal and enter the program control task process for fault handling.

[0068] Program control task process triggering steps: Confirm that both the orbiter and the lander detect the separation signal and correctly trigger the program control task process after separation. That is, confirm that both the orbiter and the lander detect the separation signal between the two vehicles and correctly trigger the program control task process after the separation of the two vehicles.

[0069] According to the working time sequence of the on-orbit flight program after the separation of the two vehicles, monitor the operating states of the orbiter and the lander, confirm that the lander correctly executes the program control task process according to the predetermined time sequence, confirm that the orbiter correctly executes the delay instruction after the separation of the two vehicles at the predetermined moment, and gradually execute the ascending orbit attitude maneuver, ascending orbit control, and attitude recovery after ascending orbit according to the predetermined time sequence.

[0070] Specifically, the orbiter and the lander are in the state of simulated separation of the two vehicles. The downlink channel of the lander telemetry data is switched from the low-frequency signal interface between vehicles to the radio frequency signal interface between vehicles. That is, the lander telemetry data is sent to the orbiter through the radio frequency signal interface between vehicles, and the orbiter transmits the lander telemetry data to the orbiter's measurement and control communication subsystem, which is distributed to the lander's measurement and control communication subsystem through the orbiter's measurement and control communication subsystem for telemetry data parsing and display.

[0071] Examples of the implementation of the radio frequency link attenuation adjustment device between vehicles are as Figure 4 shown, including a high-power fixed attenuator, a circulator, and a variable attenuator. Among them, the high-power fixed attenuator is used to attenuate the radio frequency signal between the orbiter and the lander by 20 dB, the power capacity that can be tolerated is 50 W, and the applicable frequency band covers 0 - 4 GHz; the circulator is used to realize the adaptation and transfer between the two-port transceiver unidirectional signal of the orbiter and the single-port transceiver bidirectional signal of the lander; the variable attenuator performs step-by-step adjustment on the radio frequency signal between vehicles, the step accuracy is 1 dB, and the adjustable range is 0 - 121 dB.

[0072] Post-separation monitoring step: Monitor the operating states of the orbiter and the lander according to the working time sequence of the post-separation on-orbit flight program, confirm that the lander correctly executes the programmed task process according to the predetermined working time sequence, and confirm that the orbiter correctly executes the post-separation delay instruction at the predetermined time, and gradually executes the post-separation ascending orbit attitude maneuver, ascending orbit control, and post-ascending orbit attitude recovery according to the predetermined working time sequence and separation control strategy parameters. That is, monitor the operating states of the programmed tasks and delay instructions after the separation of the orbiter and the lander according to the working time sequence of the post-separation on-orbit flight program of the two vehicles.

[0073] Troubleshooting step: If there is a discrepancy between the operating process and the predetermined working time sequence in the separation process of the orbiter and the lander, perform troubleshooting, then implement the state recovery and power-off of the orbiter and the lander, and restart the verification of the separation process of the orbiter and the lander from the first predetermined time setting step. That is, if there is a discrepancy between the operating process and the predetermined working time sequence during the deduction process of the above-mentioned separation process of the orbiter and the lander, it is necessary to perform troubleshooting, then implement the state recovery and power-off of the two vehicles, and restart the deduction process of the separation process control strategy of the two vehicles from the first predetermined time setting step.

[0074] Verification completion step: After the separation strategy process of the orbiter and the lander is verified, perform the state recovery of the orbiter and the lander, and implement the power-off.

[0075] Step S2: Perform a static equivalent separation on the orbiter and the lander according to the verified separation process.

[0076] The embodiment of the present invention also discloses a static closed-loop verification system for a Mars exploration separation strategy, including the following modules:

[0077] A static closed-loop verification system for a Mars exploration separation strategy provided by the present invention includes the following modules:

[0078] Module M1: Verify the separation process of the orbiter and the lander. Module M1 includes the following modules: First predetermined time setting module: Power on the orbiter and the lander, simulate the state setting at the first predetermined time before the separation of the orbiter and the lander, and synchronize and set the on-board time of the orbiter and the lander to the first predetermined time moment before the separation. After powering on the orbiter, the orbiter supplies power to the lander.

[0079] Upload module: Simulate the working process at the first predetermined time before the separation of the orbiter and the lander, upload the delay instruction group scheduled to be executed in the separation process to the orbiter and the lander, and upload the separation control strategy parameters to the orbiter. The delay instruction group of the lander is forwarded to the lander through the orbiter.

[0080] Confirmation module: Activate the data download function of the orbiter and the lander, confirm that the delayed command group and separation control strategy parameters received by the orbiter are correct, and confirm that the delayed command group received by the lander is correct.

[0081] Second predetermined time hopping module: Hopping the on-board time of the orbiter and the lander to the second predetermined time moment before separation.

[0082] Internal power supply setting module: According to the working timing of the on-orbit flight program before separation, turn on the battery discharge switch of the lander and turn off the power supply switch of the orbiter to the lander, and set the lander to the internal battery power supply state.

[0083] Pre-separation monitoring module: According to the working timing of the on-orbit flight program before separation, monitor the operating states of the orbiter and the lander, confirm that the orbiter and the lander correctly execute the delayed commands before separation at the predetermined moment, and confirm that the orbiter executes the deorbiting attitude maneuver, deorbiting orbit control, and separation attitude maneuver before separation according to the predetermined working timing and separation control strategy.

[0084] Separation state setting module: When the on-board time of the orbiter and the lander runs to the separation moment, send a separation signal.

[0085] Program-controlled task flow trigger module: Confirm that both the orbiter and the lander detect the separation signal and correctly trigger the program-controlled task flow after separation.

[0086] Post-separation monitoring module: According to the working timing of the on-orbit flight program after separation, monitor the operating states of the orbiter and the lander, confirm that the lander correctly executes the program-controlled task flow according to the predetermined working timing, and confirm that the orbiter correctly executes the delayed command after separation at the predetermined moment, and executes the ascending orbit attitude maneuver, ascending orbit control, and post-ascending orbit attitude recovery after separation according to the predetermined working timing and separation control strategy parameters.

[0087] Troubleshooting module: If there is a discrepancy between the operation process and the predetermined working timing in the separation process of the orbiter and the lander, after troubleshooting, implement the state recovery and power-off of the orbiter and the lander, and restart the verification of the separation process of the orbiter and the lander from the first predetermined time setting module.

[0088] Verification completion module: After the verification of the separation strategy process of the orbiter and the lander is completed, perform the state recovery of the orbiter and the lander and implement the power-off.

[0089] Module M2: Perform a static equivalent separation on the orbiter and the lander according to the verified separation process.

[0090] The present invention provides a method for deducing and verifying the separation strategy between an orbiter and a lander, aiming to meet the test and verification requirements to a certain extent during the development stage of a Mars probe. The present invention can deduce and verify the separation process between the spacecrafts during the development stage of the Mars probe, examine the correctness of the separation control strategy between the spacecrafts, evaluate the coordination of the joint operation of the two spacecrafts and the rationality of the flight program design during the separation process between the spacecrafts, identify possible fault events and design defects, and ensure that the Mars probe can successfully execute the separation process between the spacecrafts according to the designed program after being launched into orbit.

[0091] Those skilled in the art know that in addition to implementing the system and its various devices, modules, and units provided by the present invention in the form of pure computer-readable program code, the method steps can be logically programmed to enable the system and its various devices, modules, and units provided by the present invention to be implemented in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, embedded microcontrollers, etc. to achieve the same functions. Therefore, the system and its various devices, modules, and units provided by the present invention can be regarded as a hardware component, and the devices, modules, and units included therein for implementing various functions can also be regarded as the structures within the hardware component; the devices, modules, and units for implementing various functions can also be regarded as both software modules for implementing the method and the structures within the hardware component.

[0092] The specific embodiments of the present invention have been described above. 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 do not affect the essence of the present invention. Without conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A static closed-loop verification method for a Mars exploration separation strategy, characterized in that, It includes the following steps: Step S1: Verify the separation process of the orbiter and the lander; Step S2: Perform static equivalent separation of the orbiter and the lander according to the verified separation process; The said Step S1 includes the following steps: First predetermined time setting step: Power on the orbiter and the lander, simulate the state setting at the first predetermined time before the separation of the orbiter and the lander, and synchronize and set the on-board time of the orbiter and the lander to the first predetermined time moment before separation; Uploading step: Simulate the work process at the first predetermined time before the separation of the orbiter and the lander, upload a set of delay instructions scheduled to be executed in the separation process to the orbiter and the lander, and upload separation control strategy parameters to the orbiter; Confirmation step: Activate the data download function of the orbiter and the lander, confirm that the set of delay instructions and the separation control strategy parameters received by the orbiter are correct, and confirm that the set of delay instructions received by the lander is correct; Second predetermined time time jump step: Jump the on-board time of the orbiter and the lander to the second predetermined time moment before separation; Pre-separation monitoring step: Monitor the operating states of the orbiter and the lander according to the working time sequence of the on-orbit flight program before separation, confirm that the orbiter and the lander correctly execute the pre-separation delay instructions at the predetermined moment, and confirm that the orbiter executes the pre-separation orbit descent attitude maneuver, orbit descent orbit control, and separation attitude maneuver according to the predetermined working time sequence and separation control strategy; Separation state setting step: When the on-board time of the orbiter and the lander runs to the separation moment, send a separation signal; Program-controlled task process triggering step: Confirm that both the orbiter and the lander detect the separation signal and correctly trigger the program-controlled task process after separation; Post-separation monitoring step: Monitor the operating states of the orbiter and the lander according to the working time sequence of the on-orbit flight program after separation, confirm that the lander correctly executes the program-controlled task process according to the predetermined working time sequence, and confirm that the orbiter correctly executes the post-separation delay instructions at the predetermined moment, and executes the post-separation orbit ascent attitude maneuver, orbit ascent orbit control, and post-ascent attitude recovery according to the predetermined working time sequence and separation control strategy parameters; Verification completion step: The separation strategy process of the orbiter and the lander is verified, the states of the orbiter and the lander are restored, and power-off is implemented.

2. The static closed-loop verification method for the Mars exploration separation strategy according to claim 1, wherein This method further includes an internal power supply setting step; In the said first predetermined time setting step, after powering on the orbiter, the orbiter supplies power to the lander; In the said internal power supply setting step, according to the working time sequence of the on-orbit flight program before separation, turn on the battery discharge switch of the lander and turn off the power supply switch of the orbiter to the lander, and set the lander to the internal battery power supply state.

3. The static closed-loop verification method for the Mars exploration separation strategy according to claim 1, characterized in that This method further includes a troubleshooting step: If there is a discrepancy between the operating process and the predetermined working time sequence in the separation process of the orbiter and the lander, after troubleshooting, restore the states of the orbiter and the lander and implement power-off, and restart the verification of the separation process of the orbiter and the lander from the first predetermined time setting step.

4. The static closed-loop verification method for the Mars exploration separation strategy according to claim 1, characterized in that In the said uploading step, the set of delay instructions of the lander is forwarded to the lander by the orbiter.

5. A static closed-loop verification system for a Mars exploration separation strategy, characterized in that, It includes the following modules: Module M1: Verify the separation process of the orbiter and the lander; Module M2: Perform static equivalent separation on the orbiter and the lander according to the verified separation process; The module M1 includes the following modules: First predetermined time setting module: Power on the orbiter and the lander, simulate the state setting at the first predetermined time before the separation of the orbiter and the lander, and synchronize and set the on-board time of the orbiter and the lander to the moment of the first predetermined time before separation; Upload module: Simulate the working process at the first predetermined time before the separation of the orbiter and the lander, upload the delay instruction set scheduled to be executed in the separation process to the orbiter and the lander, and upload the separation control strategy parameters to the orbiter; Confirmation module: Activate the data download function of the orbiter and the lander, confirm that the delay instruction set and the separation control strategy parameters received by the orbiter are correct, and confirm that the delay instruction set received by the lander is correct; Second predetermined time jump module: Jump the on-board time of the orbiter and the lander to the moment of the second predetermined time before separation; Pre-separation monitoring module: Monitor the operating states of the orbiter and the lander according to the working time sequence of the on-orbit flight program before separation, confirm that the orbiter and the lander correctly execute the delay instructions before separation at the predetermined moment, and confirm that the orbiter executes the de-orbit attitude maneuver, de-orbit trajectory control, and separation attitude maneuver before separation according to the predetermined working time sequence and separation control strategy; Separation state setting module: When the on-board time of the orbiter and the lander runs to the separation moment, send a separation signal; Program-controlled task process trigger module: Confirm that both the orbiter and the lander detect the separation signal and correctly trigger the program-controlled task process after separation; Post-separation monitoring module: Monitor the operating states of the orbiter and the lander according to the working time sequence of the on-orbit flight program after separation, confirm that the lander correctly executes the program-controlled task process according to the predetermined working time sequence, and confirm that the orbiter correctly executes the post-separation delay instruction at the predetermined moment, and executes the post-separation orbit-raising attitude maneuver, orbit-raising trajectory control, and post-orbit-raising attitude recovery according to the predetermined working time sequence and separation control strategy parameters; Verification completion module: After the separation strategy process of the orbiter and the lander is verified, restore the states of the orbiter and the lander and implement power-off.

6. The static closed-loop verification system for the Mars exploration separation strategy according to claim 5, characterized in that The system further includes an internal power supply setting module; In the first predetermined time setting module, after the orbiter is powered on, the orbiter powers the lander; In the internal power supply setting module, according to the working time sequence of the on-orbit flight program before separation, turn on the battery discharge switch of the lander and turn off the power supply switch of the orbiter to the lander, and set the lander to the internal battery power supply state.

7. The static closed-loop verification system for the Mars exploration separation strategy according to claim 5, wherein The system further includes a troubleshooting module: If there is a discrepancy between the operating process and the predetermined working time sequence in the separation process of the orbiter and the lander, after troubleshooting, restore the states of the orbiter and the lander and implement power-off, and restart the verification of the separation process of the orbiter and the lander from the first predetermined time setting module.

8. The static closed-loop verification system for the Mars exploration separation strategy according to claim 5, characterized in that In the upload module, the delay instruction set of the lander is forwarded to the lander by the orbiter.

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