Error processing and analyzing method for whole transfer process of lunar large ellipse frozen orbit

Abstract

The invention discloses a method for processing and analyzing errors in the whole transfer process of a lunar large ellipse frozen orbit, and belongs to the technical field of deep space exploration orbits. The method comprises the following steps: setting a whole process error term; the orbit injection errors are added to orbit injection point orbit parameters, independent targeting calculation is conducted on the measured orbit errors of all the stages, and orbit control speed increments of the corresponding stages under the measured orbit errors are obtained; analyzing the influence of the whole-process error on the total speed increment and judging whether the detector can enter a task track or not; analyzing the influence of the whole-process error on the measurement and control conditions and judging whether the measurement and control conditions are met or not; and analyzing the influence of the whole-process error on the orbit safety and the autonomous deorbit at the end of the task, and judging whether the detector is controlled to deorbit, if so, ending the whole-process error analysis, and if not, reserving the speed increment to perform orbit correction. Compared with single analysis of the errors in each stage, the method is more systematic and closer to the real situation, and the analysis result can provide reference for the overall design and the design of each subsystem.

Description

Technical Field

[0001] The present invention belongs to the technical field of deep space exploration orbits, and in particular relates to a method for processing and analyzing errors in the entire process of transferring a large elliptical frozen orbit around the moon. Background Art

[0002] The communication conditions between the ground and the far side and the poles of the moon are poor. Relying solely on the tracking and control stations built on the earth can no longer meet the complex needs of future lunar exploration. Therefore, it is necessary to establish a lunar communication, navigation and remote sensing system to provide relay communication, navigation and remote sensing services for users on the moon and in the lunar circumlunar area. The large elliptical frozen orbit around the moon has become the core of the current lunar polar exploration and the establishment of lunar scientific research stations.

[0003] However, there are orbit insertion errors, orbit determination errors and orbit control errors in each stage during the transfer from orbit insertion to the large elliptical frozen orbit around the moon. The existence of errors will affect the total orbit velocity increment, measurement and control conditions, orbit safety, and deorbit at the end of the mission. In addition, a unified analysis of the errors of the entire transfer process is more systematic and more in line with engineering practice than a single analysis of errors at each stage of the transfer. Summary of the invention

[0004] In order to solve the above technical problems, the present invention provides an error processing and analysis method for the whole process of transfer of a large elliptical frozen orbit around the moon. The error of the whole process is considered in a unified way, which conforms to engineering practice. Compared with a single analysis of the errors of each stage, it is more systematic and closer to the actual situation. The analysis results can provide a reference for the overall design and each subsystem design.

[0005] To achieve the above purpose, the technical solution adopted by the present invention is as follows:

[0006] A method for error processing and analysis of the entire process of transferring a large elliptical frozen orbit around the moon includes the following steps:

[0007] Step 1: Set the error terms for the entire process, including the orbit insertion error, the orbit determination error at each stage, and the orbit control error. The stages include mid-course correction, capture braking, far-moon maneuvering, and near-moon orbit reduction.

[0008] The whole process of orbit transfer refers to the whole process of transfer maneuvers from orbit entry to large elliptical frozen orbit, including orbit entry, correction in the Earth-Moon transfer phase, lunar braking capture to enter the large elliptical orbit around the moon, maneuvering near the apogee to enter the large elliptical frozen orbit, maneuvering at the perigee to enter the large elliptical frozen orbit of a given target period, and maneuvering at the apogee to lower the perigee altitude to achieve controlled deorbit at the end of the mission; among which, the orbit control constraints are that there must be measurement and control support for at least 3 hours before each orbit control, no collision with the moon within 1 year in the lunar orbit, and autonomous deorbit after the mission is completed;

[0009] It can be seen from the maneuvers in each stage that the error terms of the whole process mainly include orbit insertion error, mid-course correction orbit determination error and orbit control error, capture braking orbit determination error and orbit control error, far-moon maneuver orbit determination error and orbit control error, and near-moon orbit determination error and orbit control error.

[0010] Step 2: Process the error term of the whole process, including adding the orbit entry error to the orbit parameter of the orbit entry point, performing separate target calculation on the orbit determination error of each stage, and obtaining the orbit control speed increment of the corresponding stage under the orbit determination error as the orbit control speed increment of the nominal orbit at the corresponding stage; calculating the orbit control error to generate speed increment error data;

[0011] Step 2.1: Processing orbit determination error; Based on the orbit determination accuracy given by the measurement and control system, randomly generate orbit determination error data for each stage according to the 3σ principle of normal distribution. With the recursion of the nominal orbit, take out the position and velocity RTN vector of the orbit control point of the corresponding stage and add the orbit determination error of the corresponding stage to perform a separate target calculation, and obtain the orbit control velocity increment of the corresponding stage under the orbit determination error. Take it as the orbit control velocity increment of the nominal orbit in the corresponding stage, and consider the orbit control error. Recurse to the next stage, repeat the separate target calculation and orbit recursion operation until entering the mission orbit;

[0012] Step 2.2: Process orbit control errors; set the azimuth and altitude error ranges of the orbit control thrust direction, randomly generate azimuth and altitude error data according to the normal distribution 3σ principle, and add them to the azimuth and altitude of the thrust direction at each stage of the nominal orbit; the orbit control velocity increment error is calculated as 0.15+0.01443dv, where dv is the orbit control velocity increment at each stage, and randomly generate velocity increment error data according to the normal distribution 1σ principle, and add them to the orbit control velocity increment value at each stage of the nominal orbit.

[0013] Step 3: Analyze the impact of the whole process error on the total velocity increment and determine whether the probe can enter the mission orbit. If so, proceed to step 4. If not, it indicates that the probe may not be able to enter the mission orbit when the error effect is considered. At this time, it is necessary to adjust the relevant characteristic parameters of the transfer orbit design.

[0014] Step 3.1: Considering the error term of the whole process, the Monte Carlo shooting simulation is performed on the whole process of orbit transfer to obtain the mean value, mean square error and maximum value of total speed increment consumption of the error term of the whole process on the total speed increment;

[0015] Step 3.2: Compare whether the maximum total velocity increment consumption is less than the velocity increment budget. If so, the probe can enter the mission orbit.

[0016] Step 4: Analyze the impact of the whole process error on the measurement and control conditions and determine whether the measurement and control conditions are met. If so, proceed to step 5. If not, the measurement and control conditions can be guaranteed by increasing or decreasing the capture period to change the orbit control time in the subsequent stage;

[0017] Since the mid-course correction and lunar capture orbit control time are fixed, the measurement and control conditions are not affected by errors. During the apogee maneuver, the apogee is near the apogee, and during the perigee is near the perigee, the measurement and control conditions will be affected by errors.

[0018] Taking the measurement and control conditions at each stage without considering the error as the reference value, the Monte Carlo target shooting simulation is calculated under the conditions of considering the orbit insertion error, orbit determination error and orbit control error, and the mean and mean square error of the measurement and control conditions at each stage compared with the reference value are obtained;

[0019] The maximum value of the measurement and control time advance of each stage is calculated according to the three-fold mean square error to determine whether the measurement and control conditions are met. If the measurement and control support time before orbit control in each stage is greater than 3 hours, the measurement and control conditions are met; otherwise, the measurement and control conditions are not met.

[0020] Step 5: Analyze the impact of the full-process error on orbital safety and autonomous deorbiting at the end of the mission, and determine whether the orbit has sufficient safety margin during the mission, and whether the probe can be controlled to deorbit. If so, the full-process error analysis is completed. If not, reserve a velocity increment for orbit correction to ensure that there is sufficient safety margin during the mission and that the conditions for deorbiting are met;

[0021] If the minimum perigee height in the lunar circumlunar phase is lower than 50 km, it is considered that there is a risk of impacting the moon. Taking into account the error of the entire process, the Monte Carlo target shooting simulation is used to obtain the proportion of the minimum perigee height in the lunar circumlunar phase lower than 50 km, thereby judging the risk of impacting the moon. If the proportion of the minimum perigee height of the mission orbit lower than 50 km is higher than the reference value, it indicates that there is a risk of impacting the moon. If the proportion of the minimum perigee height of the mission orbit lower than 50 km is lower than the reference value, it indicates that the mission orbit is safe.

[0022] If the minimum perigee height at the end of the life is less than 200km, it is considered to have the ability to deorbit in a controlled manner. Under the full process error, the Monte Carlo target simulation is used to obtain the proportion of the minimum perigee height at the end of the life below 200km to determine whether it can be deorbited in a controlled manner. If the proportion of the minimum perigee height at the end of the mission is higher than 200km is higher than the reference value, it indicates that it cannot be deorbited successfully. If the proportion of the minimum perigee height at the end of the mission is higher than 200km is lower than the reference value, it indicates that it can deorbit autonomously.

[0023] The beneficial effects of the present invention are:

[0024] The present invention includes a method for analyzing the errors of the entire process from orbit insertion to a large elliptical frozen orbit around the moon, and analyzes the impact of the errors: first, the entire process error is defined to include errors in each stage of the orbit, mainly including orbit insertion errors, orbit determination errors in each stage, and orbit control errors; then, the impact of the entire process error is analyzed, and the entire process error mainly affects the total velocity increment, measurement and control conditions, orbit safety, and deorbit. The present application solves the problem of error analysis of the entire process of transferring to a large elliptical frozen orbit around the moon, uniformly considers the errors of the entire process, conforms to engineering practice, and is more systematic and closer to the actual situation than a single analysis of the errors of each stage. The analysis results can provide a reference for the overall and subsystem designs. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] Figure 1 This is a schematic diagram of the error processing and analysis method for the entire process of transferring a large elliptical frozen orbit around the moon according to the present invention;

[0026] Figure 2 This is a curve of the telemetry and control conditions for the far-moon maneuvering in the embodiment of the present invention without considering the error;

[0027] Figure 3 is a photographed variation curve of the perigee height of an embodiment of the present invention;

[0028] Figure 4 This is a curve showing the change in height from the lunar surface before and after applying a pulse in an embodiment of the present invention. DETAILED DESCRIPTION

[0029] The present invention will be further described below in conjunction with the accompanying drawings and embodiments.

[0030] like Figure 1 As shown, it is a principle diagram of an error processing and analysis method for the whole process of transferring a large elliptical frozen orbit around the moon disclosed in the present invention. The specific implementation steps of the method are as follows:

[0031] Step 1: Set the error terms of the whole process, including the orbit insertion error, the orbit determination error of each stage and the orbit control error. The stages include mid-course correction, capture braking, far-moon maneuvering, and near-moon orbit reduction.

[0032] The whole process of orbit transfer refers to the whole process of transfer maneuvers from orbit entry to large elliptical frozen orbit, including orbit entry, correction in the Earth-Moon transfer phase, lunar braking capture to enter the large elliptical orbit around the moon, maneuvering near the apogee to enter the large elliptical frozen orbit, maneuvering at the perigee to enter the large elliptical frozen orbit of a given target period, and maneuvering at the apogee to lower the perigee altitude to achieve controlled deorbit at the end of the mission; among which, the orbit control constraints are that there must be measurement and control support for at least 3 hours before each orbit control, no collision with the moon within 1 year in the lunar orbit, and autonomous deorbit after the mission is completed;

[0033] It can be seen from the maneuvers in each stage that the error terms of the whole process mainly include orbit insertion error, mid-course correction orbit determination error and orbit control error, capture braking orbit determination error and orbit control error, far-moon maneuver orbit determination error and orbit control error, and near-moon orbit determination error and orbit control error.

[0034] Step 2: Process the error term of the whole process, including adding the orbit entry error to the orbit parameter of the orbit entry point, performing separate target calculation on the orbit determination error of each stage, and obtaining the orbit control speed increment of the corresponding stage under the orbit determination error as the orbit control speed increment of the nominal orbit at the corresponding stage; calculating the orbit control error to generate speed increment error data;

[0035] Step 2.1: Processing orbit determination error; Based on the orbit determination accuracy given by the measurement and control system, randomly generate orbit determination error data for each stage according to the 3σ principle of normal distribution. With the recursion of the nominal orbit, take out the position and velocity RTN vector of the orbit control point of the corresponding stage and add the orbit determination error of the corresponding stage to perform a separate target calculation, and obtain the orbit control velocity increment of the corresponding stage under the orbit determination error. Take it as the orbit control velocity increment of the nominal orbit in the corresponding stage, and consider the orbit control error. Recurse to the next stage, repeat the separate target calculation and orbit recursion operation until entering the mission orbit;

[0036] Step 2.2: Process orbit control errors; set the error range of azimuth and altitude angles of the thrust direction, randomly generate azimuth and altitude error data according to the 3σ principle of normal distribution, and add them to the azimuth and altitude angles of the thrust direction at each stage of the nominal orbit; the orbit control velocity increment error is calculated as 0.15+0.01443dv, where dv is the orbit control velocity increment at each stage, and randomly generate velocity increment error data according to the 1σ principle of normal distribution, and add them to the orbit control velocity increment value at each stage of the nominal orbit.

[0037] Step 3: Analyze the impact of the whole process error on the total velocity increment and determine whether the probe can enter the mission orbit. If so, proceed to step 4. If not, it indicates that the probe may not be able to enter the mission orbit when the error effect is considered. At this time, it is necessary to adjust the relevant characteristic parameters of the transfer orbit design.

[0038] Step 3.1: Considering the error term of the whole process, the Monte Carlo shooting simulation is performed on the whole process of orbit transfer to obtain the mean value, mean square error and maximum value of total speed increment consumption of the error term of the whole process on the total speed increment;

[0039] Step 3.2: Compare whether the maximum total velocity increment consumption is less than the velocity increment budget. If so, the probe can enter the mission orbit.

[0040] Given a set of orbital parameters, the Monte Carlo target shooting simulation was performed 500 times considering the full process error. The average impact of the full process error on the total velocity increment is 398.702 m / s, the mean square error is 19.4 m / s, and the maximum total velocity increment consumption is 482.5 m / s, which is less than the velocity increment budget. Therefore, considering the full process error, the probe can enter the mission orbit.

[0041] Step 4: Analyze the impact of the whole process error on the measurement and control conditions and determine whether the measurement and control conditions are met. If so, proceed to step 5. If not, the measurement and control conditions can be guaranteed by increasing or decreasing the capture period to change the orbit control time in the subsequent stage;

[0042] Since the mid-course correction and lunar capture orbit control time are fixed, the measurement and control conditions are not affected by errors. During the apogee maneuver, the apogee is near the apogee, and during the perigee is near the perigee, the measurement and control conditions will be affected by errors.

[0043] Taking the measurement and control conditions at each stage without considering the error as the reference value, the Monte Carlo target shooting simulation is calculated under the conditions of considering the orbit insertion error, orbit determination error and orbit control error, and the mean and mean square error of the measurement and control conditions at each stage compared with the reference value are obtained;

[0044] The maximum value of the measurement and control time advance of each stage is calculated according to the three-fold mean square error to determine whether the measurement and control conditions are met. If the measurement and control support time before orbit control in each stage is greater than 3 hours, the measurement and control conditions are met; otherwise, the measurement and control conditions are not met.

[0045] Without considering the error, take the measurement and control conditions of the far-moon maneuver orbit control time as an example. Figure 2 As shown, the horizontal axis is the deep space station, and the vertical axis is time, in hours. It can be seen that the far-moon maneuver orbital control moment has a large delay time margin to meet the measurement and control constraints. At the same time, the simulation results show that the measurement and control conditions can basically be met when the orbital control moment is delayed due to errors; and the advance of the orbital control moment is likely to cause the measurement and control support time to be less than 3 hours, and it is highly sensitive to the measurement and control conditions. The maneuvers in other stages are similar. For the convenience of analysis, only the measurement and control time deviation caused by the advance of the orbital control time can be considered in the future to reduce the complexity of the analysis.

[0046] Step 5: Analyze the impact of the full-process error on orbital safety and autonomous deorbiting at the end of the mission, and determine whether the orbit has sufficient safety margin during the mission, and whether the probe can be controlled to deorbit. If so, the full-process error analysis is completed. If not, reserve a velocity increment for orbit correction to ensure that there is sufficient safety margin during the mission and that the conditions for deorbiting are met;

[0047] If the minimum perigee height in the lunar circumlunar phase is lower than 50 km, it is considered that there is a risk of impacting the moon. Taking into account the error of the entire process, the Monte Carlo target shooting simulation is used to obtain the proportion of the minimum perigee height in the lunar circumlunar phase lower than 50 km, thereby judging the risk of impacting the moon. If the proportion of the minimum perigee height of the mission orbit lower than 50 km is higher than the reference value, it indicates that there is a risk of impacting the moon. If the proportion of the minimum perigee height of the mission orbit lower than 50 km is lower than the reference value, it indicates that the mission orbit is safe.

[0048] If the minimum perigee height at the end of the life is less than 200km, it is considered to have the ability to deorbit in a controlled manner. Under the full process error, the Monte Carlo target simulation is used to obtain the proportion of the minimum perigee height at the end of the life below 200km to determine whether it can be deorbited in a controlled manner. If the proportion of the minimum perigee height at the end of the mission is higher than 200km is higher than the reference value, it indicates that it cannot be deorbited successfully. If the proportion of the minimum perigee height at the end of the mission is higher than 200km is lower than the reference value, it indicates that it can deorbit autonomously.

[0049] At the end of its life, a controlled impact with the moon can avoid the generation of lunar orbital debris and reduce the threat to future lunar exploration spacecraft. Considering the situation at the end of its life, when the probe has limited fuel, the deorbit plan under the condition of insufficient fuel is studied.

[0050] The lunar satellite is greatly affected by perturbations. The law of the influence of perturbations on the height of the perigee is analyzed. The positive or negative value of the height of the perigee determines whether the probe can hit the moon. Figure 3 As shown, it can be seen that the minimum value of the perigee height in a short period of time is constantly changing. By utilizing the minimum value of the perigee height and selecting the appropriate time to apply the pulse, the lunar impact orbit control gain can be maximized. The deorbiting is achieved by applying a velocity increment at the apogee to reduce the perigee height to below the lunar surface. Combined with the minimum value of the perigee height, the deorbiting scheme is specifically designed as follows: after the mission is completed, the time of the minimum value of the perigee height within the deorbiting time limit is selected, and the orbit is moved to the apogee of the current circle, and a pulse is applied in the opposite direction of the velocity direction to achieve a lunar impact near the perigee.

[0051] In the case of controlled deorbit, the height change curve from the lunar surface before and after the pulse is applied is as follows: Figure 4 As shown in the figure, when the horizontal axis is close to 495 days, the perigee height is at a minimum value of "115.3km", and a pulse is applied from this minimum point to the apogee to achieve a moon impact near the perigee.

[0052] The specific embodiments described above further illustrate the objectives, technical solutions and beneficial effects of the present invention in detail. It should be understood that the above description is only a specific embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.

Claims

1. A method for error processing and analysis of the entire process of transfer of a large elliptical frozen orbit around the moon, characterized in that: The following steps are involved: Step 1: Set the error terms of the whole process, including the orbit insertion error, the orbit determination error of each stage and the orbit control error. The stages include mid-course correction, capture braking, far-moon maneuvering, and near-moon orbit reduction. Step 2: Process the error term of the whole process, including adding the orbit entry error to the orbit parameter of the orbit entry point, performing separate target calculations on the orbit determination errors of each stage, and obtaining the orbit control velocity increment of the corresponding stage under the orbit determination error as the orbit control velocity increment of the nominal orbit at the corresponding stage; Calculate the orbit control error to generate velocity increment error data; Step 3: Analyze the impact of the whole process error on the total velocity increment and determine whether the probe can enter the mission orbit. If so, proceed to step 4. If not, adjust the relevant characteristic parameters of the transfer orbit design. Step 4: Analyze the impact of the whole process error on the measurement and control conditions and determine whether the measurement and control conditions are met. If so, proceed to step 5. If not, change the orbit control time in the subsequent stage by increasing or decreasing the capture period to ensure the measurement and control conditions. Step 5: Analyze the impact of the full-process error on orbital safety and autonomous deorbit at the end of the mission, and determine whether the probe will be deorbited in a controlled manner. If so, the full-process error analysis is completed. If not, reserve a velocity increment for orbit correction.

2. The error processing and analysis method for the whole process of the transfer of a large elliptical frozen orbit around the moon according to claim 1 is characterized in that: The whole process in step 1 represents the whole process of transfer maneuver from orbit entry to large elliptical frozen orbit, including orbit entry, correction in the Earth-Moon transfer phase, lunar braking capture to enter the large elliptical orbit around the moon, maneuver near the apogee to enter the large elliptical frozen orbit, maneuver at the perigee to enter the large elliptical frozen orbit of a given target period, and maneuver at the apogee to lower the perigee altitude to achieve controlled deorbit at the end of the mission.

3. The error processing and analysis method for the whole process of the transfer of a large elliptical frozen orbit around the moon according to claim 2 is characterized in that: The orbit control constraints are that there must be measurement and control support for at least 3 hours before each orbit control, no collision with the moon within one year in the lunar orbit, and autonomous deorbit after the mission is completed.

4. The error processing and analysis method for the whole process of the transfer of a large elliptical frozen orbit around the moon according to claim 1 is characterized in that: The step 2 includes processing the orbit determination error and processing the orbit control error.

5. The error processing and analysis method for the whole process of the transfer of a large elliptical frozen orbit around the moon according to claim 4 is characterized in that: The processing of the orbit determination error includes the orbit determination accuracy given by the measurement and control system, randomly generating orbit determination error data for each stage according to the 3σ principle of normal distribution, and recursively taking the orbit control point position and speed RTN vector of the corresponding stage along with the nominal orbit, and adding the orbit determination error of the corresponding stage to perform a separate target calculation, so as to obtain the orbit control speed increment of the corresponding stage under the orbit determination error, and use it as the orbit control speed increment of the nominal orbit in the corresponding stage, and taking the orbit control error into consideration, the orbit is recursively transferred to the next stage, and the separate target calculation and orbit recursion operation are repeated until entering the mission orbit.

6. The error processing and analysis method for the whole process of the transfer of a large elliptical frozen orbit around the moon according to claim 4 is characterized in that: The processing of orbit control errors includes setting the error range of azimuth and altitude angle of orbit control thrust direction, randomly generating azimuth and altitude angle error data according to the normal distribution 3σ principle, and adding them to the azimuth and altitude angle of thrust direction at each stage of the nominal orbit; the orbit control speed increment error is calculated as 0.15+0.01443dv, where dv is the orbit control speed increment at each stage, and randomly generating speed increment error data according to the normal distribution 1σ principle, and adding them to the orbit control speed increment value at each stage of the nominal orbit.

7. The error processing and analysis method for the whole process of transfer of a large elliptical frozen orbit around the moon according to claim 1 is characterized in that: The step 3 comprises: Step 3.1: Considering the error term of the whole process, the Monte Carlo shooting simulation is performed on the whole process of orbit transfer to obtain the mean value, mean square error and maximum value of total speed increment consumption of the error term of the whole process on the total speed increment; Step 3.2: Compare whether the maximum total velocity increment consumption is less than the velocity increment budget. If so, the probe can enter the mission orbit.

8. The error processing and analysis method for the whole process of transfer of a large elliptical frozen orbit around the moon according to claim 1 is characterized in that: The step 4 comprises: Taking the measurement and control conditions at each stage without considering the error as the reference value, the Monte Carlo target shooting simulation is calculated under the conditions of considering the orbit insertion error, orbit determination error and orbit control error, and the mean and mean square error of the measurement and control conditions at each stage compared with the reference value are obtained; The maximum value of the measurement and control time advance of each stage is calculated according to the three-fold mean square error to determine whether the measurement and control conditions are met. If the measurement and control support time before orbit control in each stage is greater than 3 hours, the measurement and control conditions are met; otherwise, the measurement and control conditions are not met.

9. The error processing and analysis method for the whole process of the transfer of a large elliptical frozen orbit around the moon according to claim 1 is characterized in that: The step 5 comprises: If the minimum perigee height in the lunar circumlunar phase is lower than 50 km, it is considered that there is a risk of impacting the moon. Taking into account the error of the entire process, the Monte Carlo target shooting simulation is used to obtain the proportion of the minimum perigee height in the lunar circumlunar phase lower than 50 km, thereby judging the risk of impacting the moon. If the proportion of the minimum perigee height of the mission orbit lower than 50 km is higher than the reference value, it indicates that there is a risk of impacting the moon. If the proportion of the minimum perigee height of the mission orbit lower than 50 km is lower than the reference value, it indicates that the mission orbit is safe.

10. The error processing and analysis method for the whole process of transfer of a large elliptical frozen orbit around the moon according to claim 1 is characterized in that: The step 5 also includes: If the minimum perigee height at the end of the life is lower than 200 km, it is considered that there is the ability to deorbit in a controlled manner. Under the full process error, the Monte Carlo target shooting simulation is used to obtain the proportion of the minimum perigee height at the end of the life below 200 km, and judge whether it can be deorbited in a controlled manner. If the proportion of the minimum perigee height above 200 km at the end of the mission is higher than the reference value, it indicates that it cannot deorbit successfully. If the proportion of the minimum perigee height above 200 km at the end of the mission is lower than the reference value, it indicates that it can deorbit autonomously.

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