A method for processing and analyzing whole-process errors of a lunar circumlunar highly elliptical frozen orbit transfer

Abstract

The application discloses a method for processing and analyzing whole-process errors of a lunar orbit transfer of a large-ellipse frozen orbit, and belongs to the technical field of deep space exploration orbits. The method comprises the following steps: setting whole-process error terms; adding an orbit entry error to an orbit parameter at an orbit entry point, separately calculating measured orbit errors of each stage to obtain orbit control velocity increments of the corresponding stages under the measured orbit errors; analyzing the influence of the whole-process errors on total velocity increments and judging whether a probe can enter a mission orbit; analyzing the influence of the whole-process errors on measurement and control conditions and judging whether the measurement and control conditions are met; analyzing the influence of the whole-process errors on orbit safety and autonomous orbit departure at the end of a mission and judging whether the probe will be controlled to depart from the orbit, if yes, the whole-process error analysis is ended, and if no, a reserved velocity increment is used for orbit correction. Compared with single analysis of errors of each stage, the application is more systematic and closer to a real situation, and the analysis result can provide a reference for overall and each subsystem design.

Description

Technical Field

[0001] This invention belongs to the field of deep space exploration orbit technology, specifically relating to a method for error processing and analysis of the entire process of transferring a large elliptical frozen orbit around the moon. Background Technology

[0002] Communication conditions between Earth and the far side of the Moon and the polar regions are poor, and relying solely on tracking and control stations on Earth is no longer sufficient to meet the complex needs of future lunar exploration. Therefore, a lunar communication, navigation, and remote sensing system is needed to provide relay communication, navigation, and remote sensing services to users on the lunar surface and in lunar orbit. A highly elliptical frozen orbit around the Moon has become the core of current lunar polar exploration and the establishment of lunar research stations.

[0003] However, the transfer from the initial orbit to the lunar elliptical frozen orbit involves orbit insertion errors, as well as trajectory measurement and control errors at various stages. These errors will affect the total orbital velocity increment, measurement and control conditions, orbital safety, and deorbiting at the end of the mission. Furthermore, a unified analysis of the errors throughout the entire transfer process is more systematic and consistent with engineering practice than analyzing errors at individual stages of the transfer. Summary of the Invention

[0004] To address the aforementioned technical problems, this invention provides a method for error processing and analysis throughout the entire process of transferring a large elliptical frozen orbit around the moon. This method considers all errors in a unified manner, conforms to engineering practice, and is more systematic and closer to the actual situation than analyzing errors at each stage in a single analysis. The analysis results can provide a reference for the overall design and the design of each subsystem.

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

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

[0007] Step 1: Set the error items for the whole process, including the orbit insertion error, the orbit measurement error and the orbit control error at each stage. The stages include mid-course correction, capture braking, far-lunar maneuver, and near-lunar orbit descent.

[0008] The entire orbit transfer process refers to the maneuvering from orbit insertion to the highly elliptical frozen orbit, including orbit insertion, Earth-Moon transfer phase correction, lunar braking capture into the highly elliptical lunar orbit, maneuvering near the apogee into the highly elliptical frozen orbit, maneuvering at the perigee into the highly elliptical frozen orbit with a given target period, and maneuvering at the apogee to reduce the perigee altitude at the end of the mission to achieve controlled deorbiting; among them, the orbit control constraints are that there is telemetry 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 deorbiting after the mission ends.

[0009] From the maneuver of each phase, the whole-process error items mainly include the orbit insertion error, the mid-course correction orbit determination error and orbit control error, the capture braking orbit determination error and orbit control error, the far-month maneuver orbit determination error and orbit control error, and the near-month deorbit orbit determination error and orbit control error.

[0010] Step 2: processing the whole-process error items, including adding the orbit insertion error to the orbit parameters at the orbit insertion point, separately targeting the orbit determination error of each phase to obtain the orbit control velocity increment of the corresponding phase under the orbit determination error as the orbit control velocity increment of the nominal orbit at the corresponding phase, and calculating the orbit control error to generate the velocity increment magnitude error data;

[0011] Step 2.1: processing the orbit determination error; according to the orbit determination accuracy given by the TT&C system, the orbit determination error data of each phase is randomly generated according to the normal distribution 3σ principle, the corresponding phase orbit control point position and velocity RTN vector are taken out and added to the corresponding phase orbit determination error for separate targeting calculation, the orbit control velocity increment of the corresponding phase under the orbit determination error is obtained, which is taken as the orbit control velocity increment of the nominal orbit at the corresponding phase, and the orbit control error is considered, the orbit is recursively propagated to the next phase, and the separate targeting calculation and orbit recursive propagation operations are repeated until the mission orbit is entered;

[0012] Step 2.2: processing the orbit control error; the azimuth and elevation angle error range of the orbit control thrust direction is set, the azimuth and elevation angle error data are randomly generated according to the normal distribution 3σ principle and added to the azimuth and elevation angle of the orbit control thrust direction of the nominal orbit at each phase; the orbit control velocity increment magnitude error is calculated as 0.15+0.01443dv, where dv is the orbit control velocity increment at each phase, the velocity increment magnitude error data is randomly generated according to the normal distribution 1σ principle and added to the orbit control velocity increment value at each phase of the nominal orbit.

[0013] Step 3: analyzing the influence of the whole-process error on the total velocity increment and determining whether the probe can enter the mission orbit, if yes, entering step 4, if not, it indicates that the probe may not be able to enter the mission orbit when considering the error effect, and at this time, the related characteristic parameters of the transfer orbit design need to be adjusted;

[0014] Step 3.1: considering the whole-process error items, performing Monte Carlo targeting simulation on the whole-process of orbit transfer to obtain the mean value, mean square deviation and maximum value of the total velocity increment consumed by the influence of the whole-process error items on the total velocity increment;

[0015] Step 3.2: comparing whether the maximum value of the total velocity increment consumed is less than the velocity increment budget, if yes, the probe can enter the mission orbit.

[0016] Step 4: analyze the influence of the whole-process error on the measurement and control condition and judge whether the measurement and control condition is satisfied, if yes, enter step 5, if not, the measurement and control condition can be ensured by increasing or decreasing the capture period to change the orbit control time in the subsequent stage;

[0017] Since the midway correction and the lunar capture orbit control time are fixed, the measurement and control condition is not affected by the error, and the measurement and control condition is affected by the error when the far-moon maneuver is near the far-moon point and the near-moon descent is near the near-moon point;

[0018] Taking the measurement and control condition of each stage without considering the error as a reference value, the Monte Carlo shooting simulation is carried out under the condition of considering the entry orbit error, the determination orbit error and the orbit control error, and the mean value and the mean square deviation of the measurement and control condition of each stage compared with the reference value are obtained;

[0019] The maximum value of the measurement and control time of each stage is calculated according to 3 times the mean square deviation, and it is judged whether the measurement and control condition is satisfied, if the measurement and control support time before the orbit control of each stage is greater than 3 hours, the measurement and control condition is satisfied, otherwise, the measurement and control condition is not satisfied.

[0020] Step 5: analyze the influence of the whole-process error on the orbit safety and the autonomous de-orbiting at the end of the mission, and judge whether the orbit has sufficient safety margin during the mission and whether the probe can be controlled to de-orbit, if yes, the whole-process error analysis is ended, if not, a velocity increment is reserved for orbit correction to ensure that there is sufficient safety margin during the mission and the de-orbiting condition is met;

[0021] It is considered that there is a risk of moon collision when the minimum near-moon point height in the circumlunar stage is lower than 50km, the Monte Carlo shooting simulation is carried out under the whole-process error, the proportion of the minimum near-moon point height lower than 50km in the circumlunar stage is obtained, and the risk of moon collision is judged, if the proportion of the minimum near-moon point height lower than 50km in the mission orbit is higher than the reference value, it is indicated that there is a risk of moon collision, if the proportion of the minimum near-moon point height lower than 50km in the mission orbit is lower than the reference value, it is indicated that the mission orbit has safety.

[0022] It is considered that there is a controlled de-orbiting ability when the minimum near-moon point height at the end of the life is lower than 200km, the Monte Carlo shooting simulation is carried out under the whole-process error, the proportion of the minimum near-moon point height lower than 200km at the end of the life is obtained, and it is judged whether the controlled de-orbiting can be achieved, if the proportion of the minimum near-moon point height higher than 200km at the end of the mission is higher than the reference value, it is indicated that the de-orbiting cannot be successfully achieved, if the proportion of the minimum near-moon point height higher than 200km at the end of the mission is lower than the reference value, it is indicated that the autonomous de-orbiting can be achieved.

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

[0024] This invention includes a method for analyzing the entire process of transferring from the initial orbit to a lunar elliptical frozen orbit, and analyzes the impact of these errors. First, the entire process error is defined as including errors at each stage of the orbit, primarily including the initial orbit error, the measurement and control errors at each stage, and the orbit control errors. Then, the impact of the entire process error is analyzed, showing that it mainly affects the total velocity increment, measurement and control conditions, orbit safety, and de-orbiting. This application solves the problem of analyzing the entire process error during the transfer of a lunar elliptical frozen orbit, considering all process errors in a unified manner, which aligns with engineering practice. Compared to analyzing errors at each stage individually, this method is more systematic and closer to reality. The analysis results can provide a reference for the overall design and the design of each subsystem. Attached Figure Description

[0025] Figure 1 This is a schematic diagram illustrating the principle of an error processing and analysis method for the entire process of transferring a large elliptical frozen orbit around the moon, as described in this invention.

[0026] Figure 2 This is the lunar maneuvering measurement and control condition curve under the condition of not considering errors in the embodiments of the present invention;

[0027] Figure 3 This is a curve showing the perturbation variation of the lunar perigee height according to an embodiment of the present invention;

[0028] Figure 4 The curves showing the change in altitude above the lunar surface before and after applying the pulse in this embodiment of the invention. Detailed Implementation

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

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

[0031] Step 1: Set the error items for the whole process, including the orbit insertion error, the orbit measurement error and the orbit control error at each stage. The stages include mid-course correction, capture braking, far-lunar maneuver, and near-lunar orbit descent.

[0032] The entire orbit transfer process refers to the maneuvering from orbit insertion to the highly elliptical frozen orbit, including orbit insertion, Earth-Moon transfer phase correction, lunar braking capture into the highly elliptical lunar orbit, maneuvering near the apogee into the highly elliptical frozen orbit, maneuvering at the perigee into the highly elliptical frozen orbit with a given target period, and maneuvering at the apogee to reduce the perigee altitude at the end of the mission to achieve controlled deorbiting; among them, the orbit control constraints are that there is telemetry 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 deorbiting after the mission ends.

[0033] From the maneuver of each phase, the whole-process error items mainly include the orbit insertion error, the mid-course correction orbit determination error and orbit control error, the capture braking orbit determination error and orbit control error, the far-month maneuver orbit determination error and orbit control error, and the near-month deorbiting orbit determination error and orbit control error.

[0034] Step 2: processing the whole-process error items, including adding the orbit insertion error to the orbit parameters at the orbit insertion point, separately targeting the orbit determination error of each phase to obtain the orbit control velocity increment of the corresponding phase under the orbit determination error as the orbit control velocity increment of the nominal orbit at the corresponding phase, and calculating the orbit control error to generate the velocity increment magnitude error data;

[0035] Step 2.1: processing the orbit determination error; according to the orbit determination accuracy given by the TT&C system, the orbit determination error data of each phase is randomly generated according to the normal distribution 3σ principle, the corresponding phase orbit control point position and velocity RTN vector are taken out and added to the corresponding phase orbit determination error for separate targeting calculation, the orbit control velocity increment of the corresponding phase under the orbit determination error is obtained, which is taken as the orbit control velocity increment of the nominal orbit at the corresponding phase, and the orbit control error is considered, the orbit is recursively propagated to the next phase, and the separate targeting calculation and orbit recursive propagation operations are repeated until the mission orbit is entered;

[0036] Step 2.2: processing the orbit control error; the azimuth and elevation angle error range of the thrust direction is set, the azimuth and elevation angle error data are randomly generated according to the normal distribution 3σ principle, and are added to the azimuth and elevation angle of the thrust direction of the nominal orbit at each phase; the orbit control velocity increment magnitude error calculation is 0.15+0.01443dv, wherein dv is the orbit control velocity increment at each phase, the velocity increment magnitude error data is randomly generated according to the normal distribution 1σ principle, and is added to the orbit control velocity increment value at each phase of the nominal orbit.

[0037] Step 3: analyzing the influence of the whole-process error on the total velocity increment and judging whether the probe can enter the mission orbit, if yes, entering step 4, if not, it is indicated that the probe may not be able to enter the mission orbit when the error effect is considered, and at this time, the related characteristic parameters of the transfer orbit design need to be adjusted;

[0038] Step 3.1: considering the whole-process error items, performing Monte Carlo targeting simulation on the whole-process of the orbit transfer to obtain the mean value, mean square deviation and maximum value of the total velocity increment consumed by the influence of the whole-process error items on the total velocity increment;

[0039] Step 3.2: comparing whether the maximum value of the total velocity increment consumed is less than the velocity increment budget, if yes, the probe can enter the mission orbit.

[0040] Given a set of orbit parameters, consider the full-process error Monte Carlo targeting simulation 500 times. The average total velocity increment is 398.702 m / s, the mean square deviation 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, under the condition of considering the full-process error, the probe can enter the mission orbit.

[0041] Step 4: Analyze the influence of full-process error on the TT&C condition and determine whether the TT&C condition is met. If yes, go to step 5, if not, increase or decrease the capture period to change the subsequent phase orbit control time to ensure the TT&C condition;

[0042] Since the midway correction and lunar capture orbit control time are fixed, the TT&C condition is not affected by the error. The remote lunar maneuver is near the apolune, and the near-lunar descent is at the periselene. The TT&C condition will be affected by the error;

[0043] Taking the TT&C condition of each phase without considering the error as the reference value, calculate the Monte Carlo targeting simulation considering the entry error, orbit determination error and orbit control error, and get the mean value and mean square deviation of the TT&C condition of each phase compared with the reference value;

[0044] According to 3 times the mean square deviation, the maximum value of the TT&C time of each phase is calculated to determine whether the TT&C condition is met. If the TT&C support time before each phase orbit control is greater than 3 hours, the TT&C condition is met, otherwise, the TT&C condition is not met.

[0045] Without considering the error, take the TT&C condition of the remote lunar maneuver orbit control time as an example, as shown in Figure 2 The horizontal axis is the deep space station and the vertical axis is the time in hours. It can be seen that the remote lunar maneuver orbit control time has a large delay time margin that can meet the TT&C constraint. At the same time, the simulation results show that the TT&C condition is basically met under the condition of orbit control time delay caused by error; while the orbit control time advance is easy to cause the TT&C support time to be less than 3 hours, which is more sensitive to the TT&C condition. The maneuvering of other phases is similar. For ease of analysis, only the TT&C time deviation caused by the advance of the orbit control time can be considered in the future to reduce the complexity of analysis.

[0046] Step 5: Analyze the influence of full-process error on the orbit 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 yes, the full-process error analysis is complete, if not, reserve the velocity increment for orbit correction to ensure sufficient safety margin during the mission and deorbiting conditions;

[0047] The minimum perigee height below 50km in the circumlunar phase is considered to have a risk of collision with the moon, and under the whole process error, a Monte Carlo targeting simulation is performed to obtain the proportion of the minimum perigee height below 50km in the circumlunar phase, so as to judge the risk of collision with the moon. If the proportion of the minimum perigee height below 50km in the mission orbit is higher than the reference value, it is indicated that there is a risk of collision with the moon, and if the proportion of the minimum perigee height below 50km in the mission orbit is lower than the reference value, it is indicated that the mission orbit is safe.

[0048] The minimum perigee height below 200km at the end of life is considered to have a controlled orbit departure capability. Under the whole process error, a Monte Carlo targeting simulation is performed to obtain the proportion of the minimum perigee height below 200km at the end of life, so as to judge whether the controlled orbit departure can be achieved. If the proportion of the minimum perigee height above 200km at the end of life is higher than the reference value, it is indicated that the controlled orbit departure cannot be achieved, and if the proportion of the minimum perigee height above 200km at the end of life is lower than the reference value, it is indicated that the controlled orbit departure can be achieved.

[0049] At the end of life, the controllable collision with the moon can be achieved, which can avoid the generation of lunar orbit debris and reduce the threat to future lunar exploration spacecraft. Considering the limited fuel available at the end of life, the orbit departure scheme under the condition of insufficient fuel is studied.

[0050] The circumlunar satellite is greatly affected by perturbation, and the law of the perigee height affected by perturbation is analyzed. The positive and negative of the perigee height determines whether the probe can achieve collision with the moon. As shown in Figure 3 , it can be seen that the minimum value of the perigee height changes constantly in the short term. By using the minimum value of the perigee height, a suitable opportunity is selected to apply a pulse to maximize the collision orbit control gain. The scheme of applying an acceleration increment at the apogee to reduce the perigee height below the lunar surface is adopted to achieve orbit departure. Combined with the minimum value of the perigee height, the orbit departure scheme is specifically designed as follows: after the end of the mission, the time when the minimum value of the perigee height is selected within the orbit departure time limit, and the apogee is run to the apogee, a pulse is applied in the opposite direction along the velocity direction, and collision with the moon is achieved near the perigee.

[0051] Under the condition of controllable orbit departure, the change curve of the height from the moon surface before and after the pulse is applied is shown in Figure 4 , in the figure, the abscissa is close to 495 days, the perigee height is at the minimum value of “115.3km”, and the pulse is applied from the minimum value point to the apogee, and the collision with the moon is achieved near the perigee.

[0052] The above specific embodiments further illustrate the purpose, technical scheme and advantages of the present application. It should be understood that the above description is only a specific embodiment of the present application and is not used to limit the present application. Any modification, equivalent replacement, improvement, etc. made within the spirit and principles of the present application shall be included in the protection scope of the present application.

Claims

1. A method for error processing and analysis throughout the entire process of transferring a large elliptical frozen orbit around the moon, characterized in that, The method comprises the following steps: Step 1: setting whole-process error items, including orbit insertion error, determination orbit error of each stage and orbit control error, wherein the stages include midcourse correction, capture braking, far-lunar-point maneuver and near-lunar-point deorbiting; Step 2: processing the whole-process error items, including processing the determination orbit error and processing the orbit control error, wherein the processing of the determination orbit error comprises determination orbit accuracy given by a TT&C system, determination orbit error data of each stage is randomly generated according to a normal distribution 3σ principle, with the nominal orbit being recursively propagated, the position and velocity RTN vector of the orbit control point of the corresponding stage are taken out and the separate targeting calculation is carried out by adding the determination orbit error of the corresponding stage, the orbit control velocity increment of the corresponding stage under the determination orbit error is obtained, which is taken as the orbit control velocity increment of the nominal orbit at the corresponding stage, and the orbit is recursively propagated to the next stage by considering the orbit control error, and the separate targeting calculation and orbit recursive propagation operation are repeated until the mission orbit is entered; the processing of the orbit control error comprises setting the range of azimuth and elevation angle error of the orbit control thrust direction, the azimuth and elevation angle error data is randomly generated according to the normal distribution 3σ principle and is added to the azimuth and elevation angle of the orbit control thrust direction of the nominal orbit; the orbit control velocity increment size error is calculated as 0.15+0.01443dv, wherein dv is the orbit control velocity increment of each stage, the velocity increment size error data is randomly generated according to the normal distribution 1σ principle and is added to the orbit control velocity increment value of the nominal orbit at each stage; Step 3: analyzing the influence of the whole-process error on the total velocity increment and judging whether the probe can enter the mission orbit, if yes, entering step 4, if no, adjusting the related characteristic parameters of the transfer orbit design; Step 4: analyzing the influence of the whole-process error on the TT&C condition and judging whether the TT&C condition is met, if yes, entering step 5, if no, changing the orbit control time of the subsequent stage by increasing or decreasing the capture period to ensure the TT&C condition; Step 5: analyzing the influence of the whole-process error on the orbit safety and the autonomous deorbiting at the end of the mission and judging whether the probe will be controlled to deorbit, if yes, the whole-process error analysis is ended, if no, reserving the velocity increment for orbit correction. The whole process in step 1 represents the whole-process maneuver from orbit insertion to the large-elliptical frozen orbit, including orbit insertion, midcourse correction, lunar braking capture, far-lunar-point maneuver, near-lunar-point maneuver, autonomous deorbiting at the end of the mission.

2. The method according to claim 1, wherein, The orbit control constraint is that there is TT&C support at least 3 hours before each orbit control, the lunar orbit is not collided with the moon within 1 year, and autonomous deorbiting is realized after the mission is completed.

3. The method according to claim 2, wherein, Step 3 comprises:

4. The method according to claim 1, wherein, Step 3.1: considering the whole-process error items, carrying out Monte Carlo targeting simulation on the whole-process orbit transfer to obtain the mean value, mean square deviation and maximum value of the total velocity increment consumed by the influence of the whole-process error items on the total velocity increment; Step 3.2: comparing whether the maximum value of the total velocity increment consumed is less than the velocity increment budget, if yes, the probe can enter the mission orbit. Step 4 comprises:

5. The method of claim 1, wherein the method further comprises: ​ With the measurement and control conditions of each stage without considering errors as reference values, Monte Carlo shooting simulation is performed under the conditions of considering orbit insertion error, orbit determination error and orbit control error to obtain the mean value and mean square deviation of the measurement and control conditions of each stage compared with the reference values; According to the maximum value of the measurement and control time in each stage calculated by 3 times of the mean square deviation, it is judged 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.

6. The method of claim 1, wherein the method further comprises: The step 5 comprises: It is considered that there is a risk of moon collision when the minimum perigee height of the circumlunar stage is lower than 50 km, Monte Carlo shooting simulation is performed under the whole process error to obtain the proportion of the minimum perigee height of the circumlunar stage being lower than 50 km, so as to judge the risk of moon collision, if the proportion of the minimum perigee height of the mission orbit being lower than 50 km is higher than the reference value, it is indicated that there is a risk of moon collision, if the proportion of the minimum perigee height of the mission orbit being lower than 50 km is lower than the reference value, it is indicated that the mission orbit is safe.

7. The method of claim 1, wherein the method further comprises: The step 5 further comprises: It is considered that there is a controlled orbit departure ability when the minimum perigee height at the end of the life is lower than 200 km, Monte Carlo shooting simulation is performed under the whole process error to obtain the proportion of the minimum perigee height at the end of the life being lower than 200 km, so as to judge whether the controlled orbit departure is possible, if the proportion of the minimum perigee height at the end of the life being higher than 200 km is higher than the reference value, it is indicated that the controlled orbit departure is not possible, if the proportion of the minimum perigee height at the end of the life being higher than 200 km is lower than the reference value, it is indicated that the controlled orbit departure is possible.

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