Double-pulse co-switching optimal orbital transfer strategy for large lunar elliptical orbit

Abstract

According to the double-pulse co-switching optimal orbital transfer method for the lunar large elliptical orbit, firstly, a double-pulse co-switching orbital transfer control strategy based on a two-body nonlinear model is calculated according to the known initial state of the lunar orbit and the state of a target orbit, and through two times of pulse control in the tangential direction, the optimal orbital transfer control strategy is obtained; meanwhile, the average semi-major axis, the average eccentricity rate and the average latitude argument of the large elliptical orbit are adjusted, the speed increment minimum problem is converted into the one-dimensional nonlinear function minimum problem to solve the optimal orbital transfer strategy, and accurate transfer between any two elliptical orbits is achieved; then, a two-body orbit model is designed as an initial value, numerical integration is carried out on the average orbit elements under the influence of perturbation force, the two-body model is corrected, and then a real orbital transfer control strategy of the average orbit parameters under the influence of perturbation force is obtained; the problem that the problem becomes a time-varying problem and is difficult to solve due to the fact that the short-cycle term introduced by directly adopting instantaneous orbit element description is solved, and the engineering practical value of the algorithm is improved.

Description

Technical Field

[0001] The present invention relates to the field of deep space exploration orbit design, and relates to an optimal orbit change strategy for double-pulse co-tangency in a large elliptical orbit around the moon. Background Art

[0002] The lunar large elliptical frozen orbit probe has strict requirements on the semi-major axis, eccentricity and perigee of the orbit. Once the requirements are not met, the probe will gradually deviate from the target orbit and lose the freezing conditions over time. During the on-orbit flight, the high-order perturbations of the lunar gravity, light pressure perturbations, orbit control errors, attitude unloading of the engine and other factors cannot be accurately modeled. The impact of these factors on the orbit must be eliminated through an effective maintenance strategy to ensure the long-term stable operation of the probe on the target mission orbit.

[0003] Through research, it was found that there is no detector in the world that can operate stably and long-term in the lunar large elliptical frozen orbit. Therefore, based on this orbital characteristic, it is currently impossible to find an orbit control strategy that can directly meet the orbit maintenance and phase adjustment requirements of engineering applications. Summary of the invention

[0004] The technical problem solved by the present invention is: to supplement the deficiencies of the prior art and propose a new large elliptical lunar orbit double pulse co-tangent optimal orbit change strategy based on average orbit elements. The method described by the present invention uses average orbit elements to describe, and through two pulse controls along the tangent direction, the average semi-major axis, average eccentricity and average latitude angle of the large elliptical orbit can be adjusted at the same time. The main steps of the guidance strategy are as follows: The technical solution of the present invention is: a double-pulse co-cutting optimal orbit change method for a large elliptical orbit around the moon, comprising: According to the acquired initial lunar orbit elements and the target lunar orbit elements, set the guessed initial value of the angular position of the first orbit change control , calculate the two pulse velocity increments and the total velocity increment of the double-pulse tangential orbit transfer under the two-body model; Solve the angular position of the first track change control The minimum value problem of the one-dimensional nonlinear function of the total velocity increment is solved to obtain the angular position optimization solution of the first trajectory change corresponding to the two-body optimal velocity increment strategy. , Angular position optimization solution for the second track change And the optimal velocity increment of the two bodies during the two orbit changes , ; The obtained angular position optimization solution of the first orbit change, the angular position optimization solution of the second orbit change and the two-body optimal velocity increments of the two orbit changes are used as the initial values ​​of the perturbation solution, and the orbit change process is corrected based on the average orbit perturbation equation.

[0005] The initial lunar orbit elements , which is converted into the average orbital element ; In the formula, is the initial semi-major axis, is the initial eccentricity, is the initial orbit inclination, is the initial ascending node right ascension, is the initial perigee argument, is the initial mean anomaly angle; is the initial average semi-major axis, is the initial average eccentricity, is the initial average orbital inclination, is the initial mean ascending node right ascension, is the initial average perigee argument, is the initial mean anomaly angle.

[0006] The orbital elements of the target lunar orbit include: the terminal average semi-major axis , the average eccentricity of the terminal , the terminal average perigee angle , the average perigee angle difference between the initial orbit and the target orbit ; , and are the target average semi-major axis, target average eccentricity and target average perigee argument at the terminal moment, is the initial average perigee argument.

[0007] The calculation of the two pulse velocity increments and the total velocity increment of the double-pulse tangential orbit change under the two-body model includes: According to the acquired orbital elements of the initial lunar orbit and the target lunar orbit, and the initial value of the angular position of the first orbit change control, the semi-diameter of the initial orbit and the semi-diameter of the target orbit are calculated; According to the obtained semi-paths of the initial orbit and the target orbit, the average angular position difference of the two orbit changes and the angular position of the second orbit change are calculated; According to the angular positions of the first and second track changes, the angular position calculation auxiliary quantity of the two track changes is calculated, and then the magnitude of the position vector of the two track change points is obtained; The semi-diameter and eccentricity of the orbit after the first orbit change are calculated, and then the velocity increment size of the two orbit changes and the total velocity increment size are obtained according to the size of the position vectors of the two orbit change points.

[0008] The semi-diameter of the initial track , the semi-diameter of the target orbit .

[0009] The calculation of the average angular position difference between two track changes and the angular position of the second track change ,include: The average angular position difference between two track changes :

[0010] According to the first track change angle position Get the angular position of the second track change : .

[0011] The auxiliary amount of angular position calculation of the two track changes for:

[0012] ; The magnitudes of the position vectors of the two track change points and Specifically: .

[0013] The calculation of the semi-diameter and eccentricity of the track after the first track change includes: The semi-diameter of the track after the first track change ; The eccentricity of the track after the first track change .

[0014] The speed increment size of the two track changes and the total speed increment size are specifically: The speed increment between two track changes is: ; Total speed increment size : .

[0015] The angular position optimization solution for the first track change corresponding to the two-body optimal speed increment strategy is obtained ,include: In the process of calculating the two pulse velocity increments and the total velocity increment of the double-pulse tangential orbit change under the two-body model, as long as the initial angular position of the first orbit change control is given arbitrarily, The total speed increment is calculated , the two constitute a one-dimensional nonlinear function:

[0016] The single pulse optimal trajectory change problem is transformed into the problem of finding the minimum value of a one-dimensional nonlinear function, that is, ; The velocity increment is calculated using a one-dimensional nonlinear minimum solution method. The minimum corresponding angular position of the first track change , Angular position optimization solution for the second track change And the optimal velocity increment of the two bodies during the two orbit changes , .

[0017] The method uses the obtained first orbit change angular position optimization solution and the two-body optimal velocity increments of the two orbit changes as initial values ​​of the perturbation solution, and corrects the orbit change process based on the average orbit perturbation equation, including: 31) Based on the average orbital elements of the initial lunar orbit and two-body design results , the orbit is predicted by the average perturbation equation to obtain the orbit state at the terminal time ; 32) Calculate the target orbit element deviation based on the target aiming orbit parameters at the terminal time and the predicted current orbit parameters; 33) Calculate the sensitivity matrix of trajectory change strategy correction; 34) According to the target orbit element deviation and the trajectory change strategy, the sensitivity matrix is ​​corrected and the updated value of the velocity increment of the trajectory change strategy is calculated. and its corresponding correction, where is the incremental component of the tangential velocity of the first track change; is the incremental component of the tangential velocity of the second track change; is the radial velocity increment component of the second orbit change; 35) Determine whether it converges. If not, return to step 31) and repeat steps 31) to 34) until convergence. If converged, end the calculation and finally obtain the final solution of the optimal orbit change method of double pulse co-tangent in the large elliptical orbit around the moon. ,in, The final solution after optimizing the tangential velocity increment component of the first track change; The final solution after optimizing the tangential velocity increment component of the second track change; This is the final solution after optimizing the radial velocity increment component of the second orbit change.

[0018] The specific process of step 31) is as follows: From the initial moment , for the average orbital elements Initial value, using numerical integration method, solve the following average orbital root ordinary differential equation to obtain any The average number of orbital elements corresponding to the time:

[0019] in, is the average angular velocity of the probe around the moon; is the average angular velocity of the moon’s revolution; is the average rotation angular velocity of the moon; is the Earth-Moon mass characteristic ratio coefficient; is the radius of the moon; is the coefficient of the second-order harmonic term of the non-spherical gravitational field of the moon; From the initial moment Number of orbital elements The average angular position corresponding to the first track change is predicted , get the average orbital elements before the first orbit change ,exist Apply pulse to change track , get the orbital elements after the first orbit change ;from Prediction After the circle, the average angular position corresponding to the second track change is reached , get the average orbital elements before the second orbit change ,exist Apply pulse orbit change to obtain the orbit element after the second orbit change ;from Forecast to terminal time Get the orbital state at the terminal time .

[0020] The target orbit element deviation is obtained by calculation, including: According to the target average semi-major axis at the terminal time , target average eccentricity at the terminal moment and the target average latitude angle at the terminal time , and the predicted current orbital parameters , the target orbit element deviation is calculated: ; in, , and They are the differences between the mean semi-major axis, mean eccentricity and mean perigee argument at the terminal moment and the target values.

[0021] Step 33) calculates the sensitivity matrix of the trajectory change strategy correction ,include: .

[0022] Calculate the updated value of the velocity increment for the track change strategy and its corresponding correction ,include: The speed increments before updating are respectively , calculate the track change speed increment according to the following formula to obtain the updated speed increment:

[0023] Then the speed increment correction is obtained: .

[0024] The convergence criterion is .

[0025] The advantages of the present invention compared with the prior art are: 1) Based on the traditional simplified model that only considers two bodies, the present invention further adds a step of using an average orbit element model that considers the three-body gravity of the Earth and the J2 perturbation of the moon to correct the large elliptical orbit of the moon, which greatly improves the calculation accuracy and overcomes the problem of large control errors when the existing two-body model design results are applied to engineering implementation, so that the calculation results can be directly applied to engineering implementation.

[0026] 2) Based on the traditional single-pulse tangential strategy, the present invention adds the tangential and normal components of the second pulse in the plane. The aiming strategy is changed from the original single-target periodic aiming to the period, eccentricity and perigee angle multi-target joint aiming, which overcomes the defect that the single-pulse strategy cannot guarantee the original orbit freezing characteristics after control. The new two-pulse strategy can still ensure that the orbit has strict freezing characteristics after control, which can effectively reduce the maintenance frequency, and has important engineering application value for the long-term stable operation of the large elliptical frozen relay orbit.

[0027] 3) Compared with the traditional method of using instantaneous orbit elements for numerical integration, the present invention has the characteristics of clear physical meaning, full analysis of the sensitivity matrix, high efficiency without the need for numerical integration, and is particularly suitable for the implementation of on-orbit missions. It can also be applied to the realization of future satellite-borne autonomous control. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] Figure 1 This is a schematic diagram of the optimal trajectory change process of large elliptical double pulse co-cutting; Figure 2 The figure is a flowchart of the method of the present invention. DETAILED DESCRIPTION

[0029] The present invention provides a double-pulse co-tangent optimal orbit change method for a large elliptical orbit around the moon. First, according to the known initial state of the lunar orbit and the target orbit state, a double-pulse co-tangent orbit change control strategy based on a two-body nonlinear model is calculated. Through two pulse controls along the tangent direction, the average semi-major axis, average eccentricity and average latitude angle of the large elliptical orbit are adjusted at the same time, and the speed increment minimum problem is converted into a one-dimensional nonlinear function minimum problem to solve the optimal orbit change strategy; then, the two-body orbit model design is used as the initial value, and the two-body model is corrected by numerically integrating the average orbit elements under the influence of the perturbation force, thereby obtaining the real orbit change control strategy of the average orbit parameters under the influence of the perturbation. The specific steps are as follows: 1. Obtain the initial and target lunar orbit states and set the initial values ​​of optimization variables (1) Obtaining the initial instantaneous elements of the lunar orbit , converted into average orbital elements ; (2) Obtain target orbit elements and average semi-major axis , the average eccentricity , target average perigee argument , the average perigee angle difference between the initial orbit and the target orbit ; (3) Set the initial value of the optimization variable: Set the angular position of the first track change control Set to guess initial value ,Right now ; 2. Given the first orbit change position, calculate the velocity increment of the double-pulse tangential orbit change under the two-body model

[0030] (1) Calculate the semi-paths of the initial and target orbits based on the input parameters and : (1) (2) (2) Obtain the average angular position difference between the two track changes

[0031] (3) According to the known first track change angle position Get the angular position of the second track change : (4) (3) Based on calculate Two variables: (5) (6) (4) Obtain the magnitude of the two orbit change position vectors and

[0032] (7) (5) Calculate the semi-diameter of the track after the first track change

[0033] (8) Calculate the eccentricity of the orbit after the first orbit change

[0034] (9) (6) Calculate the velocity increment of the two track changes: (10) Calculate the total velocity increment size : (11) 3. Optimize the first track change position to obtain the optimal speed increment strategy According to the difference in computing power in step 2, as long as the angular position of the first track change is given The total speed increment can be calculated , the two constitute a one-dimensional nonlinear function: (12) The single pulse optimal trajectory change problem is transformed into the problem of seeking the minimum value of a one-dimensional nonlinear function.

[0035] (13) Using the one-dimensional nonlinear minimum solution method, the velocity increment can be obtained The optimal angular position of the two minimum corresponding track changes , And the optimal velocity increment of the two bodies during the two orbit changes , .

[0036] 4. Using the above results as initial values, the orbit change strategy is modified based on the average orbit perturbation equation. (1) Based on the initial state and two-body design results , the orbit is predicted by the average perturbation equation to obtain the orbit state at the terminal time .

[0037] From the initial moment , for the average orbital elements Initial value, using RKF78 numerical integration method, solve the following average orbital root ordinary differential equation to obtain any The average number of orbital elements corresponding to the time.

[0038] (14) in, is the average angular velocity of the probe around the moon; is the average angular velocity of the moon’s revolution; is the average rotation angular velocity of the moon; is the Earth-Moon mass characteristic ratio coefficient, ; is the radius of the moon; is the coefficient of the second-order harmonic term of the moon’s non-spherical gravitational field.

[0039] Set the tangential velocity increment component of the first track change The initial value of is the two-body calculation result ,Right now ; The tangential velocity increment component of the second track change The initial value of is the two-body calculation result ,Right now , the radial velocity increment component of the second orbit change The initial value of is 0, that is .

[0040] According to the above method, from the initial moment Number of orbital elements Predict the angular position corresponding to the first track change , get the average orbital elements before the first orbit change ,exist Apply pulse to change track , get the orbital elements after the first orbit change .from Prediction After the circle, it reaches the angular position corresponding to the second track change , get the average orbital elements before the second orbit change ,exist Apply pulse orbit change to obtain the orbit element after the second orbit change .from Forecast to terminal time Get the status of the terminal .

[0041] (2) Target aiming orbit parameters based on the terminal moment , and And the predicted current orbital parameters , the target orbit element deviation is calculated: (15) (3) Calculate the sensitivity matrix of trajectory change strategy correction

[0042] (13) Due to the adoption of the linear correction assumption, each orbital parameter can be substituted into the average orbital parameters at the initial time. Perform calculations.

[0043] (4) According to the terminal aiming deviation And the modified sensitivity matrix , calculate the updated value of the velocity increment of the track change strategy and its corresponding correction : Before updating, the current speed increment is recorded as , and then calculate the track change speed increment according to the following formula to get the updated speed increment: (17) Get the size of the speed increment correction: (13) (5) Iteration convergence judgment: if Then go to step (1), otherwise exit after convergence.

[0044] (6) After the algorithm converges, the output parameters are as follows: Angular position of the first track change , the corresponding track change speed increment ; Average orbital elements before the first orbit change and the average orbital element after the first orbit change ; Angular position of the second track change , the corresponding track change speed increment ; Average orbital elements before the second orbit change and the average orbital element after the second orbit change .

[0045] The present invention proposes a new dual-pulse maintenance strategy suitable for a large elliptical frozen orbit around the moon. This strategy can adjust the semi-major axis, eccentricity and perigee angle of the large elliptical orbit at the same time through two pulse controls along the tangent direction. This strategy uses the average orbital elements that take into account the three-body gravity of the earth and the J2 perturbation of the moon to describe it, which overcomes the problem of large control errors when the existing two-body model design results are applied to engineering implementation, and solves the problem that the short-period terms introduced by directly using instantaneous orbital elements to describe the problem turn the problem into a time-varying problem that is difficult to solve. In addition, by adjusting the execution time interval of the two pulses and the target semi-major axis adjustment amount, this strategy can also be directly used for the phase adjustment of the large elliptical frozen orbit. This strategy can overcome the problem that the existing single-pulse phase modulation strategy cannot ensure the strict freezing of the adjusted orbital parameters, and has important engineering application value for the long-term stable operation of the large elliptical frozen relay orbit.

Claims

1. A double-pulse co-cutting optimal orbit change method for a large elliptical orbit around the moon, characterized in that: include: According to the acquired initial lunar orbit elements and the target lunar orbit elements, set the guessed initial value of the angular position of the first orbit change control , calculate the two pulse velocity increments and the total velocity increment of the double-pulse tangential orbit transfer under the two-body model; Solve the angular position of the first track change control The minimum value problem of the one-dimensional nonlinear function of the total velocity increment is solved to obtain the angular position optimization solution of the first trajectory change corresponding to the two-body optimal velocity increment strategy. , Angular position optimization solution for the second track change And the optimal velocity increment of the two bodies during the two orbit changes , ; The obtained angular position optimization solution of the first orbit change, the angular position optimization solution of the second orbit change and the two-body optimal velocity increments of the two orbit changes are used as the initial values ​​of the perturbation solution, and the orbit change process is corrected based on the average orbit perturbation equation.

2. The method for optimal orbit change with double pulses in a lunar large elliptical orbit according to claim 1, characterized in that: The initial lunar orbit elements , which is converted into the average orbital element ; In the formula, is the initial semi-major axis, is the initial eccentricity, is the initial orbit inclination, is the initial ascending node right ascension, is the initial perigee argument, is the initial mean anomaly angle; is the initial average semi-major axis, is the initial average eccentricity, is the initial average orbital inclination, is the initial mean ascending node right ascension, is the initial average perigee argument, is the initial mean anomaly angle.

3. The method for optimal orbit change with double pulses in a lunar large elliptical orbit according to claim 2, characterized in that: The orbital elements of the target lunar orbit include: the terminal average semi-major axis , terminal average eccentricity , the terminal average perigee angle , the average perigee angle difference between the initial orbit and the target orbit ; , and are the target average semi-major axis, target average eccentricity and target average perigee argument at the terminal moment, is the initial average perigee argument.

4. The method for optimal orbit change with double pulses in a lunar large elliptical orbit according to claim 2, characterized in that: The calculation of the two pulse velocity increments and the total velocity increment of the double-pulse tangential orbit change under the two-body model includes: According to the acquired orbital elements of the initial lunar orbit and the target lunar orbit, and the initial value of the angular position of the first orbit change control, the semi-diameter of the initial orbit and the semi-diameter of the target orbit are calculated; According to the obtained semi-paths of the initial orbit and the target orbit, the average angular position difference of the two orbit changes and the angular position of the second orbit change are calculated; According to the angular positions of the first and second track changes, the angular position calculation auxiliary quantity of the two track changes is calculated, and then the magnitude of the position vector of the two track change points is obtained; The semi-diameter and eccentricity of the orbit after the first orbit change are calculated, and then the velocity increment size of the two orbit changes and the total velocity increment size are obtained according to the size of the position vectors of the two orbit change points.

5. The method for optimal orbit change with double pulses in a lunar large elliptical orbit according to claim 4 is characterized in that: The semi-diameter of the initial track , the semi-diameter of the target orbit .

6. The method for optimal orbit change with double pulses in a lunar large elliptical orbit according to claim 5, characterized in that: The calculation of the average angular position difference between two track changes and the angular position of the second track change ,include: The average angular position difference between two track changes : According to the first track change angle position Get the angular position of the second track change : 。 7. The method for optimal orbit change with double pulses in a lunar large elliptical orbit according to claim 6 is characterized in that: The auxiliary amount of angular position calculation of the two track changes for: 。 8. The method for optimal orbit change with double pulses in a lunar large elliptical orbit according to claim 7 is characterized in that: The magnitudes of the position vectors of the two track change points and Specifically: 。 9. The method for optimal orbit change with double pulses in a lunar large elliptical orbit according to claim 8, characterized in that: The calculation of the semi-diameter and eccentricity of the track after the first track change includes: The semi-diameter of the track after the first track change ; The eccentricity of the track after the first track change .

10. The method for optimal orbit change of double pulse co-cutting in a lunar large elliptical orbit according to claim 9, characterized in that: The speed increment size of the two track changes and the total speed increment size are specifically: The speed increment between two track changes is: ； Total speed increment size : 。 11. The method for optimal orbit change of double pulse co-cutting in a lunar large elliptical orbit according to claim 10, characterized in that: The angular position optimization solution for the first track change corresponding to the two-body optimal speed increment strategy is obtained ,include: In the process of calculating the two pulse velocity increments and the total velocity increment of the double-pulse tangential orbit change under the two-body model, as long as the initial angular position of the first orbit change control is given arbitrarily, The total speed increment is calculated , the two constitute a one-dimensional nonlinear function: The single pulse optimal trajectory change problem is transformed into the problem of finding the minimum value of a one-dimensional nonlinear function, that is, ； The velocity increment is calculated using a one-dimensional nonlinear minimum solution method. The minimum corresponding angular position of the first track change , Angular position optimization solution for the second track change And the optimal velocity increment of the two bodies during the two orbit changes , .

12. The method for optimal orbit change with double pulses in a lunar large elliptical orbit according to claim 11, characterized in that: The method uses the obtained first orbit change angular position optimization solution and the two-body optimal velocity increments of the two orbit changes as initial values ​​of the perturbation solution, and corrects the orbit change process based on the average orbit perturbation equation, including: 31) Based on the average orbital elements of the initial lunar orbit and two-body design results , the orbit is predicted by the average perturbation equation to obtain the orbit state at the terminal time ; 32) Calculate the target orbit element deviation based on the target aiming orbit parameters at the terminal time and the predicted current orbit parameters; 33) Calculate the sensitivity matrix of trajectory change strategy correction; 34) According to the target orbit element deviation and the trajectory change strategy, the sensitivity matrix is ​​corrected and the updated value of the velocity increment of the trajectory change strategy is calculated. and its corresponding correction, where is the incremental component of the tangential velocity of the first track change; is the incremental component of the tangential velocity of the second track change; is the radial velocity increment component of the second orbit change; 35) Determine whether it converges. If not, return to step 31) and repeat steps 31) to 34) until convergence. If converged, end the calculation and finally obtain the final solution of the optimal orbit change method of double pulse co-tangent in the large elliptical orbit around the moon. ,in, The final solution after optimizing the tangential velocity increment component of the first track change; The final solution after optimizing the tangential velocity increment component of the second track change; This is the final solution after optimizing the radial velocity increment component of the second orbit change.

13. The method for optimal orbit change of double pulse co-cutting in a lunar large elliptical orbit according to claim 12, characterized in that: The specific process of step 31) is as follows: From the initial moment , for the average orbital elements Initial value, using numerical integration method, solve the following average orbital root ordinary differential equation to obtain any The average number of orbital elements corresponding to the time: in, is the average angular velocity of the probe around the moon; is the average angular velocity of the moon’s revolution; is the average rotation angular velocity of the moon; is the Earth-Moon mass characteristic ratio coefficient; is the radius of the moon; is the coefficient of the second-order harmonic term of the non-spherical gravitational field of the moon; From the initial moment Number of orbital elements The average angular position corresponding to the first track change is predicted , get the average orbital elements before the first orbit change ,exist Pulse change , get the orbital elements after the first orbit change ;from Prediction After the circle, the average angular position corresponding to the second track change is reached , get the average orbital elements before the second orbit change ,exist Apply pulse orbit change to obtain the orbit element after the second orbit change ;from Forecast to terminal time Get the orbital state at the terminal time .

14. The method for optimal orbit change of double pulse co-cutting in a lunar large elliptical orbit according to claim 13, characterized in that: The target orbit element deviation is obtained by calculation, including: According to the target average semi-major axis at the terminal time , target average eccentricity at the terminal moment and the target average latitude angle at the terminal time , and the predicted current orbital parameters , the target orbit element deviation is calculated: ； in, , and They are the differences between the mean semi-major axis, mean eccentricity and mean perigee argument at the terminal moment and the target values.

15. The method for optimal orbit change of double pulse co-cutting in a lunar large elliptical orbit according to claim 14, characterized in that: Step 33) calculates the sensitivity matrix of the trajectory change strategy correction ,include: 。 16. The method for optimal orbit change of double pulse co-cutting in a lunar large elliptical orbit according to claim 15, characterized in that: Calculate the updated value of the velocity increment for the track change strategy and its corresponding correction ,include: The speed increments before updating are respectively , calculate the track change speed increment according to the following formula to obtain the updated speed increment: Then the speed increment correction is obtained: 。 17. The method for optimal orbit change of double pulse co-cutting in a lunar large elliptical orbit according to claim 16, characterized in that: The convergence criterion is ; After convergence, the output parameters are as follows: Angular position of the first track change , the corresponding track change speed increment ; Average orbital elements before the first orbit change and the average orbital element after the first orbit change ; Angular position of the second track change , the corresponding track change speed increment ; Average orbital elements before the second orbit change and the average orbital element after the second orbit change .

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