An optimal orbit change strategy for double-pulse co-intersection in a highly elliptical lunar orbit

Abstract

The present invention discloses a dual-pulse co-tangent optimal orbit change method for a large elliptical orbit around the moon. First, based on the known initial state and target orbit state of the lunar orbit, a dual-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 argument of the large elliptical orbit are adjusted simultaneously. The velocity increment minimum problem is converted into a one-dimensional nonlinear function minimum problem to solve the optimal orbit change strategy, thereby achieving precise transfer between any two elliptical orbits. Then, using the two-body orbit model design as the initial value, the two-body model is corrected by numerically integrating the average orbit elements under the influence of perturbations, thereby obtaining a real orbit change control strategy for the average orbit parameters under the influence of perturbations. This solves the problem that the short-period term introduced by directly using the instantaneous orbit element description turns the problem into a time-varying problem that is difficult to solve, thereby improving the engineering practical value of the algorithm.

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 a double-pulse co-tangent large elliptical orbit around the moon. Background Art

[0002] The Lunar Large Elliptical Frozen Orbit (LIFO) probe has strict requirements for the semi-major axis, eccentricity, and perigee. If these requirements are not met, the probe will gradually deviate from the target orbit over time and lose the freezing conditions. During on-orbit flight, factors such as high-order lunar gravitational perturbations, light pressure perturbations, orbital control errors, and engine attitude unloading cannot be accurately modeled. Their impact on the orbit must be mitigated through an effective maintenance strategy to ensure the long-term stable operation of the probe in its 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 large elliptical frozen orbit of the moon. 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 addressed by this invention is to address the deficiencies of existing technologies by proposing a novel dual-pulse co-tangent optimal trajectory change strategy for a large elliptical lunar orbit based on average orbital elements. This method, described using average orbital elements, simultaneously adjusts the average semi-major axis, average eccentricity, and average latitudinal argument of the large elliptical orbit through two pulses along the tangential direction. The key steps of this guidance strategy are as follows:

[0005] The technical solution of the present invention is: a method for optimal orbit change with double pulse co-tangency in a large elliptical orbit around the moon, comprising:

[0006] According to the obtained initial lunar orbit elements and the target lunar orbit elements, the initial guess value of the angular position of the first orbit control is set. , calculate the two pulse velocity increments and the total velocity increment of the double-pulse tangential orbit transfer under the two-body model;

[0007] Solve the angular position of the first track change control The angular position optimization solution of the first trajectory change corresponding to the two-body optimal velocity increment strategy is obtained by solving the one-dimensional nonlinear function minimum problem of the total velocity increment. , angular position optimization solution for the second track change And the optimal velocity increment of the two bodies during the two orbit changes 、 ;

[0008] The obtained angular position optimization solutions of the first orbit change, the angular position optimization solutions 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.

[0009] The initial lunar orbit elements , which is converted into the average orbital element Where, is the initial semi-major axis, is the initial eccentricity, is the initial orbit inclination, is the initial right ascension of the ascending node, is the initial perigee argument, is the initial mean anomaly; is the initial mean semi-major axis, is the initial average eccentricity, is the initial mean orbital inclination, is the initial mean ascending node right ascension, is the initial mean perigee argument, is the initial mean anomaly.

[0010] The orbital elements of the target lunar orbit include: the terminal mean semi-major axis , terminal average eccentricity , terminal average perigee argument , the average perigee argument difference between the initial orbit and the target orbit ; 、 and are the target mean semi-major axis, target mean eccentricity and target mean perigee argument at the terminal moment, is the initial mean perigee argument.

[0011] The calculation of the two pulse velocity increments and the total velocity increment of the double-pulse tangential trajectory change under the two-body model includes:

[0012] Based on the obtained orbital elements of the initial lunar orbit and the target lunar orbit, as well as the initial angular position value of the first orbit change control, the semi-path of the initial orbit and the semi-path of the target orbit are calculated;

[0013] Based on 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;

[0014] According to the angular positions of the first and second track changes, the auxiliary amount of the angular position calculation of the two track changes is calculated, and then the magnitude of the position vector of the two track change points is obtained;

[0015] The semi-diameter and eccentricity of the track after the first track change are calculated, and then the speed increments of the two track changes and the total speed increment are obtained based on the size of the position vectors of the two track change points.

[0016] The semi-diameter of the initial track , the semi-path of the target orbit .

[0017] The calculation of the average angular position difference between two track changes and the angular position of the second track change ,include:

[0018] The average angular position difference between two track changes :

[0019]

[0020] According to the first track change angle position Get the angular position of the second track change :

[0021] .

[0022] The auxiliary amount of angular position calculation of the two track changes for:

[0023]

[0024] ;

[0025] The magnitudes of the position vectors of the two track change points and Specifically:

[0026] .

[0027] The calculation of the semi-diameter and eccentricity of the track after the first track change includes:

[0028] The semi-diameter of the track after the first track change ;

[0029] The eccentricity of the track after the first track change .

[0030] The speed increments of the two track changes and the total speed increment are specifically:

[0031] The speed increment between two track changes:

[0032] ;

[0033] Total speed increment size :

[0034] .

[0035] The angular position optimization solution for the first trajectory change corresponding to the two-body optimal velocity increment strategy is obtained ,include:

[0036] In the process of calculating the two pulse velocity increments and the total velocity increment of the double-pulse tangential trajectory change under the two-body model, as long as the initial value of the angular position of the first trajectory change control is given arbitrarily, The total speed increment is calculated , the two constitute a one-dimensional nonlinear function:

[0037]

[0038] 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,

[0039] ;

[0040] 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 after two orbit changes 、 .

[0041] The method uses the obtained optimized solution of the first orbit change angular position and the two-body optimal velocity increments of the two orbit changes as initial values ​​for the perturbation solution, and corrects the orbit change process based on the average orbit perturbation equation, including:

[0042] 31) Based on the average orbital elements of the initial lunar orbit and two-body design results , orbit prediction is performed through the mean perturbation equation to obtain the orbit state at the terminal moment ;

[0043] 32) Calculate the target orbit element deviation based on the target aiming orbit parameters at the terminal time and the predicted current orbit parameters;

[0044] 33) Calculate the sensitivity matrix of trajectory change strategy correction;

[0045] 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 during the second track change; is the radial velocity increment component of the second orbit change;

[0046] 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.

[0047] The specific process of step 31) is as follows:

[0048] 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:

[0049]

[0050] in, is the average angular velocity of the probe around the moon; is the average angular velocity of the moon; 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 second-order harmonic coefficient of the moon's non-spherical gravitational field;

[0051] From the initial moment Number of orbital elements The average angular position corresponding to the first orbit 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 Forecast arrival 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 moment .

[0052] The target orbit element deviation is obtained by calculating:

[0053] According to the target mean semi-major axis at the terminal moment , target average eccentricity at the terminal moment and the target average latitude argument at the terminal moment , and the predicted current orbital parameters , the target orbit element deviation is calculated:

[0054] ;

[0055] in, 、 and are the differences between the mean semi-major axis, mean eccentricity and mean argument of perigee at the terminal moment and the target values, respectively.

[0056] Step 33) calculates the sensitivity matrix of the trajectory change strategy correction ,include:

[0057] .

[0058] Calculate the updated value of the velocity increment of the trajectory change strategy and its corresponding correction ,include:

[0059] The speed increments before updating are respectively recorded as , calculate the track change speed increment according to the following formula to obtain the updated speed increment:

[0060]

[0061] Then the speed increment correction is obtained:

[0062] .

[0063] The convergence criterion is .

[0064] The advantages of the present invention compared with the prior art are:

[0065] 1) Based on the traditional simplified model that only considers two bodies, this invention further adds a correction step for the lunar large elliptical orbit using an average orbit element model that considers the Earth's three-body gravity and the lunar J2 perturbation. This greatly improves the calculation accuracy and overcomes the problem of large control errors in the application of existing two-body model design results to engineering implementation, allowing the calculation results to be directly applied to engineering implementation.

[0066] 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 expanded from the original single-target periodic aiming to the multi-target joint aiming of period, eccentricity and perigee angle. This 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 large elliptical frozen relay orbits.

[0067] 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 spaceborne autonomous control. BRIEF DESCRIPTION OF THE DRAWINGS

[0068] Figure 1 Schematic diagram of the optimal trajectory change process of large elliptical double-pulse co-cutting;

[0069] Figure 2 This is a flowchart of the method of the present invention. DETAILED DESCRIPTION

[0070] The present invention provides a dual-pulse co-tangent optimal orbit change method for a large elliptical orbit around the moon. First, based on the known initial state and target orbit state of the lunar orbit, a dual-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 argument of the large elliptical orbit are adjusted simultaneously. The problem of minimizing the velocity increment is converted into a one-dimensional nonlinear function minimum problem to solve the optimal orbit change strategy. Then, using the two-body orbit model design as the initial value, the two-body model is modified by numerically integrating the average orbit elements under the influence of perturbations, thereby obtaining the actual orbit change control strategy for the average orbit parameters under the influence of perturbations. The specific steps are as follows:

[0071] 1. Obtain the initial and target lunar orbit states and set the initial values ​​of optimization variables

[0072] (1) Obtaining the initial instantaneous elements of the lunar orbit , converted into average orbital elements ;

[0073] (2) Obtain target orbit elements and average semi-major axis , average eccentricity , target average perigee argument , the average perigee argument difference between the initial orbit and the target orbit ;

[0074] (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 ;

[0075] 2. Given the first orbit change position, calculate the velocity increment of the double-pulse tangential orbit change under the two-body model

[0076] (1) Calculate the semi-paths of the initial and target orbits based on the input parameters and :

[0077] (1)

[0078] (2)

[0079] (2) Obtain the average angular position difference between the two track changes

[0080] (3)

[0081] According to the known first track change angle position Get the angular position of the second track change :

[0082] (4)

[0083] (3) According to calculate Two variables:

[0084] (5)

[0085] (6)

[0086] (4) Obtain the magnitude of the two orbit change position vectors and

[0087] (7)

[0088] (5) Calculate the semi-diameter of the track after the first track change

[0089] (8)

[0090] Calculate the eccentricity of the orbit after the first orbit change

[0091] (9)

[0092] (6) Calculate the velocity increment of the two track changes:

[0093] (10)

[0094] Calculate the total velocity increment :

[0095] (11)

[0096] 3. Optimize the first track change position to obtain the optimal speed increment strategy

[0097] According to the calculation ability difference in step 2, as long as the angular position of the first track change is given arbitrarily The total speed increment can be calculated , the two constitute a one-dimensional nonlinear function:

[0098] (12)

[0099] The single-pulse optimal trajectory change problem is transformed into the problem of finding the minimum value of a one-dimensional nonlinear function.

[0100] (13)

[0101] Using the one-dimensional nonlinear minimum solution method, the velocity increment can be obtained The optimal angular position of the minimum corresponding two track changes 、 And the optimal velocity increment of the two bodies during the two orbit changes 、 .

[0102] 4. Using the above results as initial values, the orbit change strategy is modified based on the average orbit perturbation equation.

[0103] (1) According to the initial state and two-body design results , orbit prediction is performed through the mean perturbation equation to obtain the orbit state at the terminal moment .

[0104] 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.

[0105] (14)

[0106] in,

[0107] is the average angular velocity of the probe around the moon; is the average angular velocity of the moon; 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 second-order harmonic coefficient of the moon's non-spherical gravitational field.

[0108] 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 .

[0109] According to the above method, from the initial moment Number of orbital elements Predicted 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 Forecast arrival 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 terminal status .

[0110] (2) Target aiming orbit parameters according to the terminal moment , and And the predicted current orbital parameters , the target orbit element deviation is calculated:

[0111] (15)

[0112] (3) Calculate the sensitivity matrix of the trajectory change strategy correction

[0113] (13)

[0114] Due to the adoption of the linear correction assumption, each orbital parameter can be substituted into the average orbital parameters at the initial moment. Perform calculations.

[0115] (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 :

[0116] Before updating, the current speed increment is recorded as , and then calculate the track change speed increment according to the following formula to obtain the updated speed increment:

[0117] (17)

[0118] Get the size of the speed increment correction:

[0119] (13)

[0120] (5) Iteration convergence judgment, if Then go to step (1), otherwise exit after convergence.

[0121] (6) After the algorithm converges, the output is the following parameters:

[0122] Angular position of the first track change , the corresponding track change speed increment ;

[0123] Average orbital elements before the first orbit change and the average orbital element after the first orbit change ;

[0124] Angular position of the second track change , the corresponding track change speed increment ;

[0125] Average orbital elements before the second orbit change and the average orbital element after the second orbit change .

[0126] The present invention proposes a new dual-pulse maintenance strategy suitable for a large elliptical frozen orbit around the moon. This strategy can simultaneously adjust the semi-major axis, eccentricity and perigee angle of the large elliptical orbit 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, overcoming the problem of large control errors when applying the existing two-body model design results to engineering implementation, and solving the problem of the short-period terms introduced by directly using instantaneous orbital elements to describe the problem, which makes the problem difficult to solve and turns it into a time-varying problem. 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 phase adjustment of the large elliptical frozen orbit. This strategy can overcome the problem that the existing single-pulse phase modulation strategy cannot ensure that the adjusted orbital parameters are strictly frozen, and has important engineering application value for the long-term stable operation of the large elliptical frozen relay orbit.

Claims

1. A method for optimal orbit change with double pulse co-cutting in a highly elliptical lunar orbit, characterized in that: include: According to the obtained initial lunar orbit elements and the target lunar orbit elements, the initial guess value of the angular position of the first orbit control is set. 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 minimum problem of the one-dimensional nonlinear function of the angular position θ1 of the first trajectory change control and the total velocity increment, and obtain the optimal solution of the angular position 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 solutions of the first orbit change, the angular position optimization solutions 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 pulse co-intersection in a lunar highly elliptical orbit according to claim 1, characterized in that: The initial lunar orbit element E0(t0)=(a0, e0, I0, Ω0, ω0, M0) is converted into the average orbit element E″0(t0)=(a″0, e″0, I″0, Ω″0, ω″0, M″0); where a0 is the initial semi-major axis, e0 is the initial eccentricity, I0 is the initial orbit inclination, Ω0 is the initial ascending node right ascension, ω0 is the initial perigee argument, and M0 is the initial mean anomaly; a″0 is the initial average semi-major axis, e″0 is the initial average eccentricity, I″0 is the initial average orbit inclination, Ω″0 is the initial average ascending node right ascension, ω″0 is the initial average perigee argument, and M″0 is the initial mean anomaly.

3. The method for optimal orbit change with double pulse co-intersection in a lunar highly elliptical orbit according to claim 2, characterized in that: The orbital elements of the target lunar orbit include: the terminal mean semi-major axis Terminal average eccentricity Terminal average perigee argument The average perigee argument difference between the initial orbit and the target orbit is Δω″=ω″ f -ω″0; and are the target mean semi-major axis, target mean eccentricity, and target mean argument of perigee at the terminal moment, respectively, and ω″0 is the initial mean argument of perigee.

4. The method for optimal orbit change with double pulse co-intersection in a lunar highly elliptical orbit according to claim 3 is characterized in that: The calculation of the two pulse velocity increments and the total velocity increment of the double-pulse tangential trajectory change under the two-body model includes: Based on the obtained orbital elements of the initial lunar orbit and the target lunar orbit, as well as the initial angular position value of the first orbit change control, the semi-path of the initial orbit and the semi-path of the target orbit are calculated; Based on 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 auxiliary amount of the angular position calculation 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 track after the first track change are calculated, and then the speed increments of the two track changes and the total speed increment are obtained based on the size of the position vectors of the two track change points.

5. The method for optimal orbit change with double pulse co-intersection in a lunar highly elliptical orbit according to claim 4 is characterized in that: The semi-diameter of the initial track Semi-diameter of target orbit 6. The method for optimal orbit change with double pulse co-intersection in a lunar highly elliptical orbit according to claim 5, characterized in that: The calculating the average angular position difference Δθ of the two track changes and the angular position θ2 of the second track change includes: The average angular position difference Δθ between two track changes is: The angular position θ2 of the second track change is obtained based on the angular position θ1 of the first track change: θ2=θ1+Δθ。 7. The method for optimal orbit change with double pulse co-intersection in a lunar highly elliptical orbit according to claim 6, characterized in that: The auxiliary values ​​(η1, η2) for calculating the angular position of the two track changes are: η2=(p′ f ′ / p′0′) / η1。 8. The method for optimal orbit change with double pulse co-cutting in a lunar highly elliptical orbit according to claim 7, characterized in that: The sizes r1″ and r2″ of the position vectors of the two track change points are specifically:

9. The method for optimal orbit change with double pulse co-intersection in a lunar highly 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 with double pulse co-intersection in a lunar highly elliptical orbit according to claim 9, characterized in that: The speed increments of the two track changes and the total speed increment are specifically: The speed increment between two track changes: The total velocity increment ΔV total : ΔV total =|ΔV 1t |+|ΔV 2t |。 11. The method for optimal orbit change with double pulse co-intersection in a lunar highly elliptical orbit according to claim 10, characterized in that: The angular position optimization solution for the first trajectory change corresponding to the two-body optimal velocity 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 trajectory change under the two-body model, as long as the initial value of the angular position θ1 of the first trajectory change control is given arbitrarily, a total velocity increment ΔV can be calculated. total , the two constitute a one-dimensional nonlinear function: ΔV total =function(θ1) 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 ΔV is calculated using a one-dimensional nonlinear minimum solution method. total 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 pulse co-intersection in a lunar highly elliptical orbit according to claim 11, characterized in that: The method uses the obtained optimized solution of the first orbit change angular position and the two-body optimal velocity increments of the two orbit changes as initial values ​​for the perturbation solution, and corrects the orbit change process based on the average orbit perturbation equation, including: 31) Based on the average orbital element E0″(t0) of the initial lunar orbit and the two-body design results The orbit state E′ at the terminal moment is obtained by orbit prediction through the average perturbation equation. f '; 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 sensitivity matrix of the orbit change strategy, the updated value of the velocity increment of the orbit change strategy (ΔV 1t ,ΔV 2t ,ΔV 2c ) and its corresponding correction, where ΔV 1t is the tangential velocity increment component of the first track change; ΔV 2t is the tangential velocity increment component of the second track change; ΔV 2c 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 with double pulse co-intersection in a lunar highly elliptical orbit according to claim 12, characterized in that: The target orbit element deviation is obtained by calculating: According to the target mean semi-major axis at the terminal moment Target average eccentricity at the terminal moment and the target average latitude argument at the terminal moment And the predicted current orbital parameter E′ f ′(t f )=(a′ f ′,e′ f ′,I′ f ′,Ω′ f ′,ω′ f ′,M′ f ′), the target orbit element deviation is calculated: Where Δa′ f ′、Δe′ f ′ and Δω′ f ′ are the differences between the mean semi-major axis, mean eccentricity and mean argument of perigee at the terminal moment and the target values.

14. The method for optimal orbit change with double pulse co-intersection in a lunar highly elliptical orbit according to claim 13, characterized in that: Step 33) Calculate the trajectory change strategy correction sensitivity matrix M, including:

15. The method for optimal orbit change with double pulse co-intersection in a lunar highly elliptical orbit according to claim 14, characterized in that: Calculate the updated value of the velocity increment of the trajectory change strategy (ΔV 1t ,ΔV 2t ,ΔV 2c ) and its corresponding correction value (δΔV 1t ,δΔV 2t ,δΔV 2c ),include: The speed increments before updating are respectively recorded as The track change speed increment is calculated according to the following formula to obtain the updated speed increment: Then the speed increment correction is obtained:

16. The method for optimal orbit change with double pulse co-intersection in a lunar highly elliptical orbit according to claim 15, characterized in that: The convergence criterion is |[δΔV 1t ,δΔV 2t ,δΔV 2c ] T |>10 -3 m / s; After convergence, the following parameters are output: Angular position of the first track change The corresponding track change speed increment ΔV1=(ΔV 1t ,0,0) T ; 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 ΔV2=(ΔV 2t ,ΔV 2c ,0) T ; Average orbital elements before the second orbit change and the average orbital element after the second orbit change

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