Monopulse non-coplanar optimal near-moon capture method

Abstract

The invention discloses a monopulse non-coplanar optimal near-moon capture method, which adopts a pure analytical method for calculation, does not need numerical integration, parameter optimization and search, can directly obtain an optimal orbital transfer point position and an orbital transfer speed increment vector direction and size, and is very suitable for on-board autonomous calculation. The method can be directly applied to programming of software on a future autonomous near-moon brake controller, and also has important reference value for autonomous fault diagnosis and disposal of near-moon brake.

Description

Technical Field

[0001] The invention belongs to the technical field of deep space exploration orbit design, and in particular relates to a single pulse non-coplanar optimal near-moon capture method. Background Art

[0002] As the lunar exploration mission gradually deepens, the goals and tasks of near-moon braking have gradually become more complex. The near-moon braking strategy of the early lunar exploration mission is generally a coplanar optimal energy capture strategy, which has only one target variable, that is, the target semi-major axis of the lunar orbit. With the increase in mission requirements and the traction of the timed and fixed-point landing mission, near-moon braking must not only meet the requirements of the semi-field axis of the lunar orbit after capture, but also meet the requirements of the spatial orientation of the orbital plane after capture, so as to realize the subsequent timed and fixed-point landing mission. In order to meet the requirements of the timed and fixed-point landing mission, a near-moon braking capture strategy with a normal component is usually adopted. This strategy not only meets the requirements of the target semi-major axis, but also meets the requirements of the orientation of the orbital plane in space after capture by adjusting the target orbital inclination of the Earth-Moon transfer orbit and the size of the normal component. The existing method is to use precise numerical integration, adjust the parameters such as the yaw attitude angle of the engine and the start-up point position of the engine through numerical integration and combine parameter optimization for design. This method has the problems of large amount of calculation and difficulty in numerical convergence without initial values, and can only be applied to ground offline calculations. Summary of the invention

[0003] The technical problem solved by the present invention is: to overcome the deficiencies of the prior art and provide a single-pulse non-coplanar optimal near-moon capture method, which has the advantages of small calculation amount and high calculation speed.

[0004] The object of the present invention is achieved through the following technical solutions: A single pulse non-coplanar optimal near-moon capture method, comprising: step S1: obtaining the flight time of the Earth-Moon transfer segment according to the preset Earth-Moon transfer vehicle-rocket separation point time T0 and the preset perigee arrival time T1, and obtaining the vehicle-rocket separation point geocenter distance according to the preset vehicle-rocket separation point height H0; step S2: setting the initial value of the vehicle-rocket separation point flight path angle β0, and setting the initial value of the external point pseudo-state; wherein the initial value of the position vector of the external point pseudo-state is the position vector of the moon relative to the earth Step S3: Solve the Lambert problem of flight path angle constraint according to the distance from the center of the spacecraft to the rocket separation point R0, the initial value of the flight path angle β0 of the spacecraft to the rocket separation point, the preset orbital inclination i0 of the spacecraft to the rocket separation point, the flight time ΔT of the Earth-Moon transfer phase and the initial value of the pseudo-state of the external point, and obtain the state of the spacecraft to the rocket separation point (R0, V0) and the perigee velocity V1 * ; Wherein R0 is the position vector of the separation point of the device and arrow, and V0 is the velocity vector of the separation point of the device and arrow; Step S4: Set the external point pseudo state Transition to interior point pseudo state in, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo state; Step S5: according to the interior point pseudo state The lunar center hyperbolic orbit is the Earth-Moon transfer orbit vector Step S6: Based on the Earth-Moon transfer orbit vector The preset target lunar orbit parameters (r e ,v e ) and the preset initial value u1 of the latitude angle of the near-moon braking and orbit change point to obtain the optimized near-moon braking and orbit change point state r1 and Among them, r1 is the position vector of the near-moon braking and orbit change point, is the velocity vector before the near-moon braking and orbit change, r e is the position vector of the target lunar orbit, v e is the velocity vector of the target lunar orbit; Step S7: according to the optimized near-moon braking change point state r1 and After the two-body prediction Δt time, the updated interior point pseudo-state is obtained Step S8: Update the interior point pseudo-state Transition to a new outlier pseudostate According to the external point pseudo state and the new outlier pseudostate Get the modulus of the position error; Step S9: Determine whether the modulus of the position error converges. If so, set the new external point pseudo state Replace the outpoint pseudostate Go to step S3; if it does not converge, end.

[0005] In the above single pulse non-coplanar optimal near-moon capture method, the flight time of the Earth-Moon transfer phase is: ΔT = T1-T0; the distance from the center of the earth at the separation point of the rocket is: R0 = R E +H0; where T0 is the preset separation time of the earth-moon transfer vehicle and arrow, T1 is the preset arrival time at the perigee, ΔT is the flight time of the earth-moon transfer phase, R0 is the distance from the center of the earth to the vehicle and arrow separation point, H0 is the preset height of the vehicle and arrow separation point, R E is the average radius of the Earth.

[0006] In the above single pulse non-coplanar optimal near-moon capture method, the initial value of the flight path angle β0 at the separation point of the rocket is The initial values ​​of the external point pseudo-state are the lunar position and velocity at the time T1 when the lunar perigee is reached.

[0007] In the above single pulse non-coplanar optimal near-moon acquisition method, the interior point pseudo-state It is obtained by the following formula:

[0008]

[0009] in, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo-state, V1 is the pseudo-state position vector of the external point at time T1, * is the velocity vector of the pseudo-state of the external point at time T1, R M is the position vector of the moon relative to the earth, is the velocity vector of the interior point pseudo-state, V M is the velocity vector of the moon relative to the earth, Δt is the lookback time, which is generally taken as 0.65 times of ΔT.

[0010] In the above single pulse non-coplanar optimal near-moon acquisition method, according to the interior point pseudo-state The lunar center hyperbolic orbit is the Earth-Moon transfer orbit vector Including: According to the internal point pseudo state Calculate the eccentricity vector of a hyperbola According to the eccentricity vector of the hyperbola Get the eccentricity modulus and unit vector According to the interior point pseudo state Calculate the angular momentum vector of a hyperbola Calculating the Earth-Moon transfer orbit With the eccentricity vector Angle The unit vector Angular momentum vector Rotation Angle That is, the Earth-Moon transfer orbit vector The unit vector According to the interior point pseudo state Get the Earth-Moon transfer orbit vector The modulus of the Earth-Moon transfer orbit vector The unit vector and the Earth-Moon transfer orbit vector The modulus of the Earth-Moon transfer orbit vector

[0011] In the above single pulse non-coplanar optimal near-moon acquisition method, the eccentricity vector of the hyperbola It is obtained by the following formula:

[0012]

[0013] in, is the eccentricity vector of the hyperbola, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo-state, μ M is the lunar gravitational constant.

[0014] In the above single pulse non-coplanar optimal near-moon acquisition method, the eccentricity modulus is Unit Vector for

[0015] In the above single pulse non-coplanar optimal near-moon acquisition method, the angular momentum vector of the hyperbola It is obtained by the following formula:

[0016]

[0017] in, is the angular momentum vector of the hyperbola, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo state.

[0018] In the above single pulse non-coplanar optimal near-moon capture method, the Earth-Moon transfer orbit With the eccentricity vector Angle It is obtained by the following formula:

[0019]

[0020] In the above single pulse non-coplanar optimal near-moon capture method, the Earth-Moon transfer orbit vector The modulus value is obtained by the following formula:

[0021]

[0022] in, is the Earth-Moon transfer orbit vector The modulus value, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo-state, μ M is the lunar gravitational constant.

[0023] In the above single pulse non-coplanar optimal near-moon capture method, the Earth-Moon transfer orbit vector for

[0024] In the above single pulse non-coplanar optimal near-moon acquisition method, the new external point pseudo-state for:

[0025]

[0026] in, is the updated external point pseudo-state position vector, is the updated external point pseudo-state velocity vector, is the updated interior point pseudo-state position vector, is the updated interior point pseudo-state velocity vector, Δt is the backtracking time, R M is the position vector of the moon relative to the earth, V M is the velocity vector of the moon relative to the earth.

[0027] Compared with the prior art, the present invention has the following beneficial effects:

[0028] (1) The present invention has a small amount of calculation and is very suitable for on-orbit real-time calculation by satellite-borne computers;

[0029] (2) The present invention directly obtains the optimal track change position without iteration, and can provide initial value information with sufficient accuracy for ground precise search;

[0030] (3) The present invention has a fast calculation speed and can perform large-scale and large-scale parameter calculation and optimization during the preliminary orbit design to obtain the speed increment budget of the system's near-month braking. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] Various other advantages and benefits will become apparent to those of ordinary skill in the art by reading the detailed description of the preferred embodiments below. The accompanying drawings are only for the purpose of illustrating the preferred embodiments and are not to be considered as limiting the present invention. Moreover, the same reference symbols are used throughout the accompanying drawings to represent the same components. In the accompanying drawings:

[0032] Figure 1 is a flow chart of a single pulse non-coplanar optimal near-moon capture method provided by an embodiment of the present invention;

[0033] Figure 2 It is a schematic diagram of the angle relationship provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0034] Exemplary embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. Although exemplary embodiments of the present disclosure are shown in the accompanying drawings, it should be understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided in order to enable a more thorough understanding of the present disclosure and to be able to fully convey the scope of the present disclosure to those skilled in the art. It should be noted that, in the absence of conflict, the embodiments of the present invention and the features described in the embodiments can be combined with each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0035] At present, autonomous operation and autonomous management of spacecraft are a major trend and direction in the design of future spacecraft. In order to meet the needs of autonomous implementation and autonomous fault handling of the probe's near-moon braking process, this paper proposes a non-coplanar optimal near-moon braking capture strategy suitable for autonomous computing on onboard computing machines. This strategy uses a purely analytical method for calculation, without the need for numerical integration, and can directly obtain the optimal orbit change point position and the direction and size of the orbit change velocity increment vector. It is very suitable for on-board autonomous computing, with a small amount of calculation, and can be directly applied to the compilation of software on future autonomous near-moon braking controllers. It also has important reference value for autonomous fault diagnosis and handling of near-moon braking.

[0036] Figure 1 1 is a flow chart of a single pulse non-coplanar optimal near-moon capture method provided by an embodiment of the present invention. Figure 1 As shown, the single pulse non-coplanar optimal near-moon capture method includes:

[0037] Step S1: According to the preset earth-moon transfer vehicle-rocket separation point time T0 and the preset perigee arrival time T1, the earth-moon transfer segment flight time ΔT=T1-T0 is obtained, and according to the preset vehicle-rocket separation point height H0, the vehicle-rocket separation point earth center distance R0=R E +H0;

[0038] Step S2: Set the initial value of the flight path angle β0 at the separation point of the arrow Set the initial value of the external point pseudo-state to the lunar position and velocity at the time T1 when the lunar perigee is reached;

[0039] Step S3: Solve the Lambert problem of flight path angle constraint according to the distance from the center of the spacecraft to the rocket separation point R0, the initial value of the flight path angle β0 of the spacecraft to the rocket separation point, the preset orbital inclination i0 of the spacecraft to the rocket separation point, the flight time ΔT of the Earth-Moon transfer phase and the initial value of the pseudo-state of the external point, and obtain the state of the spacecraft to the rocket separation point (R0, V0) and the perigee velocity V1 * ; Wherein, R0 is the position vector of the separation point of the instrument and the arrow, and V0 is the velocity vector of the separation point of the instrument and the arrow;

[0040] Step S4: Set the external point pseudo state Transition to interior point pseudo state in, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo state;

[0041] Step S5: According to the interior point pseudo state The lunar-centered hyperbolic orbit is the Earth-Moon transfer orbit Vector;

[0042] Step S6: Based on the Earth-Moon transfer orbit vector, preset target lunar orbit parameters (r e,v e ) and the preset initial value u1 of the latitude angle of the near-moon braking and orbit change point to obtain the optimized near-moon braking and orbit change point state r1 and Among them, r1 is the position vector of the near-moon braking and orbit change point, is the velocity vector before the near-moon braking and orbit change, r e is the position vector of the target lunar orbit, v e is the velocity vector of the target lunar orbit;

[0043] Step S7: According to the optimized near-month braking change point state r1 and After the two-body prediction Δt time, the updated interior point pseudo-state is obtained

[0044] Step S8: Update the interior point pseudo-state Transition to a new outlier pseudostate According to the external point pseudo state and the new outlier pseudostate Get the modulus of the position error;

[0045] Step S9: Determine whether the modulus of the position error converges. If so, set the new external point pseudo state Replace the outpoint pseudostate Go to step S3; if it does not converge, end.

[0046] The flight time of the Earth-Moon transfer phase is: ΔT = T1-T0;

[0047] The distance from the center of the earth to the separation point of the rocket is: R0 = R E +H0;

[0048] Among them, T0 is the preset separation time of the earth-moon transfer vehicle and arrow, T1 is the preset arrival time at the perigee, ΔT is the flight time of the earth-moon transfer phase, R0 is the distance from the center of the earth to the separation point of the vehicle and arrow, H0 is the preset height of the separation point of the vehicle and arrow, R E is the average radius of the Earth.

[0049] The initial value of the flight path angle β0 at the separation point of the rocket is The initial values ​​of the external point pseudo-state are the lunar position and velocity at the time T1 when the lunar perigee is reached.

[0050] Interior point pseudostate It is obtained by the following formula:

[0051]

[0052] in, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo-state, V1 is the pseudo-state position vector of the external point at time T1,* is the pseudo-state velocity vector of the external point at time T1, R M is the position vector of the moon relative to the earth, is the velocity vector of the interior point pseudo-state, V M is the velocity vector of the moon relative to the earth; Δt is the lookback time, which is generally taken as 0.65 times of ΔT.

[0053] According to the interior point pseudo state The lunar center hyperbolic orbit is the Earth-Moon transfer orbit vector include:

[0054] According to the interior point pseudo state Calculate the eccentricity vector of a hyperbola According to the eccentricity vector of the hyperbola Get the eccentricity modulus and unit vector

[0055] According to the interior point pseudo state Calculate the angular momentum vector of a hyperbola

[0056] Calculating the Earth-Moon transfer orbit With the eccentricity vector Angle

[0057] The unit vector Angular momentum vector Rotation Angle That is, the Earth-Moon transfer orbit vector The unit vector

[0058] According to the interior point pseudo state Get the Earth-Moon transfer orbit vector The modulus of

[0059] According to the Earth-Moon transfer orbit vector The unit vector and the Earth-Moon transfer orbit vector The modulus of the Earth-Moon transfer orbit vector

[0060] Eccentricity vector of a hyperbola It is obtained by the following formula:

[0061]

[0062] in, is the eccentricity vector of the hyperbola, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo-state, μ Mis the lunar gravitational constant.

[0063] The eccentricity modulus is Unit Vector for

[0064] Hyperbolic angular momentum vector It is obtained by the following formula:

[0065]

[0066] in, is the angular momentum vector of the hyperbola, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo state.

[0067] Earth-Moon transfer orbit With the eccentricity vector Angle It is obtained by the following formula:

[0068]

[0069] Earth-Moon transfer orbit vector The modulus value is obtained by the following formula:

[0070]

[0071] in, is the Earth-Moon transfer orbit vector The modulus value, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo-state, μ M is the lunar gravitational constant.

[0072] Earth-Moon transfer orbit vector for

[0073] New Outlier Pseudostate for:

[0074]

[0075] in, is the updated external point pseudo-state position vector, is the updated external point pseudo-state velocity vector, is the updated interior point pseudo-state position vector, is the updated interior point pseudo-state velocity vector, Δt is the backtracking time, R M is the position vector of the moon relative to the earth, V M is the velocity vector of the moon relative to the earth.

[0076] Specifically, the step includes:

[0077] 1) Input conditions: Earth-Moon transfer vehicle-rocket separation time T0, arrival at perigee time T1, vehicle-rocket separation point height H0, vehicle-rocket separation point orbit inclination i0 and vehicle-rocket separation point true perigee angle f0, target lunar orbit parameters (r e ,v e ), the initial value of the latitude angle u1 of the near-month braking and orbit change point.

[0078] 2) According to the time T0 of the separation point of the earth-moon transfer vehicle and the rocket, the time T1 of reaching the perigee and the height H0 of the separation point of the vehicle and the rocket, calculate the flight time ΔT = T1-T0 of the earth-moon transfer segment and the distance R0 = R from the center of the earth to the separation point of the vehicle and the rocket. E +H0;

[0079] 3) Calculate the initial value of the flight path angle β0 at the arrow separation point:

[0080] 4) Look up the ephemeris table to obtain the lunar position at time T1 and the component R of the velocity vector in the geocentric J2000.0 inertial coordinate system M ,V M ;

[0081] 5) Set the initial value of the external point pseudo state to the moon position and velocity at time T1:

[0082] 6) Solve the Lambert problem with flight path angle constraints:

[0083] According to the distance from the center of the earth R0 to the separation point of the rocket and the flight path angle β0 of the separation point of the rocket, the Lambert algorithm is then used to solve the flight path angle constraint of the transfer segment:

[0084]

[0085] According to the state of the separation point (R0, V0), the eccentricity e0 of the arrow separation point is calculated and the flight path angle β0 is updated:

[0086]

[0087] 7) Convert the outer point pseudo state to the inner point pseudo state

[0088]

[0089] 8) If Figure 2 As shown, according to the interior point pseudo state The lunar center hyperbolic orbit is solved to obtain the Earth-Moon transfer orbit Vector:

[0090] 8.1) According to the state of the lunar center hyperbolic orbit Calculate the eccentricity vector of a hyperbola

[0091]

[0092] Get the eccentricity vector Calculate the eccentricity modulus and its unit vector

[0093] 8.2) According to the state of the lunar center hyperbolic orbit Calculate the angular momentum vector of a hyperbola

[0094]

[0095] 8.3) Calculate vector With vector Angle

[0096]

[0097] 8.4) Convert the unit vector Angular momentum Vector rotation angle You can get Unit vector of vector

[0098] 8.5) According to Calculated The modulus of a vector

[0099]

[0100] 8.6) According to Unit vector of vector and modulus Calculated Vector:

[0101]

[0102] 9) According to the calculation vector and the given target lunar orbit parameters (r e ,v e ), the initial value of the latitude argument u1 of the orbit change point, and calculate the optimal near-moon braking strategy. The specific steps are as follows:

[0103] 9.1) Calculate the angular momentum unit vector of the circular orbit around the moon

[0104]

[0105] Among them, r e is the target lunar orbit position vector, v e is the target lunar orbit velocity vector.

[0106] 9.2) Calculation The angle between the vector and the orbital plane around the moon is i ∞ :

[0107]

[0108] 9.3) Based on the given latitude argument u1 of the orbit change point on the lunar orbit, calculate the position vector of the orbit change point and velocity vector Among them, r1 - is the position vector before the track change, r1 + is the position vector after the track change, r1 is the position vector of the track change point,

[0109] 9.5) Calculate r1 + Vector with The angle Δθ between the vectors ∞ ;

[0110] 9.6) Calculate the eccentricity e of the hyperbolic orbit h for:

[0111]

[0112] Among them, r h is the modulus of the hyperbolic orbit position vector at the change point, V ∞ is the speed of the hyperbolic orbit relative to the infinite distance of the moon, μ is the gravitational constant of the moon, Δθ ∞ For r1 + Vector with The angle between the vectors.

[0113] 9.7) According to the solution of e h and the angle Δθ ∞ , the hyperbola flight path angle γ can be calculated as follows:

[0114]

[0115] 9.8) Now solve for the magnitude of the velocity vector after the change of trajectory:

[0116]

[0117] in, is the velocity vector modulus after the track change.

[0118] 9.9) According to the flight path angle γ, the velocity vector after the trajectory change can be further solved as:

[0119]

[0120] in, is the velocity vector after the trajectory change, is the position vector after the change of track, is the angular momentum vector after the change of orbit.

[0121] 9.10) Then, the track change velocity increment vector Δv1 can be solved as:

[0122]

[0123] in, is the velocity vector before the track change.

[0124] Thus, the size of the velocity increment Δv1=|Δv1| can be calculated.

[0125] According to the above algorithm, as long as the latitude argument u1 of the change point of the initial orbit is given arbitrarily, the position and velocity vector on the target elliptical orbit can be calculated, thus obtaining a satisfying The vector requires the hyperbolic track change speed increment Δv1, which together form a one-dimensional nonlinear function:

[0126] Δv1=f(u1) (16)

[0127] Among them, u1 is the latitude angle of the change point, and f(u1) is a one-dimensional nonlinear function.

[0128] The single pulse optimal trajectory change problem is transformed into the problem of finding the minimum value of the corresponding one-dimensional nonlinear function.

[0129]

[0130] 10) According to the optimized braking and track change point states r1 and After the two-body prediction Δt time, a new interior point pseudo-state is obtained

[0131]

[0132] Among them, r1 is the position vector of the track change point, is the velocity vector before the orbit change, Δt is the lookback time, TwoBody represents the two-body prediction model, is the updated interior point pseudo-state position vector, is the updated interior point pseudo-state velocity vector;

[0133] 11) The interior point pseudo state Transition to a new outlier pseudostate

[0134]

[0135] in, is the updated external point pseudo-state position vector, is the updated external point pseudo-state velocity vector, R M is the position vector of the moon relative to the earth, V M is the velocity vector of the moon relative to the earth.

[0136] 12) Calculate the new and old external point pseudo states The modulus of the position error ΔE:

[0137]

[0138] When ΔE ≥ ε, update the external point pseudo state Go to step (6), otherwise end.

[0139] Where ε is the iterative convergence threshold, which is usually taken as ε=10.

[0140] The above algorithm accurately matches the target parameters of the spacecraft-rocket separation point and obtains the orbit change point position r1 and velocity increment Δv1 for near-moon braking.

[0141] This embodiment has a small amount of calculation and is very suitable for on-orbit real-time calculation by the onboard computer; this embodiment does not require iteration and directly obtains the optimal orbit change position, which can provide initial value information with sufficient accuracy for ground-based precise search; this embodiment has a fast calculation speed and can perform large-scale and large-batch parameter calculation and optimization during preliminary orbit design to obtain a speed increment budget for near-moon braking of the system.

[0142] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A single pulse non-coplanar optimal near-moon acquisition method, characterized in that include: Step S1: Obtain the flight time of the Earth-Moon transfer segment according to the preset Earth-Moon transfer vehicle-rocket separation point time and the preset perigee arrival time, and obtain the distance from the vehicle-rocket separation point to the Earth's center according to the preset vehicle-rocket separation point height; Step S2: setting the initial value of the flight path angle of the arrow separation point and the initial value of the external point pseudo state; Step S3: Solve the Lambert problem of flight path angle constraint according to the distance from the center of the spacecraft-rocket separation point, the initial value of the flight path angle of the spacecraft-rocket separation point, the preset orbital inclination of the spacecraft-rocket separation point, the flight time of the Earth-Moon transfer phase and the initial value of the pseudo-state of the outer point, and obtain the state of the spacecraft-rocket separation point and the perigee velocity; Step S4: converting the outer point pseudo state into the inner point pseudo state; Step S5: Obtain the Earth-Moon transfer orbit vector v according to the lunar center hyperbolic orbit where the interior point pseudo-state is located ∞ - ; Step S6: obtaining an optimized near-moon braking orbit change point state according to the Earth-Moon transfer orbit vector, the preset target lunar orbit parameters and the preset near-moon braking orbit change point latitude argument initial value; Step S7: according to the optimized near-month braking and track change point state, after the two-body forecast time, an updated interior point pseudo state is obtained; Step S8: converting the updated interior point pseudo state into a new exterior point pseudo state, and obtaining a modulus value of the position error according to the exterior point pseudo state and the new exterior point pseudo state; Step S9: determine whether the modulus value of the position error converges. If so, replace the external point pseudo state with the new external point pseudo state, and go to step S3; If it does not converge, then terminate.

2. The single pulse non-coplanar optimal near-moon capture method according to claim 1, characterized in that: The flight time of the Earth-Moon transfer phase is: ΔT = T1-T0; The distance from the center of the earth to the separation point of the rocket is: R0 = R E +H0; Among them, T0 is the preset separation time of the earth-moon transfer vehicle and arrow, T1 is the preset arrival time at the perigee, ΔT is the flight time of the earth-moon transfer phase, R0 is the distance from the center of the earth to the separation point of the vehicle and arrow, H0 is the preset height of the separation point of the vehicle and arrow, R E is the average radius of the Earth.

3. The single pulse non-coplanar optimal near-moon capture method according to claim 1, characterized in that: The initial value of the flight path angle β0 at the separation point of the rocket is The initial values ​​of the external point pseudo-state are the lunar position and velocity at the time T1 when the lunar perigee is reached.

4. The single pulse non-coplanar optimal near-moon capture method according to claim 1, characterized in that: Interior point pseudostate It is obtained by the following formula: in, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo-state, V1 is the pseudo-state position vector of the external point at time T1, * is the velocity vector of the pseudo-state of the external point at time T1, R M is the position vector of the moon relative to the earth, is the velocity vector of the interior point pseudo-state, V M is the velocity vector of the moon relative to the earth, and Δt is the look-back time.

5. The single pulse non-coplanar optimal near-moon capture method according to claim 1, characterized in that: According to the interior point pseudo state The lunar center hyperbolic orbit is the Earth-Moon transfer orbit vector include: According to the interior point pseudo state Calculate the eccentricity vector of a hyperbola According to the eccentricity vector of the hyperbola Get the eccentricity modulus and unit vector According to the interior point pseudo state Calculate the angular momentum vector of a hyperbola Calculating the Earth-Moon transfer orbit With the eccentricity vector Angle The unit vector Angular momentum vector Rotation Angle That is, the Earth-Moon transfer orbit vector The unit vector According to the interior point pseudo state Get the Earth-Moon transfer orbit vector The modulus of According to the Earth-Moon transfer orbit vector The unit vector and the Earth-Moon transfer orbit vector The modulus of the Earth-Moon transfer orbit vector 6. The single pulse non-coplanar optimal near-moon acquisition method according to claim 5, characterized in that: Eccentricity vector of a hyperbola It is obtained by the following formula: in, is the eccentricity vector of the hyperbola, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo-state, μ M is the lunar gravitational constant.

7. The single pulse non-coplanar optimal near-moon acquisition method according to claim 5, characterized in that: The eccentricity modulus is Unit Vector for 8. The single pulse non-coplanar optimal near-moon acquisition method according to claim 5, characterized in that: Hyperbolic Angular Momentum Vector It is obtained by the following formula: in, is the angular momentum vector of the hyperbola, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo state.

9. The single pulse non-coplanar optimal near-moon acquisition method according to claim 5, characterized in that: Earth-Moon transfer orbit With the eccentricity vector Angle It is obtained by the following formula: Earth-Moon transfer orbit vector The modulus value is obtained by the following formula: in, is the Earth-Moon transfer orbit vector The modulus value, is the position vector of the interior point pseudo-state, is the velocity vector of the interior point pseudo-state, μ M is the lunar gravitational constant; Earth-Moon transfer orbit vector for 10. The single pulse non-coplanar optimal near-moon acquisition method according to claim 1, characterized in that: New Outlier Pseudostate for: in, is the new external point pseudo-state position vector, is the new external point pseudo-state velocity vector, is the updated interior point pseudo-state position vector, is the updated interior point pseudo-state velocity vector, Δt is the backtracking time, R M is the position vector of the moon relative to the earth, V M is the velocity vector of the moon relative to the earth.

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