A rapid iterative guidance method for long-range guidance during lunar orbit rendezvous

Abstract

The present invention discloses a rapid iterative guidance method for lunar orbit rendezvous and remote guidance, comprising: obtaining an initial orbital state corresponding to an initial moment and an orbital target state corresponding to a terminal moment; establishing a nonlinear equation group of design variables and terminal orbital states; transforming the nonlinear equation group of design variables and terminal orbital states based on the near-circular orbit element representation of the terminal orbital state and the orbital target state; determining iterative initial values ​​and convergence conditions; iteratively solving the nonlinear equation group of the transformed design variables and terminal orbital states based on the determined iterative initial values ​​and convergence conditions until the convergence conditions are met, obtaining the final design variables, and completing the guidance. The method of the present invention has high iterative efficiency and fast calculation speed, is applicable to any number of orbit changes, and the components of each orbit change can be arbitrarily specified. It is particularly applicable to the needs of on-orbit real-time generation of lunar orbit rendezvous and remote guidance orbit change strategies and reconstruction of remote guidance strategies under fault conditions.

Description

Technical Field

[0001] The present invention belongs to the technical field of deep space exploration orbit design, and in particular relates to a lunar orbit rendezvous and long-distance guidance rapid iterative guidance method. Background Art

[0002] Remote guidance for lunar rendezvous and docking fully inherits the relevant strategies for low-Earth orbit manned spaceflight. However, a key difference lies in the implementation. During the flight control implementation phase of the lunar rendezvous and docking remote guidance trajectory change strategy, the online design, generation, and injection of the orbit control strategy must be completed in a short period of time based on real-time orbit determination results. The strategy calculation must account for the Moon's high-order gravitational field, typically exceeding 100th order to ensure accuracy. Furthermore, the three-body perturbation of the Earth and the Sun must also be considered, posing new challenges to the computational efficiency of the docking algorithm. Furthermore, the guidance algorithm requires differential refinement of the sensitivity matrix, which is obtained by numerically integrating 36 equations. A recursive algorithm is required to calculate the second-order partial derivatives of the Legendre polynomials associated with the high-order gravitational field, resulting in very low computational efficiency. Due to the large number of control variables, long integration time, and high accuracy requirements for the gravitational field model, traditional methods suffer from slow computational speed and poor convergence, making them difficult to adapt to the requirements of high-precision and rapid strategy reconstruction design under emergency conditions. Summary of the Invention

[0003] The technology of the present invention solves the problem of overcoming the shortcomings of the existing technology and providing a rapid iterative guidance method for lunar orbit rendezvous and remote guidance. The method has high iteration efficiency and fast calculation speed, is applicable to any number of orbit changes, and the components of each orbit change can be arbitrarily specified. The method is particularly suitable for the needs of on-orbit real-time generation of lunar orbit rendezvous and remote guidance orbit change strategies and reconstruction of remote guidance strategies under fault conditions.

[0004] In order to solve the above technical problems, the present invention discloses a rapid iterative guidance method for lunar orbit rendezvous and remote guidance, comprising:

[0005] Get the initial orbit state X0 and terminal time T corresponding to the initial time T0 f The corresponding orbital target state X f aim ;

[0006] Establish a nonlinear equation system of design variables and terminal orbital state; among them, the design variables include: the tangential component Δv of the first pulse 1T , the normal component of the second pulse Δv 2N , the latitude argument u2 of the second pulse, the tangential component Δv of the third pulse 3T , the tangential component of the fourth pulse Δv 4T and the radial component of the fourth pulse Δv 4R ;

[0007] Determine the terminal orbit state X f The near-circular orbit element representation E f and orbit target status Representation of near-circular orbit elements

[0008] Based on E f and Transforming the nonlinear equation group of the design variables and the terminal orbital state to obtain the transformed nonlinear equation group of the design variables and the terminal orbital state;

[0009] Determine the initial value of the iteration and the convergence conditions;

[0010] Based on the determined iterative initial values ​​and convergence conditions, the analytical correction strategy and the rapid orbit prediction algorithm of the lunar GAAF high-order gravitational field are used to iteratively solve the nonlinear equations of the transformed design variables and the terminal orbital state until the convergence conditions are met, and the final tangential component of the first pulse is obtained. The normal component of the second pulse Latitude argument of the second pulse The tangential component of the third pulse The tangential component of the fourth pulse and the radial component of the fourth pulse

[0011] according to and Determine the velocity increment vector for four trajectory changes and complete the guidance.

[0012] In the above-mentioned rapid iterative guidance method for lunar orbit rendezvous and remote guidance, the nonlinear equations of the design variables and the terminal orbit state are expressed as follows:

[0013] X f =G(Δv 1T ,Δv 2N ,u2,Δv 3T ,Δv 4T ,Δv 4R )

[0014] Where G(*) represents a nonlinear function of the gauge variable and the terminal orbital state.

[0015] In the above-mentioned lunar orbit rendezvous remote guidance rapid iterative guidance method, the terminal orbit state X f The near-circular orbit element representation E f It is expressed as follows:

[0016] E f =[a,e x ,e y,i,Ω,u]

[0017] Where a is the semi-major axis, i is the orbital inclination, Ω is the right ascension of the ascending node, u is the argument of latitude, and e is the orbital inclination. x and e y They represent the components of the eccentricity vector in the x-direction and y-direction of the orbital plane respectively.

[0018] In the above-mentioned lunar orbit rendezvous remote guidance rapid iterative guidance method, u=ω+θ, e x =ecosω,e y =esinω; where ω represents the argument of perihelion, θ represents the true anomaly, and e represents the eccentricity.

[0019] In the above-mentioned lunar orbit rendezvous remote guidance rapid iterative guidance method, the orbit target state Representation of near-circular orbit elements It is expressed as follows:

[0020]

[0021] in, represents the terminal target semi-major axis, represents the terminal target orbit inclination, represents the right ascension of the terminal target ascending node, Indicates the terminal target latitude argument, and They represent the components of the terminal target eccentricity vector in the x-direction and y-direction of the orbital plane, respectively.

[0022] In the above-mentioned lunar orbit rendezvous remote guidance rapid iterative guidance method, in, represents the terminal target perihelion argument, represents the true anomaly angle of the terminal target, represents the terminal target eccentricity.

[0023] In the above-mentioned rapid iterative guidance method for lunar orbit rendezvous and remote guidance, the nonlinear equations of the transformed design variables and the terminal orbit state are expressed as follows:

[0024]

[0025] In the above-mentioned rapid iterative guidance method for lunar orbit rendezvous and remote guidance, the initial value of the iteration is determined by the following method:

[0026] The phase difference from the initial time to the terminal time is calculated based on the latitude argument difference and circle difference between the initial orbit and the terminal orbit. Phase difference from the first orbit change to the terminal moment The semi-major axis a of the reference orbit r and reference speed V r ;

[0027] The reference orbit inclination i is calculated based on the orbit inclination difference between the initial state and the terminal target state and the right ascension difference of the ascending node. r and the right ascension of the reference ascending node Ω r ;

[0028] according to and a r , calculate the total phase difference Δθ from the initial state invariant orbit prediction to the terminal time:

[0029]

[0030] Where a0 represents the semi-major axis at the initial moment;

[0031] according to V r and Δθ, calculate the initial value of the tangential component of the first pulse

[0032]

[0033] Calculate the deviation of the orbital parameters from the final target state after the first pulse is applied to the initial orbit:

[0034]

[0035] Where Δa represents a f and The deviation, a f represents the semi-major axis of the terminal moment, represents the semi-major axis after a change of orbit; Δe x Indicates e xf and The deviation, e xf represents the x-direction component of the eccentricity vector at the terminal moment in the orbital plane, represents the x-direction component of the eccentricity vector after a change of orbit; Δe y Indicates e yf and The deviation, e yf represents the y-direction component of the eccentricity vector at the terminal moment in the orbital plane, It represents the y-direction component of the eccentricity vector after a track change; Δi represents i f and The deviation of i f represents the orbital inclination at the terminal moment, represents the orbital inclination after a track change; ΔΩ represents Ωf and Deviation, Ω f represents the right ascension of the ascending node at the terminal moment, Indicates the right ascension of the ascending node after one orbit change;

[0036] According to Δa, Δe x and Δe y , calculate the initial value of the tangential component of the third pulse The initial value of the tangential component of the fourth pulse and the initial value of the radial component of the fourth pulse

[0037]

[0038] Among them, Φ -1 represents the inverse matrix of the deviation state transfer matrix;

[0039] According to Δi and ΔΩ, the initial value of the latitude argument of the second pulse is calculated

[0040]

[0041] when When , the initial value of the normal component of the second pulse is for:

[0042]

[0043] when When , the initial value of the normal component of the second pulse is for:

[0044]

[0045] In the above-mentioned lunar orbit rendezvous remote guidance rapid iterative guidance method,

[0046]

[0047] Among them, u3 represents the latitude argument of the third orbit change point, and u4 represents the latitude argument of the fourth orbit change point.

[0048] In the above-mentioned rapid iterative guidance method for lunar orbit rendezvous remote guidance, the nonlinear equations of the transformed design variables and the terminal orbit state are iteratively solved in the following way:

[0049] Step 61, setting the initial state of the iteration, includes:

[0050] Dynamic aiming parameters

[0051] Let the design variables be initial values:

[0052] Let the number of iterations k = 0;

[0053] Set the deviation tolerance

[0054] Step 62: Based on the initial orbit state X0 and the initial values ​​of the design variables, the terminal state is calculated using the orbit predictor based on the GAAF algorithm.

[0055]

[0056] in, represents the terminal state semi-major axis, represents the terminal state orbit inclination, represents the right ascension of the ascending node in the terminal state, Indicates the terminal state latitude argument, and denote the components of the terminal state eccentricity vector in the x-direction and the y-direction of the orbital plane, respectively;

[0057] Step 63, according to and Calculate the off-target amount ΔE f :

[0058]

[0059] Where Δa f express and The difference, Δe xf express and The difference, Δe yf express and The difference, Δi f express and The difference, ΔΩ f express and The difference, Δu f express and The difference,

[0060] Step 64, according to ΔE f, update dynamic aiming parameters

[0061]

[0062] in, Represents the updated dynamic aiming parameters;

[0063] Step 65, according to ΔE f , update the tangential component of the first pulse:

[0064]

[0065] in, represents the tangential component of the first pulse after update, represents the tangential component of the first pulse before updating;

[0066] Step 66, according to Calculate the orbital parameters after the first orbit change and confirm and Deviation

[0067]

[0068] in, Indicates the deviation between the semi-major axis after a trajectory change and the updated dynamic aiming semi-major axis, It represents the deviation between the x-direction component of the eccentricity vector after a change of orbit and the x-direction component of the dynamic aiming eccentricity vector after the update. It represents the deviation between the y-direction component of the eccentricity vector after a change of orbit and the y-direction component of the dynamic aiming eccentricity vector after the update. Indicates the deviation between the orbital inclination after a change of orbit and the updated dynamic aiming orbital inclination. It represents the deviation between the right ascension of the ascending node after one orbit change and the right ascension of the ascending node after the dynamic sighting. Indicates the deviation between the latitude argument after a track change and the updated dynamic aiming latitude argument;

[0069] Step 67, according to Update the normal component of the second pulse, the argument of latitude of the second pulse, the tangential component of the third pulse, the tangential component of the fourth pulse, and the radial component of the fourth pulse:

[0070]

[0071] in, and represent the updated normal component of the second pulse, the latitude argument of the second pulse, the tangential component of the third pulse, the tangential component of the fourth pulse, and the radial component of the fourth pulse respectively;

[0072] Step 68, update the design variables:

[0073] Step 69, determine ΔE f Is it less than If ΔE f Less than Then stop the iteration and get and If ΔE f Not less than Then return to step 62 to perform iterative calculation until ΔE f Less than

[0074] The present invention has the following advantages:

[0075] (1) The present invention discloses a rapid iterative guidance method for long-range lunar rendezvous, which does not require calculation of a sensitivity matrix and has high iteration efficiency. Aiming is performed using an analytical algorithm based on linear approximation, and the correction iteration format is simple, typically achieving convergence within 4 to 6 iterations.

[0076] (2) The present invention discloses a rapid iterative guidance method for lunar orbit rendezvous and remote guidance, which has a fast computational speed. The computational speed is two orders of magnitude faster than that of traditional numerical integration algorithms, and the terminal state prediction accuracy error does not exceed 10m for position in three directions and 0.001m / s for velocity. The method is particularly suitable for the real-time on-orbit generation of lunar orbit rendezvous and remote guidance trajectory change strategies and the reconstruction of remote guidance strategies under fault conditions. BRIEF DESCRIPTION OF THE DRAWINGS

[0077] Figure 1 This is a flowchart of the steps of a rapid iterative guidance method for lunar orbit rendezvous and remote guidance in an embodiment of the present invention. DETAILED DESCRIPTION

[0078] In order to make the objectives, technical solutions and advantages of the present invention more clear, the embodiments disclosed in the present invention will be described in further detail below with reference to the accompanying drawings.

[0079] Reference Figure 1 In this embodiment, the lunar orbit rendezvous remote guidance rapid iterative guidance method includes:

[0080] Step 1: Get the initial orbit state X0 corresponding to the initial time T0 and the terminal time T f Corresponding orbit target state

[0081] Step 2: Establish a nonlinear equation set of design variables and terminal orbital states.

[0082] In this embodiment, the design variables mainly include: the tangential component Δv of the first pulse 1T , the normal component of the second pulse Δv 2N , the latitude argument u2 of the second pulse, the tangential component Δv of the third pulse 3T , the tangential component of the fourth pulse Δv 4T and the radial component of the fourth pulse Δv 4R .

[0083] The nonlinear equations of the design variables and the terminal orbital state are expressed as follows:

[0084] X f =G(Δv 1T ,Δv 2N ,u2,Δv 3T ,Δv 4T ,Δv 4R )

[0085] Where G(*) represents a nonlinear function of the gauge variable and the terminal orbital state.

[0086] Step 3: Determine the terminal orbit state X f The near-circular orbit element representation E f and orbit target status Representation of near-circular orbit elements

[0087] In this embodiment, the terminal track state X f The near-circular orbit element representation E f It is expressed as follows:

[0088] E f =[a,e x ,e y ,i,Ω,u]

[0089] Where a is the semi-major axis, i is the orbital inclination, Ω is the right ascension of the ascending node, u is the argument of latitude, and e is the orbital inclination. x and e y They represent the components of the eccentricity vector in the x and y directions of the orbital plane respectively; u=ω+θ, e x =ecosω,e y =esinω, ω represents the argument of perihelion, θ represents the true anomaly, and e represents the eccentricity.

[0090] Orbital target status Representation of near-circular orbit elements It is expressed as follows:

[0091]

[0092] in, represents the terminal target semi-major axis, represents the terminal target orbit inclination, represents the right ascension of the terminal target ascending node, Indicates the terminal target latitude argument, and They represent the components of the terminal target eccentricity vector in the x-direction and y-direction of the orbital plane respectively; represents the terminal target perihelion argument, represents the true anomaly angle of the terminal target, represents the terminal target eccentricity.

[0093] Step 4, based on E f and The nonlinear equation group of the design variables and the terminal orbital state is transformed to obtain the transformed nonlinear equation group of the design variables and the terminal orbital state.

[0094] In this embodiment, the nonlinear equations of the transformed design variables and the terminal orbital state are expressed as follows:

[0095]

[0096] Step 5: Determine the initial value of the iteration and the convergence condition.

[0097] In this embodiment, the initial value of the iteration can be determined by the following method:

[0098] The phase difference from the initial time to the terminal time is calculated based on the latitude argument difference and circle difference between the initial orbit and the terminal orbit. Phase difference from the first orbit change to the terminal moment The semi-major axis a of the reference orbit r and reference speed V r .

[0099] The reference orbit inclination i is calculated based on the orbit inclination difference between the initial state and the terminal target state and the right ascension difference of the ascending node. r and the right ascension of the reference ascending node Ω r .

[0100] according to and a r , calculate the total phase difference Δθ from the initial state invariant orbit prediction to the terminal time:

[0101]

[0102] Where a0 represents the semi-major axis at the initial moment.

[0103] according to V r and Δθ, calculate the initial value of the tangential component of the first pulse

[0104]

[0105] Calculate the deviation of the orbital parameters from the final target state after the first pulse is applied to the initial orbit:

[0106]

[0107] Where Δa represents a f and The deviation, a f represents the semi-major axis of the terminal moment, represents the semi-major axis after a change of orbit; Δe x Indicates e xf and The deviation, e xf represents the x-direction component of the eccentricity vector at the terminal moment in the orbital plane, represents the x-direction component of the eccentricity vector after a change of orbit; Δe y Indicates e yf and The deviation, e yf represents the y-direction component of the eccentricity vector at the terminal moment in the orbital plane, It represents the y-direction component of the eccentricity vector after a track change; Δi represents i f and The deviation of i f represents the orbital inclination at the terminal moment, represents the orbital inclination after a track change; ΔΩ represents Ω f and Deviation, Ω f represents the right ascension of the ascending node at the terminal moment, Indicates the right ascension of the ascending node after one orbit change.

[0108] According to Δa, Δe x and Δe y , calculate the initial value of the tangential component of the third pulse The initial value of the tangential component of the fourth pulse and the initial value of the radial component of the fourth pulse

[0109]

[0110] The inverse matrix Φ of the deviation state transfer matrix -1 It is expressed as follows:

[0111]

[0112] Among them, u3 represents the latitude argument of the third orbit change point, and u4 represents the latitude argument of the fourth orbit change point.

[0113] According to Δi and ΔΩ, the initial value of the latitude argument of the second pulse is calculated

[0114]

[0115] when When , the initial value of the normal component of the second pulse is for:

[0116]

[0117] when When , the initial value of the normal component of the second pulse is for:

[0118]

[0119] Step 6: Based on the determined initial values ​​and convergence conditions, the analytical correction strategy and the rapid orbit prediction algorithm of the lunar GAAF high-order gravitational field are used to iteratively solve the nonlinear equations of the transformed design variables and the terminal orbital state until the convergence conditions are met, and the final tangential component of the first pulse is obtained. The normal component of the second pulse Latitude argument of the second pulse The tangential component of the third pulse The tangential component of the fourth pulse and the radial component of the fourth pulse

[0120] In this embodiment, the nonlinear equations of the transformed design variables and the terminal track state can be iteratively solved in the following manner:

[0121] Step 61, set the initial state of the iteration:

[0122] Dynamic aiming parameters Let the design variables be initial values: Set the number of iterations k = 0; set the deviation tolerance

[0123] Step 62: Based on the initial orbit state X0 and the initial values ​​of the design variables, the terminal state is calculated using the orbit predictor based on the GAAF algorithm.

[0124]

[0125] in, represents the terminal state semi-major axis, represents the terminal state orbit inclination, represents the right ascension of the ascending node in the terminal state, Indicates the terminal state latitude argument, and are the components of the terminal state eccentricity vector in the x and y directions of the orbital plane, respectively.

[0126] Step 63, according to and Calculate the off-target amount ΔE f :

[0127]

[0128] Where Δa f express and The difference, Δe xf express and The difference, Δe yf express and The difference, Δi f express and The difference, ΔΩ f express and The difference, Δu f express and The difference,

[0129] Step 64, according to ΔE f , update dynamic aiming parameters

[0130]

[0131] in, Indicates the updated dynamic aiming parameters.

[0132] Step 65, according to ΔEf , update the tangential component of the first pulse:

[0133]

[0134] in, represents the tangential component of the first pulse after update, Represents the tangential component of the first pulse before updating.

[0135] Step 66, according to Calculate the orbital parameters after the first orbit change and confirm and Deviation

[0136]

[0137] in, Indicates the deviation between the semi-major axis after a trajectory change and the updated dynamic aiming semi-major axis, It represents the deviation between the x-direction component of the eccentricity vector after a change of orbit and the x-direction component of the dynamic aiming eccentricity vector after the update. It represents the deviation between the y-direction component of the eccentricity vector after a change of orbit and the y-direction component of the dynamic aiming eccentricity vector after the update. Indicates the deviation between the orbital inclination after a change of orbit and the updated dynamic aiming orbital inclination. It represents the deviation between the right ascension of the ascending node after one orbit change and the right ascension of the ascending node after the dynamic sighting. It indicates the deviation between the latitude argument after a track change and the updated dynamic aiming latitude argument.

[0138] Step 67, according to Update the normal component of the second pulse, the argument of latitude of the second pulse, the tangential component of the third pulse, the tangential component of the fourth pulse, and the radial component of the fourth pulse:

[0139]

[0140] in, and They respectively represent the updated normal component of the second pulse, the latitude argument of the second pulse, the tangential component of the third pulse, the tangential component of the fourth pulse and the radial component of the fourth pulse.

[0141] Step 68, update the design variables:

[0142] Step 69, determine ΔE fIs it less than If ΔE f Less than Then stop the iteration and get and If ΔE f Not less than Then return to step 62 to perform iterative calculation until ΔE f Less than

[0143] Step 7, according to and Determine the velocity increment vector for four trajectory changes and complete the guidance.

[0144] Although the present invention has been disclosed above in terms of preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solutions 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 solutions of the present invention shall fall within the scope of protection of the technical solutions of the present invention.

[0145] The contents not described in detail in the specification of the present invention belong to the common knowledge of professionals in this field.

Claims

1. A rapid iterative guidance method for lunar orbit rendezvous and remote guidance, characterized in that: include: Get the initial orbit state X0 and terminal time T corresponding to the initial time T0 f Corresponding orbit target state Establish a nonlinear equation system of design variables and terminal orbital state; among them, the design variables include: the tangential component Δv of the first pulse 1T , the normal component of the second pulse Δv 2N , the latitude argument u2 of the second pulse, the tangential component Δv of the third pulse 3T , the tangential component of the fourth pulse Δv 4T and the radial component of the fourth pulse Δv 4R ; Determine the terminal orbit state X f The near-circular orbit element representation E f and orbit target status Representation of near-circular orbit elements Based on E f and Transforming the nonlinear equation group of the design variables and the terminal orbital state to obtain the transformed nonlinear equation group of the design variables and the terminal orbital state; Determine the initial value of the iteration and the convergence conditions; Based on the determined iterative initial values ​​and convergence conditions, the analytical correction strategy and the rapid orbit prediction algorithm of the lunar GAAF high-order gravitational field are used to iteratively solve the nonlinear equations of the transformed design variables and the terminal orbital state until the convergence conditions are met, and the final tangential component of the first pulse is obtained. The normal component of the second pulse Latitude argument of the second pulse The tangential component of the third pulse The tangential component of the fourth pulse and the radial component of the fourth pulse according to and Determine the velocity increment vector for four trajectory changes and complete the guidance.

2. The rapid iterative guidance method for lunar orbit rendezvous and remote navigation according to claim 1, characterized in that: The nonlinear equations of the design variables and the terminal orbital state are expressed as follows: X f =G(Δv 1T ,Δv 2N ,u2,Δv 3T ,Δv 4T ,Δv 4R ) Where G(*) represents the nonlinear function of the design variables and the terminal orbital state.

3. The rapid iterative guidance method for lunar orbit rendezvous and remote navigation according to claim 2, characterized in that: Terminal orbit state X f The near-circular orbit element representation E f It is expressed as follows: E f =[a,e x ,e y ,i,Ω,u] Where a is the semi-major axis, i is the orbital inclination, Ω is the right ascension of the ascending node, u is the argument of latitude, and e is the orbital inclination. x and e y They represent the components of the eccentricity vector in the x-direction and y-direction of the orbital plane respectively.

4. The rapid iterative guidance method for lunar orbit rendezvous and remote navigation according to claim 3, characterized in that: u=ω+θ,e x =ecosω,e y =esinω; where ω represents the argument of perihelion, θ represents the true anomaly, and e represents the eccentricity.

5. The rapid iterative guidance method for lunar orbit rendezvous and remote navigation according to claim 4, characterized in that: Orbital target status Representation of near-circular orbit elements It is expressed as follows: in, represents the terminal target semi-major axis, represents the terminal target orbit inclination, represents the right ascension of the terminal target ascending node, Indicates the terminal target latitude argument, and They represent the components of the terminal target eccentricity vector in the x-direction and y-direction of the orbital plane, respectively.

6. The rapid iterative guidance method for lunar orbit rendezvous and remote navigation according to claim 5, characterized in that: in, represents the terminal target perihelion argument, represents the true anomaly angle of the terminal target, represents the terminal target eccentricity.

7. The rapid iterative guidance method for lunar orbit rendezvous and remote navigation according to claim 6, characterized in that: The nonlinear equations of the transformed design variables and terminal orbital states are expressed as follows:

8. The rapid iterative guidance method for lunar orbit rendezvous and remote navigation according to claim 7, characterized in that: The initial value of the iteration is determined as follows: The phase difference from the initial time to the terminal time is calculated based on the latitude argument difference and circle difference between the initial orbit and the terminal orbit. Phase difference from the first orbit change to the terminal moment The semi-major axis a of the reference orbit r and reference speed V r ; The reference orbit inclination i is calculated based on the orbit inclination difference between the initial state and the terminal target state and the right ascension difference of the ascending node. r and the right ascension of the reference ascending node Ω r ; according to and a r , calculate the total phase difference Δθ from the initial state invariant orbit prediction to the terminal time: Where a0 represents the semi-major axis at the initial moment; according to V r and Δθ, calculate the initial value of the tangential component of the first pulse Calculate the deviation of the orbital parameters from the final target state after the first pulse is applied to the initial orbit: Where Δa represents a f and The deviation, a f represents the semi-major axis of the terminal moment, represents the semi-major axis after a change of orbit; Δe x Indicates e xf and The deviation, e xf represents the x-direction component of the eccentricity vector at the terminal moment in the orbital plane, represents the x-direction component of the eccentricity vector after a change of orbit; Δe y Indicates e yf and The deviation, e yf represents the y-direction component of the eccentricity vector at the terminal moment in the orbital plane, It represents the y-direction component of the eccentricity vector after a track change; Δi represents i f and The deviation of i f represents the orbital inclination at the terminal moment, represents the orbital inclination after a track change; ΔΩ represents Ω f and Deviation, Ω f represents the right ascension of the ascending node at the terminal moment, Indicates the right ascension of the ascending node after one orbit change; According to Δa, Δe x and Δe y , calculate the initial value of the tangential component of the third pulse The initial value of the tangential component of the fourth pulse and the initial value of the radial component of the fourth pulse Among them, Φ -1 represents the inverse matrix of the deviation state transfer matrix; According to Δi and ΔΩ, the initial value of the latitude argument of the second pulse is calculated when When , the initial value of the normal component of the second pulse is for: when When , the initial value of the normal component of the second pulse is for:

9. The rapid iterative guidance method for lunar orbit rendezvous and remote navigation according to claim 8, characterized in that: Among them, u3 represents the latitude argument of the third orbit change point, and u4 represents the latitude argument of the fourth orbit change point.

10. The rapid iterative guidance method for lunar orbit rendezvous and remote navigation according to claim 9, characterized in that: The nonlinear equations of the transformed design variables and the terminal orbital state are iteratively solved as follows: Step 61, setting the initial state of the iteration, includes: Dynamic aiming parameters Let the design variables be initial values: Let the number of iterations k = 0; Set the deviation tolerance Step 62: Based on the initial orbit state X0 and the initial values ​​of the design variables, the terminal state is calculated using the orbit predictor based on the GAAF algorithm. in, represents the terminal state semi-major axis, represents the terminal state orbit inclination, represents the right ascension of the ascending node in the terminal state, Indicates the terminal state latitude argument, and denote the components of the terminal state eccentricity vector in the x-direction and the y-direction of the orbital plane, respectively; Step 63, according to and Calculate the off-target amount ΔE f : Where Δa f express and The difference, Δe xf express and The difference, Δe yf express and The difference, Δi f express and The difference, ΔΩ f express and The difference, Δu f express and The difference, Step 64, according to ΔE f , update dynamic aiming parameters in, Represents the updated dynamic aiming parameters; Step 65, according to ΔE f , update the tangential component of the first pulse: in, represents the tangential component of the first pulse after update, represents the tangential component of the first pulse before updating; Step 66, according to Calculate the orbital parameters after the first orbit change and confirm and Deviation in, Indicates the deviation between the semi-major axis after a trajectory change and the updated dynamic aiming semi-major axis, It represents the deviation between the x-direction component of the eccentricity vector after a change of orbit and the x-direction component of the dynamic aiming eccentricity vector after the update. It represents the deviation between the y-direction component of the eccentricity vector after a change of orbit and the y-direction component of the dynamic aiming eccentricity vector after the update. Indicates the deviation between the orbital inclination after a change of orbit and the updated dynamic aiming orbital inclination. It represents the deviation between the right ascension of the ascending node after one orbit change and the right ascension of the ascending node after the dynamic sighting. Indicates the deviation between the latitude argument after a track change and the updated dynamic aiming latitude argument; Step 67, according to Update the normal component of the second pulse, the argument of latitude of the second pulse, the tangential component of the third pulse, the tangential component of the fourth pulse, and the radial component of the fourth pulse: in, and represent the updated normal component of the second pulse, the latitude argument of the second pulse, the tangential component of the third pulse, the tangential component of the fourth pulse, and the radial component of the fourth pulse respectively; Step 68, update the design variables: Step 69, determine ΔE f Is it less than If ΔE f Less than Then stop the iteration and get and If ΔE f Not less than Then return to step 62 to perform iterative calculation until ΔE f Less than

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