Remote guidance rapid iterative guidance method for lunar orbit rendezvous

Abstract

The invention discloses a lunar orbit rendezvous remote guidance rapid iterative guidance method. The method comprises the following steps: acquiring an initial orbit state corresponding to an initial moment and an orbit target state corresponding to a terminal moment; establishing a nonlinear equation set of a design variable and a terminal track state; based on the near-circular orbit element representation forms of the terminal orbit state and the orbit target state, the nonlinear equation set of the design variables and the terminal orbit state is converted; determining an iteration initial value and a convergence condition; and on the basis of the determined iteration initial value and the convergence condition, iterative solution is conducted on the converted design variable and the nonlinear equation set of the terminal track state till the convergence condition is met, the final design variable is obtained, and guidance is completed. The method is high in iteration efficiency, high in calculation speed and suitable for any orbital transfer frequency, the component of each orbital transfer can be specified at will, and the method is particularly suitable for meeting the requirements for on-orbit real-time generation of the remote guidance orbital transfer strategy of the moon orbit intersection and remote guidance strategy reconstruction under the fault condition.

Description

Technical Field

[0001] The 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] The remote guidance of lunar orbit rendezvous and docking fully inherits the relevant strategies of manned spaceflight in low-Earth orbit. An important difference between the two lies in the implementation level. In the flight control implementation stage of the remote guidance of lunar orbit rendezvous and docking, it is necessary to complete the online design, generation and injection of the orbit control strategy based on the real-time orbit measurement results in a short time. The high-order gravitational field of the moon needs to be considered in the strategy calculation. To ensure the accuracy, it usually needs to be greater than 100 orders. At the same time, the three-body perturbation of the earth and the sun must also be considered, which poses new challenges to the efficiency of the docking algorithm calculation. On the other hand, the guidance algorithm requires differential improvement of the sensitivity matrix, which needs to be obtained by numerical integration of 36 equations, and the second-order partial derivatives of the Legendre polynomials of the high-order gravitational field need to be calculated by recursive algorithm, which has very low calculation efficiency. Due to the large number of control variables, long integration time and high accuracy requirements of the gravitational field model, the traditional method is difficult to adapt to the requirements of high-precision and fast strategy reconstruction design under emergency conditions due to its slow calculation speed and poor convergence. Summary of the invention

[0003] The technology of the present invention solves the problem: overcomes the shortcomings of the prior art and provides a rapid iterative guidance method for lunar orbit rendezvous and remote guidance, which 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, and 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 lunar orbit rendezvous remote guidance rapid iterative guidance method, comprising:

[0005] Get the initial time T 0 The corresponding initial orbital state X 0 and terminal time T f The corresponding orbital target state X f aim ;

[0006] Establish a nonlinear equation group 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 u of the second pulse 2 , the tangential component of the third pulse Δv 3T , the tangential component of the fourth pulse Δv 4Tand the radial component of the fourth pulse Δv 4R ;

[0007] Determine 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 orbit state to obtain the transformed nonlinear equation group of the design variables and the terminal orbit state;

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

[0010] Based on the determined initial values ​​and convergence conditions of the iteration, 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 orbit 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 orbit changes and complete the guidance.

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

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

[0014] Where G(*) represents a nonlinear function of the counter 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 latitude argument, 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.

[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 perigee, θ represents the true anomaly, and e represents the eccentricity.

[0019] In the above-mentioned lunar orbit rendezvous remote guidance rapid iteration 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, represents the terminal target latitude argument, and They represent the components of the terminal target eccentricity vector in the x and y directions of the orbital plane respectively.

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

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

[0024]

[0025] In the above-mentioned lunar orbit rendezvous remote guidance rapid iterative guidance method, the iteration initial value 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] Among them, a 0 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 applying the first pulse on the initial orbit:

[0034]

[0035] Among them, 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, It 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, represents the y-direction component of the eccentricity vector after a change of orbit; Δi represents if and The deviation of f represents the orbital inclination at the terminal time, represents the orbital inclination after a change of orbit; ΔΩ represents Ω f and Deviation of Ω f represents the right ascension of the ascending node at the terminal moment, It 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 is 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 angle 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, u 3 Indicates the latitude argument of the third orbit change point, u 4 Indicates the latitude argument of the fourth orbit change point.

[0048] In the above-mentioned lunar orbit rendezvous remote guidance rapid iterative guidance method, 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 iteration, includes:

[0050] Dynamic aiming parameters

[0051] Let each design variable be its initial value:

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

[0053] Set the deviation tolerance

[0054] Step 62: according to the initial track state X 0 and the initial values ​​of the design variables, and the orbit predictor based on the GAAF algorithm is used to calculate the terminal state

[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 They represent 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] Among them, 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, Indicates 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, It represents 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. It represents the deviation between the orbital inclination after a change of orbit and the updated dynamic aiming orbital inclination. It indicates the deviation between the right ascension of the ascending node after one orbit change and the right ascension of the dynamic aiming ascending node after update. 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 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:

[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 fast iterative guidance method for long-range guidance of lunar orbit rendezvous, which does not require the calculation of sensitivity matrix and has high iteration efficiency. It adopts an analytical algorithm based on linear approximation for aiming, and the correction iteration format is simple, and convergence can usually be completed through 4 to 6 iterations.

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

[0077] Figure 1 It is a flowchart of the steps of a lunar orbit rendezvous remote guidance rapid iterative guidance method 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 further described in 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 time T 0 The corresponding initial orbital state X 0 and terminal time T f Corresponding orbital target state

[0081] Step 2: Establish a set of nonlinear equations between the design variables and the terminal orbital state.

[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 u of the second pulse 2 , the tangential component of the third pulse Δv 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 ,u 2 ,Δv 3T ,Δv 4T ,Δv 4R )

[0085] Where G(*) represents a nonlinear function of the counter 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 latitude argument, 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 perigee, 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, represents the terminal target latitude argument, and They represent the components of the terminal target eccentricity vector in the x-direction and the y-direction of the orbital plane respectively; represents the terminal target perigee argument, represents the true near-point 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 iteration initial value 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] Among them, a 0 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 applying the first pulse on the initial orbit:

[0106]

[0107] Among them, 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, It 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, represents the y-direction component of the eccentricity vector after a change of orbit; Δi represents i f and The deviation of f represents the orbital inclination at the terminal time, represents the orbital inclination after a change of orbit; ΔΩ represents Ω f and Deviation of Ωf represents the right ascension of the ascending node at the terminal moment, Indicates the right ascension of the ascending node after an 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 is and the initial value of the radial component of the fourth pulse

[0109]

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

[0111]

[0112] Among them, u 3 Indicates the latitude argument of the third orbit change point, u 4 Indicates the latitude argument of the fourth orbit change point.

[0113] According to Δi and ΔΩ, the initial value of the latitude angle 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 orbit 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 each design variable be its initial value: Set the number of iterations k = 0; set the deviation tolerance

[0123] Step 62: according to the initial track state X 0 and the initial values ​​of the design variables, and the orbit predictor based on the GAAF algorithm is used to calculate the terminal state

[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 They represent 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] Among them, 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, Δuf 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 ΔE f , 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, It represents 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. It represents the deviation between the orbital inclination after a change of orbit and the updated dynamic aiming orbital inclination. It indicates the deviation between the right ascension of the ascending node after one orbit change and the right ascension of the dynamic aiming ascending node after update. 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 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:

[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 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

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

[0144] 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.

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

Claims

1. A lunar orbit rendezvous remote guidance rapid iterative guidance method, characterized in that: include: Get the initial orbit state X0 and terminal time T corresponding to the initial time T0 f Corresponding orbital target state Establish a nonlinear equation group 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 angle u2 of the second pulse, and 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 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 orbit state to obtain the transformed nonlinear equation group of the design variables and the terminal orbit state; Determine the initial value of the iteration and the convergence condition; Based on the determined initial values ​​and convergence conditions of the iteration, 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 orbit 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 orbit changes and complete the guidance.

2. The rapid iterative guidance method for lunar orbit rendezvous and remote guidance 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 a nonlinear function of the counter variable and the terminal orbital state.

3. The rapid iterative guidance method for lunar orbit rendezvous and remote guidance 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 latitude argument, 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.

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

5. The rapid iterative guidance method for lunar orbit rendezvous and remote guidance 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, represents the terminal target latitude argument, and They represent the components of the terminal target eccentricity vector in the x and y directions of the orbital plane respectively.

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

7. The rapid iterative guidance method for lunar orbit rendezvous and remote guidance 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 guidance 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 reference orbit ascending node right ascension 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 applying the first pulse on the initial orbit: Among them, 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, It 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, represents the y-direction component of the eccentricity vector after a change of orbit; Δi represents i f and The deviation of f represents the orbital inclination at the terminal moment, represents the orbital inclination after a change of orbit; ΔΩ represents Ω f and Deviation of Ω f represents the right ascension of the ascending node at the terminal moment, It 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 is 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 angle 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 guidance 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 lunar orbit rendezvous remote guidance rapid iterative guidance method according to claim 9, characterized in that: The nonlinear equations of the transformed design variables and the terminal orbital states are iteratively solved as follows: Step 61, setting the initial state of iteration, includes: Dynamic aiming parameters Let each design variable be its initial value: 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 orbit predictor based on the GAAF algorithm is used to calculate the terminal state 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 They represent 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 : Among them, 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, Indicates 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, It represents 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. It represents the deviation between the orbital inclination after a change of orbit and the updated dynamic aiming orbital inclination. It indicates the deviation between the right ascension of the ascending node after one orbit change and the right ascension of the dynamic aiming ascending node after update. 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 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: 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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