Control method and device for entering large elliptical frozen orbit of moon

Abstract

The invention discloses a control method and device for entering a lunar large elliptical frozen orbit, and the method comprises the steps: carrying out the optimization solving of a far lunar maneuver pulse according to a particle swarm algorithm based on an orbit after the first near lunar braking of a satellite, and obtaining a pulse initial value of the far lunar maneuver. The pulse initial value of the far-moon maneuver represents the speed increment initial value of the satellite at the initial value of the far-moon maneuver moment. When the terminal parameter deviation corresponding to the pulse initial value is smaller than a threshold value, according to the pulse initial value of the far-moon maneuver and the target semi-major axis initial value of the second-time near-moon braking, combined planning is conducted on the far-moon maneuver and the second-time near-moon braking based on the Newton iteration method, and when the orbit parameter deviation at the third-time near-moon point moment is smaller than a convergence threshold value, the far-moon maneuver and the second-time near-moon braking are completed. The control parameters of far-moon maneuver and the control parameters of second-time near-moon braking are determined, the control parameters of third-time near-moon braking are calculated, and the control parameters for entering the lunar large elliptical frozen orbit are accurately calculated.

Description

Technical Field

[0001] The present invention relates to the field of aerospace measurement and control, and in particular to a control method and device for entering a lunar large elliptical freezing orbit. Background Art

[0002] Lunar satellites have a lunar large elliptical frozen orbit due to the north-south asymmetry of the gravitational field. The lunar large elliptical frozen orbit is a stable orbit whose eccentricity and perigee angle remain within a certain range. Since the frozen orbit has a stable state that remains unchanged or changes minimally for a long time, it is almost unnecessary to consume propellant for orbit maintenance, which greatly saves fuel consumption; at the same time, the apsidal directional characteristics of the lunar large elliptical frozen orbit can better achieve tracking and communication of specific positions or states.

[0003] For a probe that departs from the Earth, it generally takes four major orbital controls to finally enter the lunar large elliptical frozen orbit. After the first near-moon braking, the probe enters the capture orbit segment. Three controls are required from the capture orbit segment to entering the target frozen orbit, namely the far-moon maneuver control, the second near-moon braking, and the third near-moon braking. When designing the frozen orbit entry control strategy, it is also necessary to strictly target the designed control parameters, including orbital epoch time, semi-major axis, eccentricity, inclination, right ascension of the ascending node, perigee angle, etc. In addition, the capture orbit has a large semi-major axis, and the orbital elements are significantly affected by the Earth's gravitational perturbations, making it difficult to analytically solve the frozen orbit entry pulse.

[0004] Therefore, how to accurately calculate the control parameters for entering a frozen orbit is a problem that needs to be solved urgently. Summary of the invention

[0005] The embodiment of the present invention provides a control method for entering a large elliptical frozen orbit of the moon, which uses a particle swarm algorithm to optimize and solve, and obtains the initial value of the far-moon maneuver pulse. After obtaining the initial value, the Newton iteration method is used to further solve the far-moon maneuver and the second near-moon braking, and finally the control parameters of the third near-moon braking are calculated, thereby obtaining the control parameters of the far-moon maneuver, the second near-moon braking, and the third near-moon braking, and accurately calculating the control parameters for entering a large elliptical frozen orbit of the moon.

[0006] In a first aspect, an embodiment of the present invention provides a control method for entering a lunar large elliptical frozen orbit, comprising:

[0007] Based on the orbit of the satellite after the first near-moon braking, the far-moon maneuver pulse of the satellite is optimized and solved according to the particle swarm algorithm to obtain the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value, wherein the pulse initial value of the far-moon maneuver represents the initial value of the satellite's velocity increment at the initial value of the far-moon maneuver;

[0008] When the terminal parameter deviation corresponding to the pulse initial value is less than a threshold, the orbit parameter deviation corresponding to the third perigee time is iterated based on the Newton iteration method according to the pulse initial value of the far-moon maneuver and the target semi-major axis initial value of the second perigee braking;

[0009] When the orbital parameter deviation corresponding to the third perigee time is less than the convergence threshold, the control parameters of the far-moon maneuver and the control parameters of the second perigee braking are determined, and the control parameters of the third perigee braking are calculated, so that the satellite can enter the lunar large elliptical frozen orbit based on the control parameters of the far-moon maneuver, the control parameters of the second perigee braking and the control parameters of the third perigee braking.

[0010] In the above technical scheme, after the first near-moon braking, the satellite will enter the lunar capture orbit, and then it needs to enter the lunar large elliptical frozen orbit through the far-moon maneuver, the second near-moon braking and the third near-moon braking. Therefore, it is necessary to calculate the control parameters of the far-moon maneuver, the second near-moon braking and the third near-moon braking, so that the satellite can enter the lunar large elliptical frozen orbit based on the control parameters. Therefore, firstly, based on the capture orbit, the far-moon maneuver pulse of the satellite will be optimized and solved according to the particle swarm algorithm to obtain the initial value of the far-moon maneuver pulse and the terminal parameter deviation corresponding to the initial value of the pulse. The initial value of the far-moon maneuver pulse represents the initial value of the satellite's velocity increment at the initial value of the far-moon maneuver, including the initial value of the far-moon maneuver and the three-direction components of the initial value of the satellite's velocity increment in orbit at the initial value of the far-moon maneuver. The terminal parameters include the inclination of the satellite at the orbital perigee, the right ascension of the ascending node, the perigee height and the perigee argument.

[0011] Then, when the terminal parameter deviation corresponding to the pulse initial value is less than the threshold, the orbit parameter deviation corresponding to the third perigee time will be iterated based on the pulse initial value of the far-moon maneuver and the target semi-major axis initial value of the second perigee braking, and based on the Newton iteration method. The orbit parameters corresponding to the third perigee time include the third perigee time, inclination, right ascension of the ascending node, perigee height, and perigee amplitude. Until the orbit parameter deviation corresponding to the third perigee time is less than the convergence threshold, the control parameters of the far-moon maneuver and the control parameters of the second perigee braking are determined, and the control parameters of the third perigee braking are calculated, so as to accurately calculate the control parameters for entering the lunar large elliptical frozen orbit.

[0012] Optionally, based on the orbit of the satellite after the first near-moon braking, the far-moon maneuver pulse of the satellite is optimized and solved according to the particle swarm algorithm to obtain the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value, including:

[0013] Determine the range of far-moon maneuvering time based on the time when the satellite passes the southernmost or northernmost end of the lunar large elliptical frozen orbit when descending / ascending;

[0014] According to the value range of the initial value of the speed increment of the far-moon maneuver and the value range of the far-moon maneuver time, the far-moon maneuver pulse is optimized and solved based on the particle swarm algorithm. When the performance index of the far-moon maneuver is minimized, the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value are obtained; the performance index of the far-moon maneuver is obtained by weighted summation of the terminal parameter deviation based on the far-moon maneuver.

[0015] In the above technical scheme, after the first near-moon braking, the satellite enters the capture orbit, and then the moment when the satellite descends / ascends through the southernmost or northernmost end of the lunar large elliptical frozen orbit is used as the initial moment to determine the value range of the far-moon maneuvering moment. Finally, the far-moon maneuvering pulse is optimized and solved based on the particle swarm algorithm. When the performance index of the far-moon maneuver is minimized, the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value are obtained, and the solution problem of the far-moon maneuvering control strong nonlinear equation group is transformed into a global optimization problem. The particle swarm algorithm is used to optimize the pulse initial value of the far-moon maneuver, so as to obtain a better pulse initial value of the far-moon maneuver.

[0016] Optionally, after obtaining the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value, the method further includes:

[0017] When the terminal parameter deviation corresponding to the pulse initial value is greater than a threshold, the terminal parameter deviation is iterated based on the Newton iteration method according to the pulse initial value of the far-moon maneuver;

[0018] When the terminal parameter deviation is less than the threshold, the first pulse value of the corresponding far-moon maneuver is determined as the initial pulse value of the far-moon maneuver.

[0019] In the above technical solution, when the terminal parameter deviation corresponding to the pulse initial value is greater than a threshold value, the terminal parameter deviation is iterated according to the Newton iteration method based on the pulse initial value of the far-moon maneuver to reduce the terminal parameter deviation. When the terminal parameter deviation is less than the threshold value, the first pulse value of the corresponding far-moon maneuver is obtained and used as the pulse initial value of the far-moon maneuver, thereby improving the accuracy of the pulse initial value of the far-moon maneuver.

[0020] Optionally, according to the initial pulse value of the far-moon maneuver, the terminal parameter deviation is iterated based on the Newton iteration method, including:

[0021] Establishing a first partial derivative matrix between the initial pulse value of the far-moon maneuver and the terminal parameter deviation;

[0022] Iteration is performed based on the first partial derivative matrix until the terminal parameter deviation is less than a threshold.

[0023] Optionally, according to the initial pulse value of the far-moon maneuver and the initial target semi-major axis value of the second near-moon braking, the orbit parameter deviation corresponding to the third perigee time is iterated based on the Newton iteration method, including:

[0024] Establishing a second partial derivative matrix between the initial pulse value of the far-moon maneuver, the initial target semi-major axis value of the second near-moon braking, and the orbital parameter deviation corresponding to the third perigee time;

[0025] Iteration is performed based on the second partial derivative matrix until the orbit parameter deviation corresponding to the third perigee is less than a convergence threshold.

[0026] Optionally, after iterating the orbit parameter deviation corresponding to the third perigee time based on the Newton iteration method, the method further includes:

[0027] If the orbital parameter deviation corresponding to the third perigee time cannot be less than the convergence threshold, a third partial derivative matrix is ​​established between the pulse initial value of the far-moon maneuver, the target semi-major axis initial value of the second perigee braking, the target semi-major axis initial value of the third perigee braking and the orbital parameter deviation corresponding to the fourth perigee time;

[0028] Iteration is performed based on the third partial derivative matrix until the orbital parameter deviation corresponding to the fourth perigee moment is less than a convergence threshold, thereby determining control parameters of the far-moon maneuver, control parameters of the second perigee braking, and control parameters of the third perigee braking.

[0029] In a second aspect, an embodiment of the present invention provides a control device for entering a lunar large elliptical freezing orbit, comprising:

[0030] An acquisition module is used to optimize and solve the far-moon maneuver pulse of the satellite based on the orbit of the satellite after the first near-moon braking according to the particle swarm algorithm, and obtain the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value, wherein the pulse initial value of the far-moon maneuver represents the initial value of the satellite's velocity increment at the initial value of the far-moon maneuver moment;

[0031] A processing module, configured to iterate the orbit parameter deviation corresponding to the third perigee time based on the Newton iteration method according to the pulse initial value of the far-moon maneuver and the target semi-major axis initial value of the second perigee braking when the terminal parameter deviation corresponding to the pulse initial value is less than a threshold value;

[0032] When the orbital parameter deviation corresponding to the third perigee time is less than the convergence threshold, the control parameters of the far-moon maneuver and the control parameters of the second perigee braking are determined, and the control parameters of the third perigee braking are calculated, so that the satellite can enter the lunar large elliptical frozen orbit based on the control parameters of the far-moon maneuver, the control parameters of the second perigee braking and the control parameters of the third perigee braking.

[0033] Optionally, the processing module is specifically used for:

[0034] Determine the range of far-moon maneuvering time based on the time when the satellite passes the southernmost or northernmost end of the lunar large elliptical frozen orbit when descending / ascending;

[0035] According to the value range of the initial value of the speed increment of the far-moon maneuver and the value range of the far-moon maneuver time, the far-moon maneuver pulse is optimized and solved based on the particle swarm algorithm. When the performance index of the far-moon maneuver is minimized, the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value are obtained; the performance index of the far-moon maneuver is obtained by weighted summation of the terminal parameter deviation based on the far-moon maneuver.

[0036] Optionally, the processing module is further used for:

[0037] When the terminal parameter deviation corresponding to the pulse initial value is greater than a threshold, the terminal parameter deviation is iterated based on the Newton iteration method according to the pulse initial value of the far-moon maneuver;

[0038] When the terminal parameter deviation is less than the threshold, the first pulse value of the corresponding far-moon maneuver is determined as the initial pulse value of the far-moon maneuver.

[0039] Optionally, the processing module is specifically used for:

[0040] Establishing a first partial derivative matrix between the initial pulse value of the far-moon maneuver and the terminal parameter deviation;

[0041] Iteration is performed based on the first partial derivative matrix until the terminal parameter deviation is less than a threshold.

[0042] Optionally, the processing module is specifically used for:

[0043] Establishing a second partial derivative matrix between the initial pulse value of the far-moon maneuver, the initial target semi-major axis value of the second near-moon braking, and the orbital parameter deviation corresponding to the third perigee time;

[0044] Iteration is performed based on the second partial derivative matrix until the orbit parameter deviation corresponding to the third perigee is less than a convergence threshold.

[0045] Optionally, the processing module is further used for:

[0046] If the orbital parameter deviation corresponding to the third perigee time cannot be less than the convergence threshold, a third partial derivative matrix is ​​established between the pulse initial value of the far-moon maneuver, the target semi-major axis initial value of the second perigee braking, the target semi-major axis initial value of the third perigee braking and the orbital parameter deviation corresponding to the fourth perigee time;

[0047] Iteration is performed based on the third partial derivative matrix until the orbital parameter deviation corresponding to the fourth perigee moment is less than a convergence threshold, thereby determining control parameters of the far-moon maneuver, control parameters of the second perigee braking, and control parameters of the third perigee braking.

[0048] In a third aspect, an embodiment of the present invention further provides a computer device, including:

[0049] A memory for storing program instructions;

[0050] The processor is used to call the program instructions stored in the memory and execute the above-mentioned control method for entering the lunar large elliptical freezing orbit according to the obtained program.

[0051] In a fourth aspect, an embodiment of the present invention further provides a computer-readable storage medium, wherein the computer-readable storage medium stores computer-executable instructions, and the computer-executable instructions are used to enable a computer to execute the above-mentioned control method for entering a lunar large elliptical frozen orbit.

[0052] In a fifth aspect, an embodiment of the present invention further provides a computer program product, which includes an executable program, and the executable program is executed by a processor to implement the above-mentioned control method for entering a lunar large elliptical frozen orbit. BRIEF DESCRIPTION OF THE DRAWINGS

[0053] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the drawings required for use in the description of the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0054] Figure 1 A schematic diagram of orbital parameter changes provided by an embodiment of the present invention;

[0055] Figure 2 A schematic diagram of a system architecture provided by an embodiment of the present invention;

[0056] Figure 3 A schematic flow chart of a control method for entering a lunar large elliptical frozen orbit provided by an embodiment of the present invention;

[0057] Figure 4A schematic diagram of the structure of a control device for entering a lunar large elliptical freezing orbit provided in an embodiment of the present invention. DETAILED DESCRIPTION

[0058] In order to make the purpose, technical scheme and advantages of the present invention clearer, the present invention will be further described in detail below with reference to the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0059] The application scenarios described in the embodiments of the present application are intended to more clearly illustrate the technical solutions protected by the embodiments of the present application, and do not constitute a limitation on the technical solutions provided by the embodiments of the present application. It is known to those of ordinary skill in the art that with the emergence of new application scenarios, the technical solutions provided by the embodiments of the present application are also applicable to similar technical problems. The terms "first" and "second" in the specification and claims of this application and the above-mentioned drawings are used to distinguish different objects, rather than to describe a specific order. In the description of the present application, unless otherwise specified, the meaning of "multiple" is two or more.

[0060] Before introducing a control method for entering a lunar large elliptical frozen orbit provided in an embodiment of the present application, in order to facilitate understanding, the following terms and background technologies involved in the embodiment of the present application are first introduced.

[0061] Due to the north-south asymmetry of the gravitational field, lunar satellites have large elliptical frozen orbits (Lunar Large Elliptical Frozen Orbit). The lunar large elliptical frozen orbit is a stable orbit whose eccentricity and perigee argument remain in a relatively small range. Since the lunar large elliptical frozen orbit has a stable state that remains unchanged or changes minimally for a long time, the propellant used for orbit maintenance is minimal, which greatly saves fuel consumption; at the same time, the apsidal orientation characteristics of the lunar large elliptical frozen orbit can better achieve tracking and communication at a specific position or state. According to relevant theories, the perigee argument of the lunar large elliptical inclined frozen orbit is around 90° or 270°, and there is a specific relationship between the orbital inclination and eccentricity. Due to the constraints of the mission objectives (such as the relay communication duration, start and end times, etc. for the service target), the orbital period, phase and ascending node right ascension of the lunar large elliptical frozen orbit must also meet specific requirements. The safety of the perigee height during long-term operation of the lunar large elliptical frozen orbit, the change in the perigee argument (reflecting the degree of deviation of the apsidal line), and the visible arc segment of the service target are sensitive to the initial orbit parameters. Therefore, the initial parameters of the lunar large elliptical frozen orbit must be strictly designed. At the same time, when designing the entry control strategy of the lunar large elliptical frozen orbit, the designed initial parameters must also be strictly aimed at, including time, semi-major axis, eccentricity, inclination, right ascension of the ascending node, perigee argument, etc. (the true anomaly f=0 is the integration termination condition).

[0062] For a probe that departs from the Earth, it generally takes four major orbital controls to finally enter the lunar large elliptical frozen orbit. Since the orbit of the spacecraft relative to the moon is a hyperbolic orbit when it arrives at the moon, the first near-moon braking (LOI1, Lunar Orbit Insert 1) is mainly to make the spacecraft form a large elliptical orbit around the moon. This control is implemented near the satellite's first perigee on the moon. The pulse direction is in the opposite direction of the perigee speed, and the probe is decelerated and braked. Regarding the design of the first near-moon braking target, two ideas can be adopted. One is to be independent of the subsequent control and only realize the capture of the moon; the other is to jointly plan with the subsequent control so that the subsequent control can finally enter the target lunar large elliptical frozen orbit with the minimum speed increment consumption under the measurement and control constraints.

[0063] After the first near-moon braking, the probe enters the capture orbit segment. Three controls are required from the capture orbit segment to entering the target lunar large elliptical frozen orbit, namely, the far-moon maneuver control, the second near-moon braking, and the third near-moon braking. The far-moon maneuver control is the most important of the three controls. Its control quantities include the three-directional components of control timing and speed. The control targets are the inclination, right ascension of the ascending node, amplitude of the perigee, and altitude of the perigee of the target lunar large elliptical frozen orbit. This control is represented by LAM (Lunar Apogee Maneuver). The second near-moon braking (LOI2, Lunar OrbitInsert 2) is implemented when the satellite reaches the perigee again. The main function of this control is phase adjustment. By adjusting the orbital period, the time when the probe reaches the perigee next time is the orbital epoch in the initial parameters of the lunar large elliptical frozen orbit. It is implemented using tangential pulses. The third near-moon braking (LOI3, Lunar Orbit Insert 3) was implemented when the probe reached the perigee for the third time. At this time, the deviation of the probe from the target lunar large elliptical frozen orbit was only the semi-major axis. Therefore, the goal of this control was to adjust the semi-major axis of the orbit to be consistent with the target orbit, and tangential pulses were used for implementation.

[0064] After the spacecraft enters the capture orbit, it enters the lunar large elliptical frozen orbit with all orbital elements meeting the target requirements through three pulses. Since the three pulses, especially the first two pulses, influence each other, they jointly aim at the target lunar large elliptical frozen orbit parameters. Therefore, when solving the control parameters for entering the lunar large elliptical frozen orbit, it is necessary to jointly solve the three-pulse orbit control. In these three pulses, the far-moon maneuver controls four quantities at a time, and at this time the semi-major axis of the orbit is large and the orbit is relatively unstable. Therefore, the core problem of frozen orbit entry control lies in the solution of the far-moon maneuver pulse.

[0065] The capture orbit has a large semi-major axis and is significantly affected by the Earth's gravity. If we only consider the influence on the inclination, right ascension of the ascending node, amplitude of the perigee and the altitude of the perigee, there are two sets of solutions for the far-moon maneuver pulse. One set of solutions is obtained under the condition of small changes in orbital elements and can be solved analytically, but the semi-major axis of the orbit is small and the modulus of the velocity pulse is extremely large; the other set of solutions uses the Earth's gravitational perturbation to change the orbital plane, so the semi-major axis of the controlled orbit is large and the velocity pulse is small, but it is difficult to obtain the initial value analytically. Of these two solutions, the second one is the desired solution, and the key problem to be solved is the solution of the strongly nonlinear equations.

[0066] The far-moon maneuver is carried out in the capture orbit segment, when the orbital period is large and the perturbation effect of the Earth's gravity on the orbital elements is very significant. Figure 1 As shown, Figure 1A schematic diagram of orbital parameter changes provided for an embodiment of the present invention. The figure shows the instantaneous orbital element changes of an orbit with a semi-major axis of 30,000 km. Among the orbital element changes shown in the figure, the maximum inclination perturbation is about 11.2°, the maximum right ascension perturbation of the ascending node is 18.3°, the maximum perigee angle perturbation is about 3.3°, and the maximum perigee altitude perturbation is about 230km. Analysis shows that if the semi-major axis is further increased, the perturbation of the orbital elements will be more significant. For example, for an orbit with a semi-major axis of 50,000 km, the maximum inclination perturbation is about 35°, the maximum right ascension perturbation of the ascending node is 180°, the maximum perigee angle perturbation is about 52°, and the maximum perigee altitude perturbation is about 2100km. Since the orbital element perturbation is very drastic, it is no longer suitable to use the analytical method to calculate the initial pulse value at this time, because the analytical method is based on the small change of the orbital element. The use of Newton iteration method or MinPack method also has high requirements for the initial pulse value, otherwise the convergence speed will be extremely slow or divergent.

[0067] Newton's method: also known as the Newton-Raphson method, is a method proposed by Newton in the 17th century to approximate the solution of equations in the real and complex domains. This method approximates the roots of the equations through continuous iterations, uses Taylor series expansion and ignores higher-order terms to linearize nonlinear equations, thereby simplifying the solution process. Newton's method is not only applicable to solving the roots of a single equation, but can also be extended to solving differential equations and integral equations. The basic idea is to approximate the original function with the tangent of the function, and update the estimated value of the solution through continuous iterations until the required accuracy is reached.

[0068] Particle Swarm Optimization (PSO): An algorithm for optimizing a flock of foraging birds, where each "bird" represents a particle and the "food" the flock is looking for is the optimal solution. For a flying search space of m particles in D dimensions, the following definition is made.

[0069] The coordinates of the ith particle are:

[0070] The running speed of the i-th particle is:

[0071] The optimal position experienced by the i-th particle is:

[0072] The optimal position experienced by all particles in the population is:

[0073] The iterative results of particle motion speed and particle coordinates can be expressed as:

[0074]

[0075] Where: ω is the inertia weight, which plays a role in balancing global and local searches. A larger ω has better global convergence ability, and vice versa. c1 and c2 are learning factors, which control the ability of particles to find individual optimal positions and global optimal positions, respectively. r1 and r2 are pseudo-random numbers distributed between [0,1].

[0076] Figure 2 A system architecture applicable to an embodiment of the present invention is exemplarily shown. The system architecture includes a server 200 . The server 200 may include a processor 210 , a communication interface 220 , and a memory 230 .

[0077] The communication interface 220 is used to transmit data.

[0078] The processor 210 is the control center of the server 200, and uses various interfaces and routes to connect various parts of the entire server 200, and executes various functions and processes data of the server 200 by running or executing software programs and / or modules stored in the memory 230, and calling data stored in the memory 230. Optionally, the processor 210 may include one or more processing units.

[0079] The memory 230 can be used to store software programs and modules. The processor 210 executes various functional applications and data processing by running the software programs and modules stored in the memory 230. The memory 230 can mainly include a program storage area and a data storage area, wherein the program storage area can store an operating system, at least one application required for a function, etc.; the data storage area can store data created according to business processing, etc. In addition, the memory 230 can include a high-speed random access memory, and can also include a non-volatile memory, such as at least one disk storage device, a flash memory device, or other volatile solid-state storage devices.

[0080] It should be noted that the above Figure 2 The structure shown is only an example and is not limited to this embodiment of the present invention.

[0081] Based on the above description, Figure 3 A schematic flow chart of a control method for entering a lunar large elliptical frozen orbit provided by an embodiment of the present invention is exemplarily shown, and the flow chart can be executed by a control device for entering a lunar large elliptical frozen orbit.

[0082] like Figure 3 As shown in the figure, the process specifically includes:

[0083] Step 301, based on the orbit of the satellite after the first near-moon braking, the far-moon maneuver pulse of the satellite is optimized and solved according to the particle swarm algorithm to obtain the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value, wherein the pulse initial value of the far-moon maneuver represents the initial value of the satellite's velocity increment at the initial value of the far-moon maneuver moment.

[0084] In the embodiment of the present invention, after the satellite performs the first near-moon braking, it will enter the lunar capture orbit and can obtain the parameters of the capture orbit. Then, through the far-moon maneuver, the second near-moon braking and the third near-moon braking, it will enter the lunar large elliptical frozen orbit. It can be understood that the parameters of the target lunar large elliptical frozen orbit are known. For example, the capture orbit parameters of the lunar center level equatorial inertial system after the first near-moon braking are as follows:

[0085] Orbital epoch time: T0 = 2024-03-25T00:00:00.000

[0086] Semi-major axis a0 = 30000.0 km

[0087] Eccentricity e0 = 0.938753

[0088] Inclination angle i0 = 113.0°

[0089] Ascending node right ascension Ω0=276.0°

[0090] Perilunar angle ω0 = 127.0°

[0091] Mean anomaly M0 = 0°

[0092] The parameters of the target lunar great elliptical freezing orbit are as follows:

[0093] Orbital epoch time: T e =2024-03-31T21:25:00.000

[0094] Semi-major axis a e =9977.0km

[0095] Perilunar height Hp e =620.0km

[0096] Inclination angle i e =115.0°

[0097] Ascending node right ascension Ω e =230.0°

[0098] Perilunar argument ω e =88.0°

[0099] Mean anomaly angle M e =0°

[0100] First, after the first near-moon braking, the satellite enters the capture orbit, and then, according to the time when the satellite descends / ascends through the southernmost or northernmost end of the target orbit, the range of values ​​of the far-moon maneuvering time is determined. Exemplarily, the orbit after the first near-moon braking is integrated to the descending / ascending orbit passing the southernmost or northernmost end of the target orbit (the highest latitude, the southernmost end is taken for subsequent explanation), and its time is recorded. The point with the largest moon center distance is selected as the initial time, recorded as T0, and the time range of the far-moon maneuver is selected as [T0, T0+6h]. For example: the corresponding ascending orbit passing -65° south latitude is 2024-03-25T07:34:56, and the time range of the far-moon maneuver is [2024-03-25T07:34:56, 2024-03-25T13:34:56].

[0101] Then, determine the range of the initial value of the speed increment for the far-moon maneuver. The range of the initial value of the speed increment for the far-moon maneuver is a preset range based on experience, for example: ΔV x : Far-moon maneuver X-axis velocity increment, the value range is [-100m / s100m / s]; ΔV Y : Y-axis velocity increment for far-moon maneuver, the value range is [-100m / s 100m / s]; ΔV Z : The velocity increment in Z direction during the far-moon maneuver. The value range is [-100m / s 100m / s]. The value range of the pulse initial value is not specifically limited here.

[0102] Then, according to the value range of the initial value of the velocity increment of the far-moon maneuver and the value range of the far-moon maneuver time, the far-moon maneuver pulse is optimized and solved based on the particle swarm algorithm. When the performance index of the far-moon maneuver is minimized, the initial value of the pulse of the far-moon maneuver and the terminal parameter deviation corresponding to the initial value of the pulse are obtained. Among them, the initial value of the pulse of the far-moon maneuver represents the initial value of the velocity increment of the satellite at the initial value of the far-moon maneuver time, including the initial value of the far-moon maneuver time and the three-direction components of the initial value of the velocity increment of the satellite in orbit at the initial value of the far-moon maneuver time. The terminal parameters include the inclination of the satellite at the orbital perigee, the right ascension of the ascending node, the perigee height and the perigee argument. The performance index of the far-moon maneuver is obtained by weighted summation based on the terminal parameter deviation of the far-moon maneuver. Exemplarily, the orbit is integrated without control to the next perigee, and the deviations of the inclination i, the right ascension of the ascending node Ω, the perigee argument ω and the perigee height H from the target value are calculated, and the deviation results are recorded as Di, DΩ, Dω and DH. Then, according to the initial orbit and the apogee maneuvering time and velocity increment, the deviations of the inclination, right ascension of the ascending node, argument of the perigee and perigee height from the target values ​​are di, dΩ, dω and dH respectively. Finally, the performance index J is obtained as:

[0103]

[0104] The maximum value of the above performance index is 0. When the far-moon maneuver pulse is not considered, the performance index is -4. For the construction of the performance index, the size of the initial deviation of the terminal parameter is taken into account. Therefore, different weight coefficients can be assigned to each terminal parameter according to the initial deviation of the terminal parameter. The weight of the terminal parameter with a large deviation is larger, and the weight of the terminal parameter with a small deviation is smaller.

[0105] Exemplarily, the process of optimizing and solving the far-moon maneuver pulse based on the particle swarm algorithm includes: when using the particle swarm algorithm to optimize and solve the far-moon maneuver pulse, set the population size to 100, the number of generations to 5, the inertia weight to 1.0, the cognitive coefficient to 1.5, and the social coefficient to 1.5, and use the three-directional components of the pulse time deviation and the speed increment as optimization variables, limit the optimization variable range, and do not limit the speed range. Then, based on the relevant parameters set above, perform optimization and solve, and obtain the result of the particle swarm algorithm solution process:

[0106] Table 1. Results of the particle swarm optimization algorithm solution process

[0107]

[0108]

[0109] From the above table, we can see that the third generation of the particle swarm algorithm has the best result. It is selected as the initial pulse value for accurate calculation of the far-moon maneuver. The terminal parameter deviation corresponding to the initial value is shown in the following table:

[0110] Table 2. Terminal parameter deviations of far-moon maneuver results solved by particle swarm optimization algorithm

[0111] Parameters ΔH(km) Δi(°) Δω(°) ΔΩ(°) Deviation from target parameters 12.607 -0.497 0.095134 -6.432859

[0112] In one possible implementation, when the terminal parameter deviation corresponding to the pulse initial value is greater than a threshold, the terminal parameter deviation is iterated based on the Newton iteration method according to the pulse initial value of the far-moon maneuver. The threshold may be a value preset based on experience. For example, the perigee height threshold is set to 10 km, and the inclination, ascending node right ascension and perigee angle thresholds are set to 4.0°. The threshold is not specifically defined herein. Specifically, a first partial derivative matrix is ​​established between the pulse initial value of the far-moon maneuver and the terminal parameter deviation. Then, several iterations are performed based on the first partial derivative matrix until the terminal parameter deviation is less than the threshold. Exemplarily, it is assumed that the precise correction amount for the far-moon maneuver is [Δt a , Δv x , Δv y , Δv z ], then the detector is at T0+Δt a time, with [v x +Δv x , v y +Δvy , v z +Δv z ], and the terminal parameter when reaching the perigee is the inclination Ascending node right ascension Argument of perigee and perigee height It is expressed as:

[0113]

[0114] Among them, f, g, h, k are the constructed inclination angles Ascending node right ascension Argument of perigee and perigee height and T0+Δt a Time and speed x +Δv x , v y +Δv y , v z +Δv z ] Then perform Taylor expansion on the left side of the above equation:

[0115]

[0116] Then rearrange the above formula to get:

[0117]

[0118] Further sorting gives:

[0119]

[0120] The left side of the equation is the correction of the velocity increment component value and time, the first term on the right side is the inverse matrix of the partial derivative matrix of the four terminal parameters to the three velocity component values ​​and the time correction, and the second term on the right side is the difference between the actual terminal parameters and the expected terminal parameters. In addition, since it is difficult to write an analytical partial derivative matrix for the dynamic model of the real gravitational field, the partial derivative matrix can be approximated by the finite difference method, exemplarily using the control variable t a For example, as shown below:

[0121] (1) Calculate terminal parameters: e ,Ω e ,ω e , H e ] = f(t, r, v) (5)

[0122] (2) Under the control variable t a Apply a very small disturbance ε (for example, ε = 1s)

[0123] (3) Calculate the terminal parameters after disturbance:

[0124]

[0125] Calculate the four terminal parameters for the disturbance Δt a The partial derivative column vector of :

[0126]

[0127] Similarly, the perturbation ε is applied to the three-directional components of the velocity to obtain the complete partial derivative matrix. Repeat the iteration several times until the terminal parameter deviation is less than the threshold.

[0128] Exemplarily, the iterative calculation convergence process is shown in the following table:

[0129] Table 3. Convergence of the iterative solution of the far-moon maneuver

[0130]

[0131] As shown in the table above, the terminal parameter deviation corresponding to the far-moon maneuver is less than the threshold after the 15th iteration. When the terminal parameter deviation is less than the threshold, the first pulse value of the corresponding far-moon maneuver is determined as the initial pulse value of the far-moon maneuver.

[0132] Step 302, when the terminal parameter deviation corresponding to the pulse initial value is less than the threshold, the orbit parameter deviation corresponding to the third perigee time is iterated based on the Newton iteration method according to the pulse initial value of the far-moon maneuver and the target semi-major axis initial value of the second near-moon braking.

[0133] In the embodiment of the present invention, the initial pulse value of the far-moon maneuver calculated through the above steps is relatively close to the actual pulse value of the far-moon maneuver, and the terminal parameter deviation corresponding to the far-moon maneuver is already small. Therefore, the far-moon maneuver and the second near-moon braking can be jointly solved to obtain the control parameters of the far-moon maneuver, the control parameters of the second near-moon braking, and the control parameters of the third near-moon braking.

[0134] Specifically, when the terminal parameter deviation corresponding to the far-moon maneuver is less than the threshold, the second partial derivative matrix between the pulse initial value of the far-moon maneuver, the target semi-major axis initial value of the second near-moon braking, and the orbital parameter deviation corresponding to the third perigee time is established. Among them, the target semi-major axis initial value of the second near-moon braking can be the semi-major axis of the target lunar large elliptical frozen orbit, or it can be calculated according to the time difference (i.e., orbital period) between the second near-moon braking moment and the initial epoch of the target lunar large elliptical frozen orbit (the third near-moon braking moment) to obtain the target semi-major axis initial value, and then the true target semi-major axis is obtained through iterative optimization. The method of obtaining the target semi-major axis initial value is not specifically limited here. The orbital parameters corresponding to the third perigee time include the third perigee time, inclination, right ascension of ascending node, perigee height, and perigee amplitude angle. Exemplarily, let the target semi-major axis initial value of the second near-moon braking control be A2, then the new control variable is [Δt a , Δv x , Δv y , Δv z , A2], the target variable is the third near-month point at time t e3 、Inclination angle i e3 , right ascension of ascending node Ω e3 、perilunar argument ω e3 and perigee height H e3 The second partial derivative matrix is ​​established using formulas (1) to (7) in the above example.

[0135] The second partial derivative matrix is ​​iterated until the orbit parameter deviation corresponding to the third perigee is less than the convergence threshold. The convergence threshold can be a value preset based on experience, for example, the convergence threshold is that the deviation of all orbit parameters is less than 1, or infinitely close to 0, and the convergence threshold is not specifically limited here. Exemplarily, the iterative calculation process is shown in the following table.

[0136] Table 4. Convergence of orbital parameter deviations in the joint planning of far-moon maneuvers and the second near-moon braking

[0137]

[0138]

[0139] Step 303, when the orbital parameter deviation corresponding to the third perigee time is less than the convergence threshold, the control parameters of the far-moon maneuver and the control parameters of the second perigee braking are determined, and the control parameters of the third perigee braking are calculated, so that the satellite can enter the lunar large elliptical frozen orbit based on the control parameters of the far-moon maneuver, the control parameters of the second perigee braking and the control parameters of the third perigee braking.

[0140] In the embodiment of the present invention, when the orbit parameter deviation corresponding to the third perigee time is less than the convergence threshold, the control parameters of the far-moon maneuver and the control parameters of the second perigee braking are determined, and then the control parameters of the third perigee braking are calculated. The control parameters include the start-up time and the three components of the velocity increment at the start-up time.

[0141] After obtaining the control parameters of the far-moon maneuver and the second near-moon braking, the current orbit and the target orbit mainly have a difference in semi-major axis. The purpose of the third near-moon braking is to adjust the orbit semi-major axis to be consistent with the target orbit. Therefore, the control parameters of the third near-moon braking can be solved separately. The control parameters of the third near-moon braking can be obtained according to the vitality formula. The vitality formula is shown in the following formula:

[0142]

[0143] Among them, r is the distance from the moon center, a is the semi-major axis, and V is the speed on the orbit.

[0144] For example, assuming that the semi-major axis of the orbit before the third near-moon braking is A 3b , the target semi-major axis is Ae, then the velocity increment of the satellite at the perigee of the orbit before and after the third perigee braking can be calculated according to the vitality formula, the formula is as follows:

[0145]

[0146] Among them, ΔV3 is the speed increment of the third near-month braking, r p is the perigee height.

[0147] For example, continuing with the above example, the final calculation results of the control parameters for the far-moon maneuver, the second near-moon braking, and the third near-moon braking are shown in the following table:

[0148] Table 5. Results of joint planning control parameters for far-moon maneuvers and near-moon braking

[0149]

[0150] In the table, v m is the modulus of the velocity. At this time, the pulse of the third near-moon braking is small, so the pulse of the third near-moon braking has little effect on the orbit perigee angle and perigee height, and the pulse of the third near-moon braking can be omitted in the joint solution.

[0151] After obtaining the control parameters of the far-moon maneuver, the control parameters of the second near-moon braking and the control parameters of the third near-moon braking, the satellite can operate based on the control parameters of the far-moon maneuver, the control parameters of the second near-moon braking and the control parameters of the third near-moon braking to enter a lunar large elliptical frozen orbit.

[0152] In a possible implementation, if the orbital parameter deviation corresponding to the third perigee cannot be less than the convergence threshold (mainly when the accuracy requirement is high, the perigee angle and perigee height caused by the limited thrust of the orbit control engine cannot converge), then the third partial derivative matrix between the pulse initial value of the far-moon maneuver, the target semi-major axis initial value of the second perigee braking, the target semi-major axis initial value of the third perigee braking and the orbital parameter deviation corresponding to the fourth perigee time is established. Among them, the target semi-major axis initial value of the third perigee braking is the target frozen orbit semi-major axis. The orbital parameters corresponding to the fourth perigee time include the fourth perigee time, semi-major axis, inclination, right ascension of ascending node, perigee height and perigee angle. Exemplarily, let the target semi-major axis initial value of the second perigee braking control be A2, and the target semi-major axis initial value of the third perigee braking control be A3, then the new control variable is [Δt a , Δv x , Δv y , Δv z , A2, A3], the fourth near-moon time of the target variable is t e4 , semi-major axis A e4 、Inclination angle i e4 , right ascension of ascending node Ω e4 、perilunar argument ω e4 and perigee height H e4 Then, the third partial derivative matrix is ​​established using formulas (1) to (7) in the above example. It can be understood that at this time, the pulse of the third near-moon braking is relatively large. Due to the use of limited thrust control, it will have a certain impact on the orbital perigee angle and perigee height. Therefore, the pulse of the third near-moon braking needs to be considered in the joint solution.

[0153] Iterate based on the third partial derivative matrix until the orbit parameter deviation corresponding to the fourth perigee is less than the convergence threshold, and determine the control parameters of the far-moon maneuver, the second perigee braking, and the third perigee braking. The satellite operates based on the control parameters of the far-moon maneuver, the second perigee braking, the third perigee braking, and the fourth perigee braking to enter the lunar large elliptical frozen orbit.

[0154] In the above example, the final control time of the far-moon maneuver is 2024-03-25T10:48:52, which is about 3 hours away from the time when the target orbit passes the southernmost latitude circle by searching for the ascending orbit (2024-03-25T07:34:56), which is consistent with expectations. In the process of using the particle swarm algorithm to calculate the initial solution of the far-moon maneuver pulse, the Y component of the optimal solution from the third generation is 100, indicating that the range of the Y-direction velocity increment is small, but it is sufficient to provide an initial value for the subsequent precise solution calculation. The convergence speed of the precise solution of the far-moon maneuver alone is slow (15 times), and the convergence speed of the joint solution with the subsequent near-moon braking is fast (4 times). This is because when the far-moon maneuver is accurately solved, the orbital parameters of the target orbit are directly used as the perigee parameters after the far-moon maneuver. In fact, the two are separated by one circle in time, and there is a certain deviation in the orbital elements. This phenomenon further shows that the accuracy requirement for the precise solution of the far-moon maneuver alone is not high.

[0155] In the embodiment of the present invention, the performance index function is first constructed by calculating the time range of the far-moon maneuver, and then the particle swarm algorithm is used to find the optimal solution to obtain the pulse initial value and terminal parameter deviation of the far-moon maneuver. Then, when the terminal parameter deviation is too large, the Newton iteration method is used to accurately solve the far-moon maneuver pulse, so that the terminal parameter deviation is reduced, and the pulse initial value of the far-moon maneuver is updated. Finally, the pulse initial value of the far-moon maneuver is used as the initial value of the joint planning of the far-moon maneuver and the near-moon braking, and the joint planning exact solution is calculated to obtain the control parameters of the satellite. In this scheme, the performance index is constructed by weighted summation of the terminal parameter deviation, and the control variable is used as the optimization variable, so that the problem of solving the strong nonlinear equation group is converted into an optimization problem, and the control parameters for entering the large elliptical frozen orbit of the moon are accurately calculated. In addition, the present invention essentially targets all elements of the orbital parameters, and is therefore also applicable to other situations that require full-element targeting control (such as: initial design of low-energy entry of the target mission orbit, lunar orbit rendezvous and docking, and emergency orbit reconstruction of lunar capture anomalies).

[0156] Based on the same technical concept, Figure 4 A schematic diagram of a control structure for entering a lunar large elliptical frozen orbit provided by an embodiment of the present invention is exemplarily shown. The device can execute the process of a control method for entering a lunar large elliptical frozen orbit.

[0157] like Figure 4 As shown, the device specifically includes:

[0158] The acquisition module 410 is used to optimize and solve the far-moon maneuver pulse of the satellite based on the orbit of the satellite after the first near-moon braking according to the particle swarm algorithm to obtain the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value, wherein the pulse initial value of the far-moon maneuver represents the initial value of the satellite's velocity increment at the initial value of the far-moon maneuver;

[0159] Processing module 420 is used to iterate the orbit parameter deviation corresponding to the third perigee time based on the Newton iteration method according to the pulse initial value of the far-moon maneuver and the target semi-major axis initial value of the second perigee braking when the terminal parameter deviation corresponding to the pulse initial value is less than a threshold value;

[0160] When the orbital parameter deviation corresponding to the third perigee time is less than the convergence threshold, the control parameters of the far-moon maneuver and the control parameters of the second perigee braking are determined, and the control parameters of the third perigee braking are calculated, so that the satellite can enter the lunar large elliptical frozen orbit based on the control parameters of the far-moon maneuver, the control parameters of the second perigee braking and the control parameters of the third perigee braking.

[0161] Optionally, the processing module 420 is specifically used for:

[0162] Determine the range of far-moon maneuvering time based on the time when the satellite passes the southernmost or northernmost end of the lunar large elliptical frozen orbit when descending / ascending;

[0163] According to the value range of the initial value of the speed increment of the far-moon maneuver and the value range of the far-moon maneuver time, the far-moon maneuver pulse is optimized and solved based on the particle swarm algorithm. When the performance index of the far-moon maneuver is minimized, the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value are obtained; the performance index of the far-moon maneuver is obtained by weighted summation of the terminal parameter deviation based on the far-moon maneuver.

[0164] Optionally, the processing module 420 is further configured to:

[0165] When the terminal parameter deviation corresponding to the pulse initial value is greater than a threshold, the terminal parameter deviation is iterated based on the Newton iteration method according to the pulse initial value of the far-moon maneuver;

[0166] When the terminal parameter deviation is less than the threshold, the first pulse value of the corresponding far-moon maneuver is determined as the initial pulse value of the far-moon maneuver.

[0167] Optionally, the processing module 420 is specifically used for:

[0168] Establishing a first partial derivative matrix between the initial pulse value of the far-moon maneuver and the terminal parameter deviation;

[0169] Iteration is performed based on the first partial derivative matrix until the terminal parameter deviation is less than a threshold.

[0170] Optionally, the processing module 420 is specifically used for:

[0171] Establishing a second partial derivative matrix between the initial pulse value of the far-moon maneuver, the initial target semi-major axis value of the second near-moon braking, and the orbital parameter deviation corresponding to the third perigee time;

[0172] Iteration is performed based on the second partial derivative matrix until the orbit parameter deviation corresponding to the third perigee is less than a convergence threshold.

[0173] Optionally, the processing module 420 is further configured to:

[0174] If the orbital parameter deviation corresponding to the third perigee time cannot be less than the convergence threshold, a third partial derivative matrix is ​​established between the pulse initial value of the far-moon maneuver, the target semi-major axis initial value of the second perigee braking, the target semi-major axis initial value of the third perigee braking and the orbital parameter deviation corresponding to the fourth perigee time;

[0175] Iteration is performed based on the third partial derivative matrix until the orbital parameter deviation corresponding to the fourth perigee moment is less than a convergence threshold, thereby determining control parameters of the far-moon maneuver, control parameters of the second perigee braking, and control parameters of the third perigee braking.

[0176] Based on the same technical concept, an embodiment of the present invention further provides a computer device, including:

[0177] A memory for storing program instructions;

[0178] The processor is used to call the program instructions stored in the memory and execute the control method for entering the lunar large elliptical freezing orbit according to the obtained program.

[0179] Based on the same technical concept, an embodiment of the present invention also provides a computer-readable storage medium, which stores computer-executable instructions, and the computer-executable instructions are used to enable a computer to execute the above-mentioned control method for entering a lunar large elliptical frozen orbit.

[0180] Based on the same technical concept, an embodiment of the present invention further provides a computer program product, characterized in that the computer program product includes an executable program, and the executable program is executed by a processor to implement the above-mentioned control method for entering a lunar large elliptical frozen orbit.

[0181] Those skilled in the art will appreciate that the embodiments of the present application may be provided as methods, systems, or computer program products. Therefore, the present application may adopt the form of a complete hardware embodiment, a complete software embodiment, or an embodiment in combination with software and hardware. Moreover, the present application may adopt the form of a computer program product implemented in one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) that include computer-usable program code.

[0182] The present application is described with reference to the flowcharts and / or block diagrams of the methods, devices (systems), and computer program products according to the present application. It should be understood that each process and / or box in the flowchart and / or block diagram, as well as the combination of the processes and / or boxes in the flowchart and / or block diagram, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing device to produce a machine, so that the instructions executed by the processor of the computer or other programmable data processing device generate instructions for implementing the processes in the flowchart and / or block diagram. Figure 1 A process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in a block or multiple blocks.

[0183] These computer program instructions may also be stored in a computer-readable memory capable of directing a computer or other programmable data processing device to operate in a specific manner, so that the instructions stored in the computer-readable memory produce an article of manufacture comprising an instruction device, which implements the process Figure 1 A process or multiple processes and / or boxes Figure 1 A function specified in one or more boxes.

[0184] These computer program instructions can also be loaded onto a computer or other programmable data processing device so that a series of operating steps are executed on the computer or other programmable device to produce a computer-implemented process, thereby providing instructions for implementing the process. Figure 1 A process or multiple processes and / or boxes Figure 1 The steps for the functions specified in one or more boxes.

[0185] Obviously, those skilled in the art can make various changes and modifications to the present application without departing from the spirit and scope of the present application. Thus, if these modifications and variations of the present application fall within the scope of the claims of the present application and their equivalents, the present application is also intended to include these modifications and variations.

Claims

1. A control method for entering a lunar large elliptical freezing orbit, characterized in that: include: Based on the orbit of the satellite after the first near-moon braking, the far-moon maneuver pulse of the satellite is optimized and solved according to the particle swarm algorithm to obtain the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value, wherein the pulse initial value of the far-moon maneuver represents the initial value of the satellite's velocity increment at the initial value of the far-moon maneuver; When the terminal parameter deviation corresponding to the pulse initial value is less than a threshold, the orbit parameter deviation corresponding to the third perigee time is iterated based on the Newton iteration method according to the pulse initial value of the far-moon maneuver and the target semi-major axis initial value of the second perigee braking; When the orbital parameter deviation corresponding to the third perigee time is less than the convergence threshold, the control parameters of the far-moon maneuver and the control parameters of the second perigee braking are determined, and the control parameters of the third perigee braking are calculated, so that the satellite can enter the lunar large elliptical frozen orbit based on the control parameters of the far-moon maneuver, the control parameters of the second perigee braking and the control parameters of the third perigee braking.

2. The method according to claim 1, characterized in that Based on the orbit of the satellite after the first near-moon braking, the far-moon maneuver pulse of the satellite is optimized and solved according to the particle swarm algorithm to obtain the initial pulse value of the far-moon maneuver and the terminal parameter deviation corresponding to the initial pulse value, including: Determine the range of far-moon maneuvering time based on the time when the satellite passes the southernmost or northernmost end of the lunar large elliptical frozen orbit when descending / ascending; According to the value range of the initial value of the speed increment of the far-moon maneuver and the value range of the far-moon maneuver time, the far-moon maneuver pulse is optimized and solved based on the particle swarm algorithm. When the performance index of the far-moon maneuver is minimized, the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value are obtained; the performance index of the far-moon maneuver is obtained by weighted summation of the terminal parameter deviation based on the far-moon maneuver.

3. The method according to claim 1, characterized in that After obtaining the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value, the method further includes: When the terminal parameter deviation corresponding to the pulse initial value is greater than a threshold, the terminal parameter deviation is iterated based on the Newton iteration method according to the pulse initial value of the far-moon maneuver; When the terminal parameter deviation is less than the threshold, the first pulse value of the corresponding far-moon maneuver is determined as the initial pulse value of the far-moon maneuver.

4. The method according to claim 3, characterized in that According to the initial pulse value of the far-moon maneuver, the terminal parameter deviation is iterated based on the Newton iteration method, including: Establishing a first partial derivative matrix between the initial pulse value of the far-moon maneuver and the terminal parameter deviation; Iteration is performed based on the first partial derivative matrix until the terminal parameter deviation is less than a threshold.

5. The method according to claim 1, characterized in that According to the initial pulse value of the far-moon maneuver and the initial target semi-major axis value of the second near-moon braking, the orbit parameter deviation corresponding to the third perigee time is iterated based on the Newton iteration method, including: Establishing a second partial derivative matrix between the initial pulse value of the far-moon maneuver, the initial target semi-major axis value of the second near-moon braking, and the orbital parameter deviation corresponding to the third perigee time; Iteration is performed based on the second partial derivative matrix until the orbit parameter deviation corresponding to the third perigee is less than a convergence threshold.

6. The method according to any one of claims 1 to 5, characterized in that: After iterating the orbit parameter deviation corresponding to the third perigee time based on the Newton iteration method, the method further includes: If the orbital parameter deviation corresponding to the third perigee time cannot be less than the convergence threshold, a third partial derivative matrix is ​​established between the pulse initial value of the far-moon maneuver, the target semi-major axis initial value of the second perigee braking, the target semi-major axis initial value of the third perigee braking and the orbital parameter deviation corresponding to the fourth perigee time; Iteration is performed based on the third partial derivative matrix until the orbital parameter deviation corresponding to the fourth perigee moment is less than a convergence threshold, thereby determining control parameters of the far-moon maneuver, control parameters of the second perigee braking, and control parameters of the third perigee braking.

7. A control device for entering a lunar large elliptical freezing orbit, characterized in that: include: An acquisition module is used to optimize and solve the far-moon maneuver pulse of the satellite based on the orbit of the satellite after the first near-moon braking according to the particle swarm algorithm, and obtain the pulse initial value of the far-moon maneuver and the terminal parameter deviation corresponding to the pulse initial value, wherein the pulse initial value of the far-moon maneuver represents the initial value of the satellite's velocity increment at the initial value of the far-moon maneuver moment; A processing module, configured to iterate the orbit parameter deviation corresponding to the third perigee time based on the Newton iteration method according to the pulse initial value of the far-moon maneuver and the target semi-major axis initial value of the second perigee braking when the terminal parameter deviation corresponding to the pulse initial value is less than a threshold value; When the orbital parameter deviation corresponding to the third perigee time is less than the convergence threshold, the control parameters of the far-moon maneuver and the control parameters of the second perigee braking are determined, and the control parameters of the third perigee braking are calculated, so that the satellite can enter the lunar large elliptical frozen orbit based on the control parameters of the far-moon maneuver, the control parameters of the second perigee braking and the control parameters of the third perigee braking.

8. A computer device, characterized in that: include: A memory for storing program instructions; A processor, configured to call the program instructions stored in the memory, and execute the method according to any one of claims 1 to 6 according to the obtained program.

9. A computer-readable storage medium, characterized in that: The computer-readable storage medium stores computer-executable instructions, and the computer-executable instructions are used to enable a computer to execute the method according to any one of claims 1 to 6.

10. A computer program product, characterized in that The computer program product comprises an executable program, which is executed by a processor to implement the method according to any one of claims 1 to 6.

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