A control method and apparatus for entering a highly elliptical frozen orbit on the moon

Abstract

This invention discloses a control method and apparatus for entering a highly elliptical frozen orbit around the moon, comprising: based on the satellite's orbit after the first lunar braking, optimizing the apolunar maneuver pulse using a particle swarm optimization algorithm to obtain the initial value of the apolunar maneuver pulse. The initial value of the apolunar maneuver pulse represents the initial value of the satellite's velocity increment at the initial value of the apolunar maneuver. When the deviation of the terminal parameters corresponding to the initial value of the pulse is less than a threshold, based on the initial value of the apolunar maneuver pulse and the initial value of the target semi-major axis of the second lunar braking, jointly planning the apolunar maneuver and the second lunar braking using the Newton-Raphson iteration method, until the orbital parameter deviation at the third perihelion is less than the convergence threshold, determining the control parameters of the apolunar maneuver and the second lunar braking, and calculating the control parameters of the third lunar braking, thereby achieving accurate calculation of the control parameters for entering a highly elliptical frozen orbit around the moon.

Description

Technical Field

[0001] This invention relates to the field of aerospace telemetry and control, and in particular to a control method and apparatus for entering a highly elliptical frozen orbit on the moon. Background Technology

[0002] Lunar satellites exist in highly elliptical frozen orbits due to the north-south asymmetry of their gravitational field. A highly elliptical frozen orbit is a stable orbit, with its eccentricity and perigee angle remaining within a certain range. Because frozen orbits maintain a stable state with minimal or no change over long periods, almost no propellant is needed for orbital maintenance, significantly saving fuel consumption. Furthermore, utilizing the arch-shaped orientation characteristics of the highly elliptical frozen orbit allows for effective tracking and communication at specific locations or in specific states.

[0003] For a probe originating from Earth, entering a highly elliptical frozen orbit around the Moon typically involves four major orbital control maneuvers. After the first lunar braking maneuver, the probe enters the capture orbit segment. From the capture orbit segment to entering the target frozen orbit, three more control maneuvers are required: the apolunar maneuver, the second lunar braking maneuver, and the third lunar braking maneuver. When designing the frozen orbit entry control strategy, it is crucial to strictly adhere to the designed control parameters, including orbital epoch time, semi-major axis, eccentricity, inclination, right ascension of the ascending node, and argument of the perigee. Furthermore, the capture orbit has a large semi-major axis, and its orbital elements are significantly affected by Earth's gravitational perturbations, making analytical solutions for the frozen orbit entry pulse difficult.

[0004] Therefore, accurately calculating the control parameters for entering the frozen track is a problem that urgently needs to be solved. Summary of the Invention

[0005] This invention provides a control method for entering a highly elliptical frozen orbit around the moon. A particle swarm optimization algorithm is used to obtain the initial value of the apogee maneuver pulse. After obtaining the initial value, Newton's iteration method is used to further solve for the apogee maneuver and the second lunar braking, and finally, the control parameters for the third lunar braking are calculated. This yields the control parameters for the apogee maneuver, the second lunar braking, and the third lunar braking, enabling accurate calculation of the control parameters for entering a highly elliptical frozen orbit around the moon.

[0006] In a first aspect, embodiments of the present invention provide a control method for entering a highly elliptical frozen orbit of the moon, comprising:

[0007] Based on the satellite's orbit after its first lunar braking, the satellite's lunar maneuver pulses are optimized using the particle swarm optimization algorithm to obtain the initial value of the lunar maneuver pulses and the corresponding terminal parameter deviations. The initial value of the lunar maneuver pulses represents the initial value of the satellite's velocity increment at the moment of the lunar maneuver.

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

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

[0010] In the above technical solution, after the satellite performs its first lunar braking maneuver, it enters the lunar capture orbit. Then, it needs to perform apolunar maneuvers, a second lunar braking maneuver, and a third lunar braking maneuver to enter the highly elliptical frozen lunar orbit. Therefore, it is necessary to calculate the control parameters for the apolunar maneuver, the second lunar braking maneuver, and the third lunar braking maneuver to ensure the satellite enters the highly elliptical frozen lunar orbit based on these parameters. Therefore, the apolunar maneuver pulse is first optimized based on the capture orbit using a particle swarm optimization algorithm to obtain the initial value of the apolunar maneuver pulse and the corresponding terminal parameter deviations. The initial value of the apolunar maneuver pulse represents the initial value of the satellite's velocity increment at the moment of the apolunar maneuver, including the initial value at the moment of the apolunar maneuver and the three-directional components of the initial velocity increment in the orbit at that moment. The terminal parameters include the satellite's inclination at the lunar perigee, right ascension of the ascending node, lunar perigee altitude, and lunar perigee argument.

[0011] Then, when the deviation of the terminal parameters corresponding to the initial pulse value is less than the threshold, the orbital parameter deviation corresponding to the third perilune time is iterated based on the initial pulse value of the apolunar maneuver and the initial target semi-major axis value of the second perilune braking, using Newton's iteration method. The orbital parameters corresponding to the third perilune time include the third perilune time, inclination, right ascension of the ascending node, perilune altitude, and perilune argument. This process continues until the orbital parameter deviation corresponding to the third perilune time is less than the convergence threshold. At this point, the control parameters for the apolunar maneuver and the second perilune braking are determined, and the control parameters for the third perilune braking are calculated, thus achieving accurate calculation of the control parameters for entering the lunar highly elliptical frozen orbit.

[0012] Optionally, based on the satellite's orbit after its first lunar braking maneuver, the lunar maneuver pulses of the satellite are optimized using a particle swarm optimization algorithm to obtain the initial values ​​of the lunar maneuver pulses and the corresponding terminal parameter deviations, including:

[0013] The range of values ​​for the apolunar maneuver timing is determined based on the time when the satellite descends / ascends through the southernmost or northernmost point of the lunar highly elliptical frozen orbit.

[0014] Based on the range of initial values ​​for the velocity increment of the apogee maneuver and the range of values ​​for the apogee maneuver time, the apogee maneuver pulse is optimized using a particle swarm optimization algorithm. When the performance index of the apogee maneuver is minimized, the initial value of the apogee maneuver pulse and the corresponding terminal parameter deviation are obtained. The performance index of the apogee maneuver is obtained by weighted summation based on the terminal parameter deviation of the apogee maneuver.

[0015] In the above technical solution, after the satellite's first lunar braking maneuver, it enters a capture orbit. The initial time is then determined by using the moment when the satellite descends / ascends past the southernmost or northernmost point of the lunar highly elliptical frozen orbit as the initial time, thus defining the range of values ​​for the apogee maneuver timing. Finally, the apogee maneuver pulse is optimized using a particle swarm optimization algorithm. When the performance index of the apogee maneuver is minimized, the initial pulse value and the corresponding terminal parameter deviation are obtained. This transforms the problem of solving the strongly nonlinear equations governing the apogee maneuver into a global optimization problem. The particle swarm optimization algorithm is then used to optimize the initial pulse value for the apogee maneuver, resulting in a relatively good initial pulse value.

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

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

[0018] When the deviation of the terminal parameters is less than the threshold, the first pulse value of the corresponding lunar maneuver is determined and used as the initial pulse value of the lunar maneuver.

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

[0020] Optionally, based on the initial pulse value of the lunar maneuver, the terminal parameter deviation is iterated using the Newton-Raphson iteration method, including:

[0021] Establish the first partial derivative matrix between the initial pulse value of the lunar maneuver and the deviation of the terminal parameters;

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

[0023] Optionally, based on the initial pulse value of the far-lunar maneuver and the initial target semi-major axis value of the second lunar braking, the orbital parameter deviation corresponding to the third lunar perihelion moment is iterated using the Newton-Raphson iteration method, including:

[0024] Establish the second partial derivative matrix between the initial pulse value of the far-lunar maneuver, the initial target semi-major axis value of the second lunar braking, and the orbital parameter deviation corresponding to the time of the third lunar point;

[0025] The iteration is performed based on the second partial derivative matrix until the orbital parameter deviation corresponding to the third near-lunar point is less than the convergence threshold.

[0026] Optionally, after iterating the orbital parameter deviations corresponding to the third perigee time based on Newton's iteration method, the method further includes:

[0027] If the orbital parameter deviation corresponding to the third lunar perihelion time cannot be less than the convergence threshold, then establish the third partial derivative matrix between the initial pulse value of the far lunar maneuver, the initial target semi-major axis value of the second lunar braking, the initial target semi-major axis value of the third lunar braking, and the orbital parameter deviation corresponding to the fourth lunar perihelion time.

[0028] Based on the third partial derivative matrix, the iterative process continues until the orbital parameter deviation corresponding to the fourth lunar perihelion time is less than the convergence threshold, thereby determining the control parameters for the far-lunar maneuver, the second lunar braking, and the third lunar braking.

[0029] Secondly, embodiments of the present invention provide a control device for entering a highly elliptical frozen orbit of the moon, comprising:

[0030] The acquisition module is used to optimize and solve the satellite's far-moon maneuver pulse based on the satellite's orbit after the first lunar braking, according to the particle swarm optimization 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 moment.

[0031] The processing module is used to iterate the orbital parameter deviation corresponding to the third perigee moment based on Newton's iteration method, according to the initial value of the pulse for the far-lunar maneuver and the initial value of the target semi-major axis for the second near-lunar braking, when the terminal parameter deviation corresponding to the initial value of the pulse is less than the threshold.

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

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

[0034] The range of values ​​for the apolunar maneuver timing is determined based on the time when the satellite descends / ascends through the southernmost or northernmost point of the lunar highly elliptical frozen orbit.

[0035] Based on the range of initial values ​​for the velocity increment of the apogee maneuver and the range of values ​​for the apogee maneuver time, the apogee maneuver pulse is optimized using a particle swarm optimization algorithm. When the performance index of the apogee maneuver is minimized, the initial value of the apogee maneuver pulse and the corresponding terminal parameter deviation are obtained. The performance index of the apogee maneuver is obtained by weighted summation based on the terminal parameter deviation of the apogee maneuver.

[0036] Optionally, the processing module is further configured to:

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

[0038] When the deviation of the terminal parameters is less than the threshold, the first pulse value of the corresponding lunar maneuver is determined and used as the initial pulse value of the lunar maneuver.

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

[0040] Establish the first partial derivative matrix between the initial pulse value of the lunar maneuver and the deviation of the terminal parameters;

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

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

[0043] Establish the second partial derivative matrix between the initial pulse value of the far-lunar maneuver, the initial target semi-major axis value of the second lunar braking, and the orbital parameter deviation corresponding to the time of the third lunar point;

[0044] The iteration is performed based on the second partial derivative matrix until the orbital parameter deviation corresponding to the third near-lunar point is less than the convergence threshold.

[0045] Optionally, the processing module is further configured to:

[0046] If the orbital parameter deviation corresponding to the third lunar perihelion time cannot be less than the convergence threshold, then establish the third partial derivative matrix between the initial pulse value of the far lunar maneuver, the initial target semi-major axis value of the second lunar braking, the initial target semi-major axis value of the third lunar braking, and the orbital parameter deviation corresponding to the fourth lunar perihelion time.

[0047] Based on the third partial derivative matrix, the iterative process continues until the orbital parameter deviation corresponding to the fourth lunar perihelion time is less than the convergence threshold, thereby determining the control parameters for the far-lunar maneuver, the second lunar braking, and the third lunar braking.

[0048] Thirdly, embodiments of the present invention also provide a computer device, comprising:

[0049] Memory, used to store program instructions;

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

[0051] Fourthly, embodiments of the present invention also provide a computer-readable storage medium storing computer-executable instructions for causing a computer to execute the aforementioned control method for entering a highly elliptical frozen orbit of the moon.

[0052] Fifthly, embodiments of the present invention also provide a computer program product, the computer program product including an executable program, which is executed by a processor to perform the above-described control method for entering a highly elliptical frozen orbit of the moon. Attached Figure Description

[0053] To more clearly illustrate the technical solutions in the embodiments of the present invention, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

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

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

[0056] Figure 3 A flowchart illustrating a control method for entering a highly elliptical frozen orbit on the moon, provided in an embodiment of the present invention;

[0057] Figure 4This is a schematic diagram of the structure of a control device for entering a highly elliptical frozen orbit on the moon, provided as an embodiment of the present invention. Detailed Implementation

[0058] To make the objectives, technical solutions, and advantages of the present invention more apparent, the present invention will be further described in detail below with reference to the accompanying drawings. It is apparent that the embodiments described are only some, not all, of the present invention. All other embodiments derived by persons of ordinary skill in the art based on the embodiments of the present invention without creative effort are intended to fall within the scope of protection of the present invention.

[0059] The application scenarios described in this application are for the purpose of more clearly illustrating the technical solutions protected by the embodiments of this application, and do not constitute a limitation on the technical solutions provided by the embodiments of this application. Those skilled in the art will understand that with the emergence of new application scenarios, the technical solutions provided by the embodiments of this application are also applicable to similar technical problems. The terms "first" and "second" in the specification, claims, and accompanying drawings of this application are used to distinguish different objects, not to describe a specific order. In the description of this application, unless otherwise stated, "multiple" means two or more.

[0060] Before introducing the control method for entering a highly elliptical frozen orbit of the moon provided in the embodiments of this application, for ease of understanding, the terms and background technology involved in the embodiments of this application will be introduced below.

[0061] Lunar satellites exist in highly elliptical frozen orbits due to the north-south asymmetry of their gravitational fields. These orbits are stable, with their eccentricity and perigee argument remaining relatively constant. Because of this long-term stability, requiring minimal or no change, the amount of propellant needed for orbital maintenance is minimized, significantly reducing fuel consumption. Furthermore, the camber orientation of these orbits allows for effective tracking and communication at specific locations or in specific states. According to relevant theories, the perigee argument of a highly elliptical frozen orbit is around 90° or 270°, and a specific relationship exists between the orbital inclination and eccentricity. Due to mission constraints (such as relay communication duration, start and end times to service targets), the orbital period, phase, and right ascension of the ascending node must also meet specific requirements. The safety of the lunar perihelion altitude, the change in the lunar perihelion argument (reflecting the degree of deviation of the arch line), and the visible arc of the service target during long-term operation of the lunar highly elliptical frozen orbit are all highly sensitive to the initial parameters of the orbit. Therefore, the initial parameters of the lunar highly elliptical frozen orbit must be strictly designed. At the same time, when designing the entry control strategy for the lunar highly elliptical frozen orbit, the designed initial parameters must also be strictly targeted, including time, semi-major axis, eccentricity, inclination, right ascension of the ascending node, and lunar perihelion argument (the true anomaly f = 0 is the termination condition for integration).

[0062] For probes originating from Earth, entering a highly elliptical frozen orbit around the Moon typically involves four major orbital maneuvers. Since the spacecraft's orbit relative to the Moon is hyperbolic upon arrival, the first lunar orbit insertion (LOI1) primarily aims to establish the spacecraft in a highly elliptical lunar orbit. This maneuver is performed near the first perigee of the Moon, with the pulse direction opposite to the perigee velocity, thus decelerating the probe. Regarding the design of the first lunar orbit insertion target, two approaches can be adopted: one is to operate independently of subsequent maneuvers, focusing solely on lunar capture; the other is to plan it in conjunction with subsequent maneuvers to minimize the velocity increment consumed by subsequent maneuvers while meeting telemetry and control constraints before finally entering the target highly elliptical frozen lunar orbit.

[0063] After the first lunar braking maneuver, the probe enters the capture orbit phase. From the capture orbit phase to entering the target lunar highly elliptical frozen orbit, three control maneuvers are required: apolunar maneuvering, the second lunar braking maneuver, and the third lunar braking maneuver. The apolunar maneuvering maneuver is the most important of these three. Its control variables include the three-directional components of timing and velocity. The control targets are the inclination, right ascension of the ascending node, argument of the perilune, and altitude of the target lunar highly elliptical frozen orbit. This control is denoted by LAM (Lunar Apogee Maneuver). The second lunar braking maneuver (LOI2, Lunar OrbitInsert 2) is performed when the satellite is near the perilune again. The main function of this control is phasing. By adjusting the orbital period, the time when the probe next reaches the perilune is at the orbital epoch in the initial parameters of the lunar highly elliptical frozen orbit. This is implemented using tangential pulses. The third lunar orbit insertion (LOI3) was performed when the probe reached its third perigee. At this point, the probe's deviation from the target lunar elliptical frozen orbit was only the semi-major axis. Therefore, the goal of this control was to adjust the orbit's semi-major axis to match the target orbit, which was implemented using tangential pulses.

[0064] After the spacecraft enters the capture orbit, it is controlled by three pulses to enter a highly elliptical lunar frozen orbit where all orbital elements meet the target requirements. Because the three pulses, especially the first two, influence each other and jointly target the parameters of the highly elliptical lunar frozen orbit, it is necessary to jointly solve for the three pulse orbital control parameters when solving for the control parameters for entering the highly elliptical lunar frozen orbit. In these three pulses, the apogee maneuver controls four quantities, and at this time the semi-major axis of the orbit is relatively large, making the orbit less stable. Therefore, the core issue in frozen orbit entry control lies in the calculation of the apogee maneuver pulse.

[0065] The capture orbit has a large semi-major axis and is significantly affected by Earth's gravity. Considering only the effects on inclination, right ascension of the ascending node, argument of the perigee, and altitude of the perigee, there are two sets of solutions for the apolunar maneuver pulse. One set of solutions is obtained under conditions of relatively small changes in orbital parameters and can be solved analytically, but the semi-major axis is small and the velocity pulse magnitude is extremely large. The other set of solutions utilizes Earth's gravitational perturbation to change the orbital plane, resulting in a larger semi-major axis and a smaller velocity pulse after control; however, it is difficult to obtain initial values ​​analytically. Of these two solutions, the second is the desired solution, and the key problem to be solved is solving the strongly nonlinear equations.

[0066] The far-lunar maneuver is performed during the capture orbit phase, when the orbital period is relatively large, and the perturbation effect of Earth's gravity on orbital parameters is very significant. For example... Figure 1 As shown, Figure 1This diagram illustrates the variation of orbital parameters according to an embodiment of the present invention. The diagram shows the instantaneous changes in orbital elements for an orbit with a semi-major axis of 30,000 km. Among the changes in orbital elements shown in the diagram, the largest inclination perturbation is approximately 11.2°, the largest right ascension perturbation at the ascending node is 18.3°, the largest perturbation at the lunar perihelion is approximately 3.3°, and the largest perturbation at the lunar perihelion altitude is approximately 230 km. Analysis shows that if the semi-major axis increases further, the perturbations in orbital elements will be more significant. For example, for an orbit with a semi-major axis of 50,000 km, the largest inclination perturbation is approximately 35°, the largest perturbation at the ascending node is 180°, the largest perturbation at the lunar perihelion is approximately 52°, and the largest perturbation at the lunar perihelion altitude is approximately 2100 km. Due to the very drastic perturbations in orbital elements, analytical methods for calculating initial pulse values ​​are no longer suitable, as analytical methods are based on relatively small changes in orbital elements. Using Newton's iteration method or the MinPack method also requires high-quality initial pulse values; otherwise, extremely slow convergence or divergence may occur.

[0067] Newton's method, also known as the Newton-Raphson method, is a method proposed by Newton in the 17th century for approximating solutions to equations in the real and complex number domains. This method approximates the roots of the equation through repeated iterations, using Taylor series expansion and neglecting higher-order terms to linearize nonlinear equations, thus simplifying the solution process. Newton's method is not only applicable to solving the roots of single equations but can also be extended to solving differential and integral equations. Its basic idea is to approximate the antiderivative with the tangent line to the function and iteratively update the estimated value of the solution until the desired accuracy is achieved.

[0068] Particle Swarm Optimization (PSO) is an algorithm for finding the optimal solution for a flock of birds searching for food. Each "bird" represents a particle, and the "food" the flock seeks is the desired optimal solution. For a search involving m particles flying in a D-dimensional search space, the algorithm is defined as follows.

[0069] The coordinates of the i-th particle are:

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

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

[0072] The optimal positions experienced by all particles in the population are:

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

[0074]

[0075] In the formula: ω is the inertia weight, which plays a role in balancing the global and local search. A larger ω has a better global convergence ability, and vice versa. c1 and c2 are learning factors, which control the particle's ability to find the individual optimal position and the global optimal position, respectively. r1 and r2 are pseudo-random numbers that are distributed between [0,1].

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

[0077] The communication interface 220 is used for data transmission.

[0078] The processor 210 is the control center of the server 200, connecting various parts of the server 200 through various interfaces and routes. It performs various functions of the server 200 and processes data by running or executing software programs and / or modules stored in the memory 230, and by 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 may mainly include a program storage area and a data storage area. The program storage area may store the operating system, at least one application program required for a function, etc.; the data storage area may store data created according to business processing, etc. In addition, the memory 230 may include high-speed random access memory, and may also include non-volatile memory, such as at least one disk storage device, flash memory device, or other volatile solid-state storage device.

[0080] It should be noted that the above Figure 2 The structure shown is merely an example, and the embodiments of the present invention are not limited thereto.

[0081] Based on the above description Figure 3 An exemplary schematic diagram of a control method for entering a highly elliptical frozen orbit of the moon, provided by an embodiment of the present invention, is shown. This process can be executed by a control device for entering a highly elliptical frozen orbit of the moon.

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

[0083] Step 301: Based on the satellite's orbit after its first lunar braking, optimize the lunar maneuver pulse of the satellite using the particle swarm optimization algorithm to obtain the initial value of the lunar maneuver pulse and the corresponding terminal parameter deviation. The initial value of the lunar maneuver pulse represents the initial value of the satellite's velocity increment at the initial value of the lunar maneuver moment.

[0084] In this embodiment of the invention, after the satellite performs its first lunar braking maneuver, it enters the lunar capture orbit, and the parameters of the capture orbit can be obtained. Then, through apolunar maneuvers, a second lunar braking maneuver, and a third lunar braking maneuver, it enters a highly elliptical frozen lunar orbit. It is understood that the parameters of the target lunar highly elliptical frozen orbit are known. For example, the capture orbit parameters of the lunar center-equatorial inertial frame after the first lunar 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] Right ascension of the ascending node Ω0 = 276.0°

[0090] The near-lunar angle ω0 = 127.0°

[0091] The angle of approach to the point M0 = 0°

[0092] The parameters of the target's highly elliptical frozen lunar 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] Lunar perigee height Hp e =620.0km

[0096] Inclination angle i e =115.0°

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

[0098] near-lunar point argument ω e =88.0°

[0099] Plane near point angle M e =0°

[0100] First, after the first lunar braking maneuver, the satellite enters the capture orbit. Then, based on the time when the satellite descends / ascends past the southernmost or northernmost point of the target orbit, the range of values ​​for the apo-lunar maneuver time is determined. For example, the orbit after the first lunar braking is integrated until the satellite descends / ascends past the southernmost or northernmost point of the target orbit (the highest latitude; the southernmost point will be used for explanation later), and the time is recorded. The point with the larger lunar center distance is selected as the initial time, denoted as T0. Therefore, the time range for the apo-lunar maneuver is [T0, T0+6h]. For example, if the corresponding ascending orbit passes through -65°S latitude at 2024-03-25T07:34:56, then the time range for the apo-lunar maneuver is [2024-03-25T07:34:56, 2024-03-25T13:34:56].

[0101] Then, determine the range of initial values ​​for the velocity increments of the apogee maneuver. The range of initial values ​​for the velocity increments of the apogee maneuver is a preset range based on experience, for example: ΔV x : Lunar maneuver X-axis velocity increment, with a value range of [-100m / s to 100m / s]; ΔV Y : The Y-axis velocity increment of the lunar maneuver, with a value range of [-100m / s 100m / s]; ΔV Z The Z-axis velocity increment of the lunar maneuver has a range of [-100m / s 100m / s]. The range of the initial pulse value is not specifically limited here.

[0102] Based on the range of initial values ​​for the velocity increment during the apogee maneuver and the range of values ​​for the apogee maneuver timing, and using a particle swarm optimization algorithm to optimize the apogee maneuver pulse, the initial pulse value and the corresponding terminal parameter deviation are obtained when the performance index of the apogee maneuver is minimized. The initial pulse value of the apogee maneuver represents the initial velocity increment of the satellite at the initial value of the apogee maneuver timing, including the initial value of the apogee maneuver timing and the three-directional components of the initial velocity increment of the satellite in orbit at the initial value of the apogee maneuver timing. The terminal parameters include the satellite's inclination at the lunar perigee, right ascension of the ascending node, lunar perigee altitude, and lunar perigee argument. The performance index of the apogee maneuver is obtained by weighted summation of the terminal parameter deviations. For example, the orbit is integrated uncontrolled to the next lunar perigee, and the deviations of the inclination i, right ascension Ω, lunar perigee argument ω, and lunar perigee altitude H from the target values ​​are calculated. The deviation results are denoted as Di, DΩ, Dω, and DH. Then, based on the initial orbit, the time of the far-lunar maneuver, and the velocity increment, the deviations of the inclination, right ascension of the ascending node, argument of the perilune, and altitude from the target values ​​at the next perilune are calculated as di, dΩ, dω, and dH, respectively. Finally, the performance index J is obtained as follows:

[0103]

[0104] The maximum value of the above performance index is 0. Without considering the lunar maneuver pulse, the performance index is -4. Regarding the construction of the performance index, considering the magnitude of the initial deviation of the terminal parameters, different weighting coefficients can be assigned to each terminal parameter based on the initial deviation. Terminal parameters with larger deviations have higher weights, and those with smaller deviations have lower weights.

[0105] For example, the process of optimizing the apogee maneuver pulse using the particle swarm optimization (PSO) algorithm includes: setting the population size to 100, generation number to 5, inertia weight to 1.0, cognitive coefficient to 1.5, and social coefficient to 1.5; using the three-directional components of the pulse's time deviation and velocity increment as optimization variables; restricting the range of optimization variables but not the velocity range; and then, based on the above-set parameters, performing optimization to obtain the result of the PSO algorithm solution.

[0106] Table 1. Results of the Particle Swarm Optimization Algorithm Solution

[0107]

[0108]

[0109] As shown in the table above, the third-generation particle swarm optimization algorithm yields the best results. It is selected as the initial pulse value for accurately calculating the far-lunar maneuver. The corresponding terminal parameter deviations are shown in the table below:

[0110] Table 2. Terminal parameter deviations in lunar maneuver results solved using the particle swarm optimization algorithm.

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

[0112] In one possible implementation, when the terminal parameter deviation corresponding to the initial pulse value is greater than a threshold, the terminal parameter deviation is iterated based on the initial pulse value of the apogee maneuver using Newton's iteration method. The threshold can be a preset value based on experience; for example, the perilune altitude threshold is set to 10 km, and the dip angle, right ascension of the ascending node, and argument of the perilune are set to 4.0°. No specific limit is imposed on the threshold here. Specifically, a first partial derivative matrix is ​​established between the initial pulse value of the apogee 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. For example, assume the precise correction amount for the apogee maneuver is [Δt]. a Δv x Δv y Δv z Then the detector at T0+Δt a At that moment, with [v x +Δv x , v y +Δvy , v z +Δv z Starting at a speed of [missing information], the terminal parameter at the perigee is the inclination angle. Right ascension of ascending node Near-month point argument and perilune altitude Represented as:

[0113]

[0114] Where f, g, h, and k are the tilt angles of the structure. Right ascension of ascending node Near-month point argument and perilune altitude With T0+Δt a Time and speed [v] x +Δv x , v y +Δv y , v z +Δv z The functional relationship between them. Then, perform a Taylor expansion on the left side of the above equation:

[0115]

[0116] Then, after rearranging the above formula, we get:

[0117]

[0118] Further analysis revealed:

[0119]

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

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

[0122] (2) In the control variable t a Apply a very small perturbation ε (e.g., ε = 1s) to the surface.

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

[0124]

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

[0126]

[0127] Similarly, by applying perturbations ε to the three velocity components, the complete partial derivative matrix can be obtained. This process is repeated several times until the terminal parameter deviation is less than a threshold.

[0128] For example, the convergence process of the iterative calculation is shown in the table below:

[0129] Table 3. Convergence of the iterative solution for the lunar maneuver.

[0130]

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

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

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

[0134] Specifically, when the terminal parameter deviation corresponding to the apolunar maneuver is less than a threshold, a second partial derivative matrix is ​​established between the initial pulse value of the apolunar maneuver, the initial target semi-major axis value of the second lunar braking, and the orbital parameter deviation corresponding to the third lunar perihelion time. The initial target semi-major axis value of the second lunar braking can be the semi-major axis of the target lunar elliptical frozen orbit, or it can be calculated based on the time difference (i.e., orbital period) between the second lunar braking time and the initial epoch of the target lunar elliptical frozen orbit (the third lunar braking time), and then the true target semi-major axis is obtained through iterative optimization. The specific method for obtaining the initial target semi-major axis value is not specified here. The orbital parameters corresponding to the third lunar perihelion time include the third lunar perihelion time, inclination, right ascension of the ascending node, lunar perihelion altitude, and lunar perihelion argument. For example, if the initial target semi-major axis value of the second lunar braking control is A2, then the new control variable is [Δt]. a Δv x Δv y Δv z [A2], the target variable is the third near-lunar point time t e3 Inclination angle i e3 Right ascension of ascending node Ω e3 near-lunar argument ω e3 and the height of the perigee 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 orbital parameter deviation at the third perigee is less than the convergence threshold. The convergence threshold can be a preset value based on experience, such as all orbital parameter deviations being less than 1, or infinitely approaching 0. No specific limitation is made to the convergence threshold here. An example iterative calculation process is shown in the table below.

[0136] Table 4. Convergence of Orbital Parameter Deviations in the Joint Planning of the Far-Moon Maneuver and the Second Near-Moon Braking

[0137]

[0138]

[0139] Step 303: When the orbital parameter deviation corresponding to the third perigee is less than the convergence threshold, determine the control parameters of the far-moon maneuver and the control parameters of the second perigee braking, and calculate the control parameters of the third perigee braking, so that the satellite can enter the lunar highly 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 this embodiment of the invention, when the orbital parameter deviation corresponding to the third lunar perihelion is less than the convergence threshold, the control parameters for the far-lunar maneuver and the second lunar braking are determined, and then the control parameters for the third lunar braking are calculated. The control parameters include the start-up time and three components of the velocity increment at the start-up time.

[0141] After obtaining the control parameters for the far-lunar maneuver and the second lunar braking, the main difference between the current orbit and the target orbit lies in their semi-major axis. The purpose of the third lunar braking is to adjust the orbit's semi-major axis to match the target orbit. Therefore, the control parameters for the third lunar braking can be calculated separately. The control parameters for the third lunar braking can be obtained using the vitality formula, as shown below:

[0142]

[0143] Where r is the distance from the lunar center, a is the semi-major axis, and V is the velocity on the orbit.

[0144] For example, suppose the semi-major axis of the orbit before the third lunar braking is A. 3b If the semi-major axis of the target is Ae, then the velocity increment of the satellite at the perigee of its orbit before and after the third lunar braking can be calculated using the vitality formula, as shown below:

[0145]

[0146] Where ΔV3 is the velocity increment of the third lunar braking, r p This is the altitude of the perigee.

[0147] For example, continuing from the previous example, the final calculation results of the control parameters for the far-lunar maneuver, the second lunar braking, and the third lunar braking are shown in the table below:

[0148] Table 5. Results of Control Parameters for Joint Planning of Lunar Maneuvering and Lunar Braking

[0149]

[0150] In the table, v m Let be the magnitude of the velocity. At this point, the pulse of the third lunar braking is relatively small, therefore the pulse of the third lunar braking has little impact on the lunar perihelion argument and lunar perihelion altitude, and the pulse of the third lunar braking can be omitted from the joint solution.

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

[0152] In one possible implementation, if the orbital parameter deviation corresponding to the third lunar perihelion cannot be less than the convergence threshold (mainly because the limited thrust of the orbit control engine causes the lunar perihelion argument and altitude to fail to converge when high accuracy is required), then a third partial derivative matrix is ​​established between the initial values ​​of the pulse for the far-lunar maneuver, the initial values ​​of the target semi-major axis for the second lunar braking, the initial values ​​of the target semi-major axis for the third lunar braking, and the orbital parameter deviation corresponding to the fourth lunar perihelion. The initial value of the target semi-major axis for the third lunar braking is the target frozen orbit semi-major axis. The orbital parameters corresponding to the fourth lunar perihelion include the fourth lunar perihelion time, semi-major axis, inclination, right ascension of the ascending node, lunar perihelion altitude, and lunar perihelion argument. For example, if the initial value of the target semi-major axis for the second lunar braking control is A2, and the initial value of the target semi-major axis for the third lunar braking control is A3, then the new control variable is [Δt]. a Δv x Δv y Δv z [A2, A3], the target variable is the fourth near-lunar point time t e4 Semi-major axis A e4 Inclination angle i e4 Right ascension of ascending node Ω e4 near-lunar argument ω e4 and the height of the perigee 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 lunar braking is relatively large. Due to the use of finite thrust control, it will have a certain impact on the lunar perilpoint argument and lunar perilpoint height. Therefore, the pulse of the third lunar braking needs to be considered when solving the problem together.

[0153] Iterations were performed based on the third partial derivative matrix until the orbital parameter deviation at the fourth perigee was less than the convergence threshold, thus determining the control parameters for the apolunar maneuver, the second perigee braking, and the third perigee braking. The satellite then operated based on these parameters to enter a highly elliptical frozen lunar orbit.

[0154] In the above example, the final control timing of the apolunar maneuver was 2024-03-25T10:48:52, which deviated from the time of the ascent through the southernmost latitude circle of the target orbit (2024-03-25T07:34:56) by about 3 hours, consistent with expectations. During the calculation of the initial solution for the apolunar maneuver pulse using the particle swarm optimization algorithm, the Y-axis component of the optimal solution was 100 from the 3rd generation onwards, indicating that the range of the Y-axis velocity increment was relatively small, but sufficient to provide initial values ​​for subsequent accurate solutions. The convergence speed of the apolunar maneuver's individual accurate solution was slow (15 iterations), while the convergence speed of the joint solution with the subsequent lunar braking was faster (4 iterations). This is because when calculating the apolunar maneuver's accurate solution, the orbital parameters of the target orbit were directly used as the perilunar point parameters after the apolunar maneuver. In reality, the two are separated by one orbit in time, resulting in a certain deviation in orbital elements. This phenomenon further illustrates that the accuracy requirement for the apolunar maneuver's individual accurate solution is not high.

[0155] In this embodiment of the invention, a performance index function is first constructed by calculating the time range of the apogee maneuver. Then, a particle swarm optimization algorithm is used to find the optimal solution, obtaining the initial pulse value and terminal parameter deviation of the apogee maneuver. Next, when the terminal parameter deviation is too large, a Newton-Raphson iteration method is used to accurately solve the apogee maneuver pulse, reducing the terminal parameter deviation and updating the initial pulse value of the apogee maneuver. Finally, the initial pulse value of the apogee maneuver is used as the initial value for the joint planning of the apogee maneuver and the lunar braking, and the accurate solution of the joint planning is calculated to obtain the satellite's control parameters. This scheme constructs a performance index by weighted summation of terminal parameter deviations and uses control variables as optimization variables, transforming the problem of solving a strongly nonlinear equation system into an optimization problem, thus achieving accurate calculation of the control parameters for entering the lunar highly elliptical frozen orbit. Furthermore, this invention essentially targets all elements of the orbital parameters, therefore it is also applicable to other situations requiring full-element targeting control (such as: initial design for low-energy entry into the target mission orbit, lunar orbit rendezvous and docking, and emergency orbit reconstruction for lunar capture anomalies).

[0156] Based on the same technological concept Figure 4 An exemplary schematic diagram of a control structure for entering a highly elliptical frozen orbit of the moon is shown in an embodiment of the present invention. This device can execute the process of a control method for entering a highly elliptical frozen orbit of the moon.

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

[0158] The acquisition module 410 is used to optimize the lunar maneuver pulse of the satellite based on the orbit after the first lunar braking of the satellite, and to obtain the initial value of the lunar maneuver pulse and the terminal parameter deviation corresponding to the initial value of the pulse. The initial value of the lunar maneuver pulse represents the initial value of the satellite's velocity increment at the initial value of the lunar maneuver moment.

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

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

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

[0162] The range of values ​​for the apolunar maneuver timing is determined based on the time when the satellite descends / ascends through the southernmost or northernmost point of the lunar highly elliptical frozen orbit.

[0163] Based on the range of initial values ​​for the velocity increment of the apogee maneuver and the range of values ​​for the apogee maneuver time, the apogee maneuver pulse is optimized using a particle swarm optimization algorithm. When the performance index of the apogee maneuver is minimized, the initial value of the apogee maneuver pulse and the corresponding terminal parameter deviation are obtained. The performance index of the apogee maneuver is obtained by weighted summation based on the terminal parameter deviation of the apogee maneuver.

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

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

[0166] When the deviation of the terminal parameters is less than the threshold, the first pulse value of the corresponding lunar maneuver is determined and used as the initial pulse value of the lunar maneuver.

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

[0168] Establish the first partial derivative matrix between the initial pulse value of the lunar maneuver and the deviation of the terminal parameters;

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

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

[0171] Establish the second partial derivative matrix between the initial pulse value of the far-lunar maneuver, the initial target semi-major axis value of the second lunar braking, and the orbital parameter deviation corresponding to the time of the third lunar point;

[0172] The iteration is performed based on the second partial derivative matrix until the orbital parameter deviation corresponding to the third near-lunar point is less than the convergence threshold.

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

[0174] If the orbital parameter deviation corresponding to the third lunar perihelion time cannot be less than the convergence threshold, then establish the third partial derivative matrix between the initial pulse value of the far lunar maneuver, the initial target semi-major axis value of the second lunar braking, the initial target semi-major axis value of the third lunar braking, and the orbital parameter deviation corresponding to the fourth lunar perihelion time.

[0175] Based on the third partial derivative matrix, the iterative process continues until the orbital parameter deviation corresponding to the fourth lunar perihelion time is less than the convergence threshold, thereby determining the control parameters for the far-lunar maneuver, the second lunar braking, and the third lunar braking.

[0176] Based on the same technical concept, embodiments of the present invention also provide a computer device, including:

[0177] Memory, used to store program instructions;

[0178] The processor is used to call the program instructions stored in the memory and execute the above-mentioned control method for entering the lunar highly elliptical frozen orbit according to the obtained program.

[0179] Based on the same technical concept, embodiments of the present invention also provide a computer-readable storage medium storing computer-executable instructions for causing a computer to execute the above-described control method for entering a highly elliptical frozen orbit of the moon.

[0180] Based on the same technical concept, this invention also provides a computer program product, characterized in that the computer program product includes an executable program, which is executed by a processor to perform the above-described control method for entering a highly elliptical frozen orbit of the moon.

[0181] Those skilled in the art will understand that embodiments of this application can be provided as methods, systems, or computer program products. Therefore, this application can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, this application can take the form of a computer program product embodied on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code.

[0182] This application is described with reference to flowchart illustrations and / or block diagrams of methods, apparatus (systems), and computer program products according to this application. It should be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, special-purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, generate instructions for implementing the flowchart illustrations. Figure 1 a process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in one or more boxes.

[0183] These computer program instructions may also be stored in a computer readable memory that can direct a computer or other programmable data processing device to work 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 The 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 operational steps are executed on the computer or other programmable device to produce a computer-implemented process, thereby providing the instructions executed on the computer or other programmable device for implementing the process. Figure 1 a process or multiple processes and / or boxes Figure 1 A step that specifies a function in one or more boxes.

[0185] Obviously, those skilled in the art can make various modifications and variations to this application without departing from the spirit and scope of this application. Therefore, if such modifications and variations fall within the scope of the claims of this application and their equivalents, this application also intends to include such modifications and variations.

Claims

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

2. The method as described in claim 1, characterized in that, Based on the satellite's orbit after its first lunar braking maneuver, the lunar maneuver pulses of the satellite are optimized using the particle swarm optimization algorithm to obtain the initial values ​​of the lunar maneuver pulses and the corresponding terminal parameter deviations, including: The range of values ​​for the apolunar maneuver timing is determined based on the time when the satellite descends / ascends through the southernmost or northernmost point of the lunar highly elliptical frozen orbit. Based on the range of initial values ​​for the velocity increment of the apogee maneuver and the range of values ​​for the apogee maneuver time, the apogee maneuver pulse is optimized using a particle swarm optimization algorithm. When the performance index of the apogee maneuver is minimized, the initial value of the apogee maneuver pulse and the corresponding terminal parameter deviation are obtained. The performance index of the apogee maneuver is obtained by weighted summation based on the terminal parameter deviation of the apogee maneuver.

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

4. The method as described in claim 3, characterized in that, Based on the initial pulse value of the lunar maneuver, the terminal parameter deviation is iterated using the Newton-Raphson iteration method, including: Establish the first partial derivative matrix between the initial pulse value of the lunar maneuver and the deviation of the terminal parameters; The iteration is performed based on the first partial derivative matrix until the deviation of the terminal parameter is less than the threshold.

5. The method as described in claim 1, characterized in that, Based on the initial pulse value of the far-lunar maneuver and the initial target semi-major axis value of the second lunar braking, the orbital parameter deviations corresponding to the third lunar perihelion time are iterated using the Newton-Raphson iteration method, including: Establish the second partial derivative matrix between the initial pulse value of the far-lunar maneuver, the initial target semi-major axis value of the second lunar braking, and the orbital parameter deviation corresponding to the time of the third lunar point; The iteration is performed based on the second partial derivative matrix until the orbital parameter deviation corresponding to the third near-lunar point is less than the convergence threshold.

6. The method according to any one of claims 1-5, characterized in that, After iterating the orbital parameter deviations corresponding to the third perigee time using Newton's iteration method, the method further includes: If the orbital parameter deviation corresponding to the third lunar perihelion time cannot be less than the convergence threshold, then establish the third partial derivative matrix between the initial pulse value of the far lunar maneuver, the initial target semi-major axis value of the second lunar braking, the initial target semi-major axis value of the third lunar braking, and the orbital parameter deviation corresponding to the fourth lunar perihelion time. Based on the third partial derivative matrix, the iterative process continues until the orbital parameter deviation corresponding to the fourth lunar perihelion time is less than the convergence threshold, thereby determining the control parameters for the far-lunar maneuver, the second lunar braking, and the third lunar braking.

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

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

9. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores computer-executable instructions for causing a computer to perform the method according to any one of claims 1 to 6.

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

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