Method and device for orbit control of a circumlunar spacecraft

Abstract

The application discloses a method and device for orbit transfer control of a lunar spacecraft, and realizes orbit transfer of the spacecraft into a full-parameter target orbit. The method comprises the following steps: a three-impulse control strategy of a far-lunar maneuver, phase control and period control is designed; based on an orbit plane, an arch line and a perigee height, control time and omnidirectional control impulses for orbit transfer from an initial orbit to a first intermediate orbit are determined; based on an orbit phase, a first tangential impulse for orbit transfer from the first intermediate orbit to a second intermediate orbit is determined; and based on an orbit period, a second tangential impulse for orbit transfer from the second intermediate orbit to the target orbit is determined. The first tangential impulse and the second tangential impulse are taken as variables, and a far-lunar maneuver position vector deviation is taken as a fitness value, and a genetic optimization algorithm is used to solve initial values of the impulses. The method solves the problem of orbit control solution caused by nonlinear enhancement of orbit evolution under perturbation influence and complexity of omnidirectional impulse control parameters.

Description

Technical Field

[0001] This invention relates to the field of aerospace control, and in particular to a method and apparatus for orbit control of a lunar orbiter. Background Technology

[0002] A frozen lunar orbit is a stable orbit with a constant eccentricity and perigee angle. Compared to other lunar orbits, a frozen lunar orbit requires the least amount of propellant, effectively conserving fuel. Furthermore, the arch-shaped orientation of a frozen orbit allows for better tracking and communication at specific locations or in specific states. However, accurately entering a frozen lunar orbit under the influence of orbital perturbations remains a critical challenge. Summary of the Invention

[0003] This invention provides a method and apparatus for orbit control of a lunar spacecraft, enabling precise orbit change to enter the target orbit under the influence of orbital perturbations.

[0004] In a first aspect, embodiments of this application provide a method for orbital control of a lunar spacecraft, including:

[0005] Based on the orbital plane, orbital camber, and perigee altitude of the spacecraft's first intermediate orbit around the moon, the control timing and omnidirectional control pulses for the spacecraft to change its orbit from the initial orbit around the moon to the first intermediate orbit are determined; wherein, the omnidirectional control pulses include the spacecraft's velocity increments in the radial, normal, and tangential directions;

[0006] Based on the orbital phase of the second intermediate orbit, the control time and the first tangential pulse for the spacecraft to change orbit from the first intermediate orbit to the second intermediate orbit are determined;

[0007] Based on the orbital period of the target orbit for lunar orbiting, the control time and second tangential pulse for the spacecraft to change orbit from the second intermediate orbit to the target orbit are determined.

[0008] In some embodiments, after determining the control time and tangential pulse for the spacecraft to change its orbit from the second intermediate orbit to the target orbit based on the orbital period of the target lunar orbit, the method further includes:

[0009] The first intermediate track is obtained by inverse integration of the target track based on the first tangential pulse and the second tangential pulse;

[0010] Determine the line of intersection between the orbital plane of the first intermediate orbit and the orbital plane of the initial orbit, and determine the target intersection point among the intersection points of the intersection line and the initial orbit that is closest to the apogee;

[0011] Based on the latitude argument of the target intersection point, the first intermediate track is inversely integrated and the initial track is integrated in the forward direction to determine the first position vector of the first intermediate track and the second position vector of the initial track;

[0012] The first tangential pulse and the second tangential pulse are optimized and solved based on the first position vector and the second position vector.

[0013] In some embodiments, the control time for the spacecraft to change its orbit from the first intermediate orbit to the second intermediate orbit is the time when the spacecraft reaches the lunar perigee of the first intermediate orbit; the control time for the spacecraft to change its orbit from the second intermediate orbit to the target orbit is the time when the spacecraft reaches the lunar perigee of the second intermediate orbit; the step of obtaining the first intermediate orbit by inverse integration of the target orbit based on the first tangential pulse and the second tangential pulse specifically includes:

[0014] The second intermediate trajectory is determined based on the target trajectory and the second tangential pulse;

[0015] Based on the orbit extrapolation algorithm, the second intermediate orbit is inversely integrated to the moment when it reaches the previous perigee;

[0016] The first intermediate track is determined based on the integrated second intermediate track and the first tangential pulse.

[0017] In some embodiments, determining the first position vector of the first intermediate track and the second position vector of the initial track by performing inverse integration on the first intermediate track and forward integration on the initial track based on the latitude argument of the target intersection point specifically includes:

[0018] The first time point is obtained by inversely integrating the first intermediate trajectory to the latitude argument of the target intersection point, and the second time point is obtained by inversely integrating the initial trajectory to the latitude argument of the target intersection point.

[0019] Determine whether the time difference between the first moment and the second moment is less than a preset time deviation;

[0020] If it is less than, then determine the first position vector of the first intermediate track at the first time and the position vector of the initial track at the second time;

[0021] If the time difference is not less than the preset time deviation, then based on the first intermediate track after inverse integration and the initial track after forward integration, return to the step of determining the intersection line of the track planes until the time difference is less than the preset time deviation.

[0022] In some embodiments, the optimization solution for the first tangential pulse and the second tangential pulse based on the first position vector and the second position vector specifically includes:

[0023] Determine the deviation between the first position vector and the second position vector;

[0024] A genetic optimization algorithm is used to optimize the first tangential pulse and the second tangential pulse by using the deviation as the fitness value of the genetic optimization algorithm.

[0025] In some embodiments, the use of a genetic optimization algorithm, with the deviation as the fitness value of the genetic optimization algorithm, to optimize the first tangential pulse and the second tangential pulse, specifically includes:

[0026] An initial population is generated based on preset boundary settings for the first tangential pulse and the second tangential pulse, respectively; wherein, the initial population contains multiple individuals, and each individual contains a first tangential pulse and a second tangential pulse;

[0027] For each individual in the initial population, the deviation is calculated, and individuals whose calculated deviation is less than a set threshold are designated as parent individuals.

[0028] Genetic optimization algorithms employ crossover and mutation operators to perform genetic operations on the parent individuals to generate a new population. Based on the new population, the steps for calculating the deviation are repeated until the number of genetic operations reaches a threshold. Finally, the optimal individual with the smallest deviation is output.

[0029] Determine the optimal first tangential pulse and the optimal second tangential pulse contained in the optimal individual.

[0030] In some embodiments, after optimizing the first tangential pulse and the second tangential pulse based on the first position vector and the second position vector, the method further includes:

[0031] The first intermediate track is recalculated based on the first and second tangential pulses obtained from the optimized solution, and the intersection point of the initial track and the recalculated first intermediate track is calculated.

[0032] Based on the velocity vector of the initial track at the track intersection and the velocity vector of the re-solved first intermediate track at the track intersection, the omnidirectional control pulse for changing track from the initial track to the re-solved first intermediate track is determined;

[0033] The time when the orbital intersection is reached is taken as the control time for the orbital change from the initial orbit to the first intermediate orbit obtained by resolving the problem.

[0034] In some embodiments, after using the time of reaching the orbital intersection as the control time for the orbital change from the initial orbit to the re-solved first intermediate orbit, the method further includes:

[0035] Based on all the orbital parameters of the target orbit, a differential correction algorithm is used to solve for the control time, the omnidirectional control pulse, the optimized first tangential pulse, and the optimized second tangential pulse of the spacecraft changing orbit from the initial orbit to the first intermediate orbit obtained by resolving the algorithm. This allows the initial orbit to reach the target orbit after three orbital controls by the omnidirectional control pulse, the first tangential pulse, and the second tangential pulse. The total orbital parameters of the target orbit include the semi-major axis, perilune altitude, inclination, right ascension of the ascending node, argument of the perilune, and true anomaly at the target orbit time.

[0036] Secondly, this application provides a lunar orbital control device, the device comprising:

[0037] The lunar maneuver control unit is used to determine the control timing and omnidirectional control pulses for the spacecraft to change its orbit from the initial lunar orbit to the first intermediate orbit, based on the orbital plane, orbital camber, and perilune altitude of the spacecraft's first intermediate orbit around the moon; wherein the omnidirectional control pulses include the spacecraft's velocity increments in the radial, normal, and tangential directions;

[0038] The phase control unit is used to determine the control time and the first tangential pulse for the spacecraft to change its orbit from the first intermediate orbit to the second intermediate orbit based on the orbital phase of the second intermediate orbit;

[0039] The period control unit is used to determine the control time and the second tangential pulse for the spacecraft to change its orbit from the second intermediate orbit to the target orbit based on the orbital period of the target orbit for lunar orbiting.

[0040] In some embodiments, the apparatus further includes an optimization unit for:

[0041] The first intermediate track is obtained by inverse integration of the target track based on the first tangential pulse and the second tangential pulse;

[0042] Determine the line of intersection between the orbital plane of the first intermediate orbit and the orbital plane of the initial orbit, and determine the target intersection point among the intersection points of the intersection line and the initial orbit that is closest to the apogee;

[0043] Based on the latitude argument of the target intersection point, the first intermediate track is inversely integrated and the initial track is integrated in the forward direction to determine the first position vector of the first intermediate track and the second position vector of the initial track;

[0044] The first tangential pulse and the second tangential pulse are optimized and solved based on the first position vector and the second position vector.

[0045] In some embodiments, the control time for the spacecraft to change its orbit from the first intermediate orbit to the second intermediate orbit is the time when the spacecraft reaches the lunar perigee of the first intermediate orbit; the control time for the spacecraft to change its orbit from the second intermediate orbit to the target orbit is the time when the spacecraft reaches the lunar perigee of the second intermediate orbit; the optimization unit is specifically used for:

[0046] The second intermediate trajectory is determined based on the target trajectory and the second tangential pulse;

[0047] Based on the orbit extrapolation algorithm, the second intermediate orbit is inversely integrated to the moment when it reaches the previous perigee;

[0048] The first intermediate track is determined based on the integrated second intermediate track and the first tangential pulse.

[0049] In some embodiments, the optimization unit is specifically used for:

[0050] The first time point is obtained by inversely integrating the first intermediate trajectory to the latitude argument of the target intersection point, and the second time point is obtained by inversely integrating the initial trajectory to the latitude argument of the target intersection point.

[0051] Determine whether the time difference between the first moment and the second moment is less than a preset time deviation;

[0052] If it is less than, then determine the first position vector of the first intermediate track at the first time and the position vector of the initial track at the second time;

[0053] If the time difference is not less than the preset time deviation, then based on the first intermediate track after inverse integration and the initial track after forward integration, return to the step of determining the intersection line of the track planes until the time difference is less than the preset time deviation.

[0054] In some embodiments, the optimization unit is specifically used for:

[0055] Determine the deviation between the first position vector and the second position vector;

[0056] A genetic optimization algorithm is used to optimize the first tangential pulse and the second tangential pulse by using the deviation as the fitness value of the genetic optimization algorithm.

[0057] In some embodiments, the optimization unit is specifically used for:

[0058] An initial population is generated based on preset boundary settings for the first tangential pulse and the second tangential pulse, respectively; wherein, the initial population contains multiple individuals, and each individual contains a first tangential pulse and a second tangential pulse;

[0059] For each individual in the initial population, the deviation is calculated, and individuals whose calculated deviation is less than a set threshold are designated as parent individuals.

[0060] Genetic optimization algorithms employ crossover and mutation operators to perform genetic operations on the parent individuals to generate a new population. Based on the new population, the steps for calculating the deviation are repeated until the number of genetic operations reaches a threshold. Finally, the optimal individual with the smallest deviation is output.

[0061] Determine the optimal first tangential pulse and the optimal second tangential pulse contained in the optimal individual.

[0062] In some embodiments, the optimization unit is further configured to:

[0063] The first intermediate track is recalculated based on the first and second tangential pulses obtained from the optimized solution, and the intersection point of the initial track and the recalculated first intermediate track is calculated.

[0064] Based on the velocity vector of the initial track at the track intersection and the velocity vector of the re-solved first intermediate track at the track intersection, the omnidirectional control pulse for changing track from the initial track to the re-solved first intermediate track is determined;

[0065] The time when the orbital intersection is reached is taken as the control time for the orbital change from the initial orbit to the first intermediate orbit obtained by resolving the problem.

[0066] In some embodiments, the apparatus further includes a correction unit for:

[0067] Based on the orbital parameters of the target orbit, a differential correction algorithm is used to solve for the control time, the omnidirectional control pulse, the optimized first tangential pulse, and the optimized second tangential pulse of the spacecraft changing orbit from the initial orbit to the first intermediate orbit obtained by resolving the algorithm. This allows the initial orbit to reach the target orbit after three orbital controls via the omnidirectional control pulse, the first tangential pulse, and the second tangential pulse. The total orbital parameters of the target orbit include the semi-major axis, lunar altitude, inclination, right ascension of the ascending node, argument of the lunar perigee, and true anomaly at the target orbit time.

[0068] Thirdly, this application provides an electronic device, comprising:

[0069] Memory, used to store program instructions;

[0070] A processor is configured to invoke program instructions stored in the memory and execute the steps of the method described in any one of the first aspects according to the obtained program instructions.

[0071] Fourthly, this application provides a computer-readable storage medium storing a computer program, the computer program including program instructions, which, when executed by a computer, cause the computer to perform the method described in any one of the first aspects.

[0072] Fifthly, this application provides a computer program product comprising: computer program code, which, when run on a computer, causes the computer to perform the method described in any one of the first aspects.

[0073] The beneficial effects of this application are as follows:

[0074] To address the challenges of enhanced nonlinearity in orbit evolution under large perturbations, complex omnidirectional pulse control targets, and difficulties in solving multi-pulse problems due to multi-pulse coupling, this application proposes a three-pulse orbit-changing strategy involving lunar maneuvering control, phasing control, and periodic control. An optimization model is constructed, using the tangential pulses of phasing and periodic control as planning variables and the deviation of the orbital splicing position vector at the lunar maneuvering control position point as the fitness value. A genetic optimization algorithm is employed to search for the optimal values ​​of the two tangential pulses and to solve for the initial values ​​of the omnidirectional pulses for lunar maneuvering control. Finally, a differential correction algorithm is used for precise solution, yielding the accurate control parameters for entering the highly elliptical lunar orbit, thus achieving the target orbit requirements. Attached Figure Description

[0075] 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.

[0076] Figure 1 A flowchart illustrating a lunar orbiter trajectory control method provided in an embodiment of the present invention;

[0077] Figure 2 A schematic flowchart of a pulse optimization method provided in an embodiment of the present invention;

[0078] Figure 3 This is a schematic diagram illustrating the implementation process of a genetic optimization algorithm provided in an embodiment of the present invention;

[0079] Figure 4 A schematic diagram illustrating the relationship between flight time and orbital period provided in an embodiment of the present invention;

[0080] Figure 5 This is a schematic diagram illustrating the process of solving a problem using a genetic optimization algorithm, as provided in an embodiment of the present invention.

[0081] Figure 6 This is a schematic diagram illustrating the process of calculating fitness values ​​according to an embodiment of the present invention;

[0082] Figure 7 This is a schematic diagram of the structure of a lunar orbiter orbiting control device provided in an embodiment of the present invention;

[0083] Figure 8 This is a schematic diagram of the structure of an electronic device provided in an embodiment of the present invention. Detailed Implementation

[0084] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings. Obviously, the described embodiments are merely some embodiments of this invention, and not all embodiments. Based on the embodiments of this invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this invention.

[0085] With the continuous development of deep space technology, human exploration of the moon has deepened, and the scope of lunar exploration has begun to expand from the "near side" to the "far side" and its "polar regions." Due to the moon's obstruction, lunar probes at the lunar south pole and far side cannot establish direct communication with ground stations. Therefore, establishing a lunar relay communication system is an inevitable way to meet the needs of future lunar exploration. Earth (and other planets) and the moon's satellites all have frozen orbits due to the north-south asymmetry of their central gravitational fields. A frozen orbit is a stable orbit whose eccentricity and perigee argument (taking specific values) remain constant. Because frozen orbits have a stable state that remains constant or has minimal change over a long period, the amount of propellant used for orbit maintenance is minimized, greatly saving fuel consumption. At the same time, utilizing the arch orientation characteristics of frozen orbits, it is possible to achieve better tracking and communication at specific locations or states. As a lunar relay communication orbit, serving communication and navigation services at specific times and locations, frozen orbit parameters, including orbital period, orbital plane, and orbital position, have strict requirements. Therefore, it is necessary to perform full-parameter orbital entry control to freeze the orbit at a specific time and ensure that all orbital parameters meet the requirements, in order to meet the orbital requirements for lunar relay services. Additionally, other specific mission requirements also necessitate full-parameter orbital entry control.

[0086] However, spacecraft orbit entry control is affected by orbital perturbations. Orbital perturbations during lunar orbit mainly include: lunar non-spherical gravitational perturbations, Earth-centric gravitational perturbations, Earth-centric gravitational perturbations, solar-centric gravitational perturbations, radiation pressure perturbations, and planetary-centric gravitational perturbations. For highly elliptical orbits with apogee altitudes greater than 1000 km, Earth's gravitational perturbations are significantly enhanced, leading to drastic evolution of the lunar frozen orbit. The higher the apogee altitude, the more drastic the orbital evolution, with the semi-major axis evolution reaching hundreds to thousands of kilometers, resulting in significantly enhanced nonlinear characteristics. To address the difficulty in solving orbit entry control due to the enhanced nonlinearity of orbital evolution under the influence of large perturbations, this application proposes a three-pulse orbital maneuver control method to enter the target orbit where all orbital parameters meet the requirements after the spacecraft enters its initial lunar orbit. Compared to the traditional four-pulse control scheme, the three-pulse entry into the target orbit can accelerate the control of the spacecraft entering the target orbit in terms of time series.

[0087] The solution of this application will be described below with reference to embodiments. See also... Figure 1 This is a schematic flowchart illustrating a lunar orbiter trajectory control method provided in an embodiment of this application. Exemplarily, this application... Figure 1 The execution entity of the method flow shown is not limited and can be any electronic device at the ground center. Figure 1 The method flow shown specifically includes:

[0088] 101. Based on the orbital plane, orbital camber, and lunar perihelion altitude of the spacecraft's first intermediate orbit around the moon, determine the control timing and omnidirectional control pulses for the spacecraft to change its orbit from the initial orbit around the moon to the first intermediate orbit.

[0089] The omnidirectional control pulse includes the spacecraft's velocity increments in the radial, normal, and tangential directions. For example, applying the omnidirectional control pulse to the spacecraft at the control moment can ensure that the spacecraft's orbit conforms to the orbital plane, orbital camber, and perilune altitude of the first intermediate orbit. Alternatively, the omnidirectional control pulse can be applied near the apogee; applying the omnidirectional control pulse at an orbital position close to the apogee can result in a smaller normal pulse.

[0090] Optionally, the orbital plane is described by two orbital parameters: the orbital inclination and the right ascension of the ascending node, both of which are affected by the normal velocity. Therefore, when controlling the spacecraft's orbital maneuver to satisfy the orbital plane of the first intermediate orbit, the adjustment amount of the orbital inclination can be determined based on the orbital inclination of the initial orbit and the orbital inclination of the first intermediate orbit, as well as the adjustment amount of the right ascension of the ascending node based on the orbital inclination and the right ascension of the ascending node. Further, the normal pulse for the spacecraft's orbital maneuver control, i.e., the normal velocity increment, is determined based on the orbital inclination adjustment and the ascending node right ascension adjustment.

[0091] The orbital camber is controlled by the lunar peril angle, a key orbital parameter. Furthermore, changes in radial, normal, and tangential velocities all affect the lunar peril angle. Therefore, the adjustment amount of the lunar peril angle can be determined based on the lunar peril angles of the initial orbit and the first intermediate orbit. Further, the corresponding normal pulse, tangential pulse, and radial pulse are determined based on the lunar peril angle adjustment amount, thus determining the velocity increments in the three directions.

[0092] The perigee altitude is described by two orbital parameters: the semi-major axis and the eccentricity. These parameters are influenced by radial and tangential velocities. Therefore, to control the perigee altitude to meet the requirements of the first intermediate orbit, the semi-major axis adjustment and eccentricity adjustment of the initial orbit relative to the first intermediate orbit can be calculated first. Further, radial and tangential pulses are calculated based on the semi-major axis and eccentricity adjustments.

[0093] 102. Based on the orbital phase of the second intermediate orbit, determine the control time and the first tangential pulse for the spacecraft to change orbit from the first intermediate orbit to the second intermediate orbit.

[0094] The adjustment amount of the orbital phase is determined by the adjustment amount of the orbital angular velocity and the flight time, and the orbital angular velocity is described by the orbital parameter of the semi-major axis. Therefore, the corresponding tangential pulse can be determined by the adjustment amount of the orbital semi-major axis.

[0095] For example, applying a first tangential pulse to a spacecraft in a first intermediate orbit at the corresponding control time can change the orbital phase so that the orbital phase of the second intermediate orbit is satisfied after the orbital change. Optionally, the control time for applying the first tangential pulse can be the time when the spacecraft reaches the perilune of the first intermediate orbit.

[0096] 103. Based on the orbital period of the target orbit, determine the control time and the second tangential pulse for the spacecraft to change its orbit from the second intermediate orbit to the target orbit.

[0097] For example, applying a second tangential pulse to the spacecraft at the control moment can change the orbital period, ensuring that the orbital parameters after the orbital change fully meet the requirements of the target orbit, thus achieving full parameter control of the target orbit. The control moment can be the time when the spacecraft reaches the perigee of the second intermediate orbit.

[0098] Optionally, when determining the second tangential pulse, the semi-major axis adjustment of the track can be determined first based on the track period adjustment, and then the corresponding second tangential pulse can be determined based on the semi-major axis adjustment.

[0099] Based on the above scheme, the orbit change scheme proposed in this application achieves precise orbit change to the target orbit based on three-pulse control. The joint control strategy of one omnidirectional pulse and two tangential pulses, compared with the traditional four-pulse control scheme that decouples the in-plane parameter control and the orbital plane parameter control, can enable the spacecraft to enter the target orbit more quickly.

[0100] In some embodiments, before executing the three pulses, the spacecraft first needs to change its orbit from a hyperbolic orbit to its initial lunar orbit. After launch, the spacecraft enters a lunar transfer orbit and, after a period of flight, arrives at the Moon. Upon arrival, its orbit relative to the Moon is a hyperbolic orbit, requiring a change from this hyperbolic orbit to its initial lunar orbit. Optionally, this change to the initial orbit can be performed when the spacecraft reaches its perigee. See Table 1 below, which exemplarily illustrates the four orbit changes required to transition from a hyperbolic orbit to the target orbit.

[0101] Table 1. Control Strategy for Four Orbit Changes

[0102]

[0103] The process of the four orbit changes will be described in detail below with reference to Table 1:

[0104] (1) Capture Control 1 (Lunar Orbit Insert 1, LOI1)

[0105] Acquisition Control 1 is the pulse control prior to executing the three pulses proposed in this application. When the spacecraft arrives at the Moon, its orbit relative to the Moon is a hyperbolic orbit. Therefore, Acquisition Control 1 controls the spacecraft to change its orbit from the hyperbolic orbit to an elliptical orbit around the Moon, i.e., the initial orbit (for the sake of consistency, the initial orbit will be used as an example in the following description).

[0106] The orbital velocity of a spacecraft satisfies the vitality formula, which can be found in formula (1) below:

[0107]

[0108] Where μ is the gravitational constant of the Moon, v is the velocity of the spacecraft in the hyperbolic orbit, r is the distance between the spacecraft and the center of the Moon, and a is the semi-major axis of the hyperbolic orbit.

[0109] The velocity of the spacecraft in its initial orbit can be determined using the following formula (2):

[0110]

[0111] Among them, v T Let be the spacecraft's initial orbital velocity, μ be the Moon's gravitational constant, r be the spacecraft's distance from the Moon's center, and a be the velocity of the spacecraft in its initial orbit. T The semi-major axis of the initial orbit.

[0112] For example, when determining the tangential pulse of the LOI, the velocity increment of the spacecraft changing from the hyperbolic orbit to the initial orbit can be calculated. For example, the velocity increment can be determined according to the following formula (3):

[0113] Δv=v T -v Formula (3)

[0114] Where Δv is the velocity increment, v T Let v be the initial velocity of the spacecraft's orbit, and v be the velocity of the spacecraft's hyperbolic orbit.

[0115] Therefore, by applying the calculated velocity increment in the tangential direction at the perigee, the hyperbolic orbit can be changed back to the initial orbit.

[0116] Furthermore, when determining the initial orbit, the orbital period is generally used for description. The relationship between the orbital period and the semi-major axis of the orbit can be found in formula (4) below:

[0117]

[0118] Where P is the orbital period, a is the semi-major axis of the orbit, and μ is the gravitational constant of the Moon.

[0119] Therefore, orbital periodic control is equivalent to orbital semi-major axis control.

[0120] (2) Lunar Apogee Maneuver (LAM)

[0121] The apolunar maneuvering control is the first pulse control in the three-pulse orbital maneuvering scheme proposed in this application. The apolunar maneuvering control is used to control the spacecraft to perform orbital changes from its initial lunar orbit to satisfy the orbital camber, orbital plane, and perilune altitude of the first intermediate orbit. Therefore, it can be seen that when performing maneuvering control, it is necessary to control the three orbital characteristic parameters—the orbital camber, orbital plane, and perilune altitude—separately. The control of these three orbital characteristic parameters is described below:

[0122] 1. Track plane control:

[0123] The orbital plane is described by two orbital parameters: the orbital inclination and the right ascension of the ascending node. Changes in the normal velocity affect these two parameters. Based on this, this application proposes that the normal pulse can be determined by combining the adjustment of the orbital inclination and the adjustment of the right ascension of the ascending node. For example, the normal pulse can be calculated using the following formula (5):

[0124]

[0125] Where, Δv n Let Δi be the normal pulse, ΔΩ be the orbital inclination adjustment relative to the first intermediate orbit, ΔΩ be the right ascension adjustment relative to the ascending node of the first intermediate orbit, and i be the orbital inclination of the initial orbit. h is the angular momentum. Where μ is the gravitational constant of the Moon, a is the semi-major axis of the initial orbit, e is the initial orbital eccentricity, and r is the distance between the spacecraft and the center of the Moon. Where e is the initial orbital eccentricity, f is the true anomaly angle of the initial orbit, p is the semi-major diameter, and = a(1-e) 2 ).

[0126] After determining the normal pulse, the control time can be further determined. For example, the latitudinal arguments u1 and u2 of the two control points (the two intersection points of the initial orbital plane and the first intermediate orbital plane, respectively) for the joint control of the orbital inclination and the right ascension of the ascending node are determined using the following formulas (6) and (7):

[0127]

[0128]

[0129] The parameters in formulas (6) to (7) can be explained in the above formulas, and will not be repeated here.

[0130] The normal velocity increment between two control points can be determined based on the latitudinal argument of the two control points. For example, see formulas (8) and (9) below:

[0131]

[0132]

[0133] Where, Δv nu1 Let Δv be the velocity increment of one of the two control points. nu2 Let r1 be the velocity increment of the other of the two control points, where r1 and r2 represent the distances between the two control points and the center of the moon. The remaining parameters can be found in the explanation of the formula above, and will not be repeated here.

[0134] According to formulas (8) and (9), when the true perihelion angle f = π, the distance between the control point and the center of the moon is the largest, and the normal velocity increment is the smallest. Therefore, the apogee control is set to be executed at the moment when the moon is close to the apogee, so as to minimize the normal velocity increment.

[0135] 2. Track arch control:

[0136] The orbital camber is described by the near-lunar point argument, and changes in radial, tangential, and normal velocities all affect the near-lunar point argument. Therefore, this application proposes to determine the near-lunar point argument adjustment relative to the first intermediate orbit and to calculate the omnidirectional velocity increments satisfying this adjustment. As an example, the velocity increments in the three directions can be determined using the following formula (10):

[0137]

[0138] Where Δω is the perilune angle adjustment relative to the first intermediate orbit, p is the semi-circle, μ is the lunar gravitational constant, f is the true anomaly angle, e is the eccentricity, and Δv r Δv is the radial velocity increment, r is the distance between the spacecraft and the center of the moon. t Let ω be the tangential velocity increment, ω be the lunar perihelion argument of the initial orbit, i be the orbital inclination of the initial orbit, and Δv be the velocity increment. n This represents the normal velocity increment.

[0139] 3. Lunar perigee height control:

[0140] The lunar perihelion altitude is controlled by the semi-major axis and eccentricity, as detailed in formula (11) below:

[0141] h p =a(1-e)-R M Formula (11)

[0142] Among them, h p Let be the lunar perigee velocity of the orbit, 'a' be the semi-major axis of the orbit, 'e' be the eccentricity of the orbit, and 'R' be the eccentricity of the orbit. M This is the radius of the Moon's equatorial axis.

[0143] The adjustment amount of the lunar perigee altitude of the initial orbit relative to the first intermediate orbit is:

[0144] Δh p = (1-e)Δa-aΔe Formula (12)

[0145] Where Δhp is the lunar perigee height adjustment, e is the eccentricity of the initial orbit, Δa is the semi-major axis adjustment relative to the first intermediate orbit, a is the semi-major axis of the initial orbit, and Δe is the eccentricity adjustment relative to the first intermediate orbit.

[0146] Furthermore, the method for determining the radial and tangential velocity increments based on the semi-major axis adjustment and the perigee height adjustment can be found in the following formulas (13)-(14):

[0147]

[0148]

[0149] The parameters in formulas (13) to (14) can be found in the relevant explanations in the formulas above, and will not be repeated here.

[0150] According to formulas (13) and (14), the direct relationship between the lunar altitude adjustment and the radial velocity increment and tangential velocity increment can be obtained, and the details can be found in formula (15):

[0151]

[0152] The parameters in formula (15) can be found in the relevant explanations in the formula above, and will not be repeated here.

[0153] (3) Capture Control 2 (Lunar Orbit Insert 2, LOI2)

[0154] Capture control 2 is the second pulse control in the three-pulse control proposed in this application. Capture control 2 is used for track phasing, performing a track change on the first intermediate track to satisfy the track phase of the second intermediate track. Since the track phase is related to the track angular velocity, and the track angular velocity is described by the semi-major axis of the track, this application uses tangential pulse control to adjust the semi-major axis of the track.

[0155] For example, the relationship between the phase adjustment of the orbit and the angular velocity adjustment of the orbit can be seen in the following formula (16):

[0156] Δu=Δn·t Formula (16)

[0157] Where Δu is the phase adjustment amount, Δn is the angular velocity adjustment amount, and t is the flight time.

[0158] The relationship between the orbital angular velocity and the semi-major axis of the orbit can be found in the following formula (17):

[0159]

[0160] Where n is the angular velocity, μ is the gravitational constant of the Moon, and a is the semi-major axis of the orbit.

[0161] Therefore, the relationship between the track angular velocity adjustment and the track semi-major axis adjustment can be found in the following formula (18):

[0162]

[0163] Where Δn is the angular velocity adjustment, P is the orbital period, Δa is the orbital semi-major axis adjustment, and a is the orbital semi-major axis.

[0164] Therefore, the relationship between the phase adjustment of the track and the semi-major axis adjustment of the track can be found in the following formula (19):

[0165]

[0166] Where Δu is the phase adjustment of the orbit, t is the flight time, P is the orbital period, a is the semi-major axis of the orbit, and Δa is the semi-major axis adjustment.

[0167] According to the vitality formula, the semi-major axis of the track is related to the tangential velocity increment. Therefore, the phase adjustment of the track can be calculated using the following formula (20):

[0168]

[0169] The parameters in formula (20) can be found in the explanation of the above formula, and will not be repeated here.

[0170] Therefore, the tangential velocity increment from the first intermediate track to the second intermediate track can be determined by the following formula (21):

[0171]

[0172] The parameters in formula (21) can be found in the explanation of the above formula, and will not be repeated here.

[0173] (4) Capture Control 3 (Lunar Orbit Insert 3, LOI3)

[0174] The capture control 3 is the third pulse control in the three-pulse control proposed in this application. The capture control 3 is used to adjust the orbit period so that the second intermediate orbit changes to meet all the orbit parameters of the target orbit. The orbit period is related to the orbit semi-major axis (see formula (4) for details), so the tangential velocity can be used to control the change of the orbit semi-major axis so that the orbit semi-major axis meets the target orbit, and thus the orbit period meets the target orbit.

[0175] For example, the change in the semi-major axis of the second intermediate orbit relative to the target orbit can be represented by the following formula (22):

[0176]

[0177] Where Δa represents the semi-major axis adjustment relative to the target orbit, the other parameters can be found in the explanation of the above formula, and will not be repeated here.

[0178] Therefore, the tangential velocity increment applied by capture control 3 can be obtained as:

[0179]

[0180] The parameters in formula (23) can be found in the explanation of the above formulas, and will not be repeated here.

[0181] As described above, in the apolunar maneuver control, the normal velocity increment controls the orbital inclination, the right ascension of the ascending node, and the perilune argument, while the radial and tangential velocity increments control the perilune argument and perilune altitude. It can be seen that the apolunar maneuver control has mutually coupled control effects on the orbital parameters, and there is also a related coupling relationship between the apolunar maneuver control and capture control 2 and capture control 3. Therefore, the accuracy of the pulses obtained using the above calculation method is relatively poor. In view of this, this application proposes a scheme to optimize the apolunar maneuver control timing and each control pulse based on the plane intersection of the first intermediate orbit and the initial orbit, which will be described in detail below.

[0182] See Figure 2 The above is a flowchart illustrating a pulse optimization method provided in an embodiment of this application, specifically including:

[0183] 201. Based on the orbit extrapolation algorithm, the first intermediate orbit is obtained by inverse integration of the target orbit according to the first tangential pulse and the second tangential pulse.

[0184] The orbit extrapolation algorithm specifically involves calculating the spacecraft's position and velocity in chronological order using an integral algorithm based on the spacecraft's orbital dynamics equations. The first tangential pulse is applied to control the spacecraft's transition from the first intermediate orbit to the second intermediate orbit; the timing of executing the first tangential pulse is the moment the spacecraft reaches the lunar perigee of the first intermediate orbit. The second tangential pulse is applied to control the spacecraft's transition from the second intermediate orbit to the target orbit; the timing of executing the second tangential pulse is the moment the spacecraft reaches the lunar perigee of the second intermediate orbit.

[0185] For example, when performing inverse integration on the target orbit, the second intermediate orbit can be determined first based on the target orbit and the second tangential pulse.

[0186] Since the target trajectory is known, the trajectory position vector and velocity vector after the second tangential pulse are also known. Therefore, the velocity vector of the second intermediate trajectory can be calculated by reverse calculation based on the trajectory velocity vector after the second tangential pulse and the second tangential pulse itself. As an example, the velocity vector of the second intermediate trajectory can be calculated using the following formula (24):

[0187]

[0188] in, The velocity vector of the second intermediate orbit. M is the velocity vector of the target orbit. IO Let Δv be the transformation matrix from the orbital frame to the inertial frame. 3t This is the second tangential pulse.

[0189] Furthermore, the orbital parameters of the second intermediate orbit can be determined based on the velocity vector and position vector of the second intermediate orbit. After determining the second intermediate orbit, the second intermediate orbit can be inversely integrated to the moment of reaching the previous perigee based on the orbit extrapolation algorithm. The first intermediate orbit is then determined based on the integrated second intermediate orbit and the first tangential pulse. For example, the second intermediate orbit obtained after inverse integration is the orbit after the execution of the first tangential pulse. After obtaining this orbit, the velocity vector of the first intermediate orbit can be calculated based on the velocity vector of this orbit and the first tangential pulse. As an example, the velocity vector of the first intermediate orbit can be calculated using the following formula (25):

[0190]

[0191] in, The velocity vector of the first intermediate track. M is the velocity vector of the second intermediate orbit. IO Let Δv be the transformation matrix from the orbital frame to the inertial frame. 2t This is the first tangential pulse.

[0192] Furthermore, the orbital parameters of the first intermediate track can be determined based on the position vector and velocity vector of the first intermediate track.

[0193] 202. Determine the two intersection points between the orbital plane of the first intermediate orbit and the orbital plane of the initial orbit, and determine the target intersection point that is closest to the apogee among the two intersection points on the initial orbit.

[0194] For example, as described in the above embodiments, the initial orbit is an elliptical orbit obtained by transforming a hyperbolic orbit. Therefore, the hyperbolic orbit can be integrated up to the lunar perihelion to obtain the orbit before the execution of capture control 1. Let the orbital velocity of this orbit be... And let the orbital velocity of the initial orbit obtained after executing capture control 1 be... according to Know The pulse Δv of capture control 1 can be calculated using the above formulas (1)-(4). 1t .

[0195] Furthermore, the velocity vector of the initial orbit can be calculated based on the orbit integrated from the hyperbolic orbit to the lunar perihelion and the pulse of capture control 1. For example, the velocity vector of the initial orbit can be calculated using the following formula (26):

[0196]

[0197] in, The velocity vector of the initial trajectory. Let M be the velocity vector of the hyperbolic orbit. IO Let Δv be the transformation matrix from the orbital frame to the inertial frame. 1t To capture the pulse of control 1.

[0198] Furthermore, the initial trajectory can be calculated based on the velocity vector and position vector of the initial curve.

[0199] The intersection point of the calculated initial orbit and the orbital plane of the first intermediate orbit calculated in step 201 is determined. For example, the initial orbit and the first intermediate orbit are not coplanar. The intersection of these two non-coplanar orbits forms an intersection line, which is a line passing through the center of the moon. The intersection line intersects the initial orbit at two points: one near the perilunar point of the orbit, and the other near the apolunar point. As described in the above embodiments, apolunar maneuvering control should be performed near the apolunar point to reduce the normal velocity increment. Therefore, for ease of description, the intersection point closer to the apolunar point between the first intermediate orbit and the initial orbit will be referred to as the target intersection point.

[0200] 203. Based on the latitude argument of the target intersection point, perform inverse integration calculation on the first intermediate track and forward integration calculation on the initial track to determine the first position vector of the first intermediate track and the second position vector of the initial track.

[0201] Since the first intermediate track and the initial track do not connect, it is necessary to adjust the two tracks using the latitude argument of the target intersection point. For example, the first intermediate track can be adjusted using an orbital extrapolation algorithm. The latitude argument obtained by inverse integration to the target intersection point is... Track of Time And based on the orbit extrapolation algorithm, the initial orbit The latitude argument obtained by forward integration to the target intersection point is... Track of Time

[0202] because and The difference is significant, therefore iterative calculations are required: Calculation Know The two intersection points of the orbital planes of these two orbits are used to determine the target intersection point closest to the apogee, and then the result is obtained by integrating the latitude argument of the target intersection point. and renew And updates After updating, return to the steps for calculating the intersection points of the orbital planes, until... εt is a preset time deviation, and the corresponding first intermediate track is output. The first position vector at time t and the corresponding initial orbit are in The second position vector at time.

[0203] 204. The first tangential pulse and the second tangential pulse are optimized and solved based on the first position vector and the second position vector.

[0204] For example, the deviation between the first position vector and the second position vector can be calculated first. Since two unknowns—the first tangential pulse and the second tangential pulse—are used in the calculation of both the first and second position vectors, this application proposes using a genetic optimization algorithm. The deviation between the two position vectors is used as the fitness value of the genetic optimization algorithm, and the formula for calculating the deviation using the first and second tangential pulses is used as the fitness function (or performance index function). The first and second tangential pulses are then optimized based on the objective of a deviation of 0.

[0205] Genetic optimization algorithm is an adaptive global optimization probabilistic search algorithm. Borrowing terminology from biological genetics, the basic idea of ​​a genetic algorithm can be described as follows: A genetic optimization algorithm starts with a population representing the potential solutions to a problem. This population consists of a certain number of individuals encoded with genes. Each individual is essentially a chromosome carrying characteristics. Chromosomes, as the main carriers of genetic material, are collections of multiple genes. Their internal expression (genotype) is a certain combination of genes, which determines the individual's external expression (trait). Therefore, initially, a mapping from phenotype to genotype, i.e., encoding, needs to be implemented. After the initial population is generated, it evolves generation by generation according to the principle of "survival of the fittest," producing increasingly better individuals. In each generation, individuals are selected based on their fitness in the problem domain, and crossover and mutation are performed using genetic operators to generate a population representing a new set of solutions. This process leads to a population that, like natural evolution, is more adapted to the environment in subsequent generations than in previous generations. The best individual in the final generation, after decoding, can serve as an approximate optimal solution to the problem. As an example, the implementation process of the genetic optimization algorithm can be found in [link to relevant documentation]. Figure 3 As shown below. Figure 3 The implementation process of the genetic optimization algorithm is introduced as follows:

[0206] Step 1: Initialization. Set evolutionary parameters and randomly generate the initial population.

[0207] Step 2: Calculate the fitness value of each individual in the initial population. Different fitness functions are used for different problems. You can set a corresponding fitness function based on the specific problem scenario, and input each individual into the fitness function to obtain the corresponding fitness value.

[0208] Step 3: Determine if the optimization criteria are met. If they are met, end the evolution and output the approximate optimal solution; otherwise, continue the evolution.

[0209] Step 4: Selection. A selection algorithm can be pre-set to select superior individuals from the current population to serve as parents for the next generation.

[0210] Step 5: Crossover. The crossover operator is applied to the parent population. Crossover is the most important genetic operation in genetic optimization algorithms. Through crossover, a new generation of individuals is obtained, combining the characteristics of its parents.

[0211] Step Six: Mutation. Apply the mutation operator to the current population. First, randomly select an individual from the population. For the selected individual, randomly change its data with a certain probability to obtain a new individual.

[0212] After selection, crossover, and mutation operations, a new population will be obtained. Then, return to step two and continue until an approximate optimal solution is output.

[0213] In this application, the pulse of capture control 1 for the hyperbolic orbit change to the initial orbit is known, but the pulses of capture control 2 and capture control 3 are unknown. This is also related to the solution of the omnidirectional control pulse and control timing of the far-moon maneuver control. Therefore, this application proposes to use the first tangential pulse applied by capture control 2 and the second tangential pulse applied by capture control 3 as planning variables in the genetic optimization algorithm. As for the fitness function in the genetic optimization algorithm, this application proposes to use the position vector deviation between the first position vector of the first intermediate orbit located at the target intersection point and the second position vector of the initial orbit located at the target intersection point as the fitness function. Therefore, the goal of genetic optimization is to obtain a fitness value of zero according to the fitness function, that is, the position vector deviation is zero, and the initial orbit and the first intermediate orbit are consistent. For example, the formula for calculating the position vector deviation can be found in the following formula (27):

[0214]

[0215] Where d is the position vector deviation, The first position vector, This is the second position vector.

[0216] The goal of genetic optimization is to solve d. min =0, the following is a detailed introduction to the execution process of the genetic optimization algorithm.

[0217] For example, coarse initial values ​​for the planning variables can be determined, i.e., coarse initial values ​​for the first and second tangential pulses can be determined. The orbital control proposed in this application includes capture control 1, apolunar maneuver control, capture control 2, and capture control 3. Let the control time of capture control 1 (i.e., the time when the spacecraft reaches the perigee of the hyperbolic orbit) be t1, and the control time of apolunar maneuver control be t... m The control time of capture control 2 (i.e., the time when the spacecraft reaches the lunar perigee of the first intermediate orbit) is t2, the control time of capture control 3 (i.e., the time when the spacecraft reaches the lunar perigee of the second intermediate orbit) is t3, and the time of the target orbit is t. e Because of the target orbit's time t e The time of lunar perihelion of the target orbit is a known value. The control time t3 of capture control 3 is the target orbit time. Therefore, t3 = t e .

[0218] When setting coarse initial values, the influence of orbital perturbations and lunar maneuvering control on the orbital period can be ignored. The orbital flight processes of the three tangential orbital controls—Capture Control 1, Capture Control 2, and Capture Control 3—are interconnected. Therefore, the orbital period after any control is consistent with the period before the next control. Since capture control 1, capture control 2, and capture control 3 all execute control at the perigee, exhibiting the characteristic of full-cycle flight, the total flight time from the control time of capture control 1 to the control time of capture control 3 is equal to the sum of the orbital periods of the initial orbit after capture control 1 and the second intermediate orbit after capture control 2, i.e., the total flight time. in, The orbital period of the initial orbit. The orbital period of the second intermediate orbit.

[0219] As an example, see Figure 4 This is a schematic diagram illustrating the relationship between flight time and orbital period. The orbital period after capture control 1 is... The orbital period of the highly elliptical orbit is a known value. The orbital periods before and after the execution of capture control 2 are respectively... and Ignoring the effects of orbital perturbations and far-lunar maneuvering (LAM) on the orbital period, Therefore, the orbital period before capture control 2 is executed. Given that the orbital periods before and after the execution of capture control 3 are respectively... and The trajectory after capture control 3 is executed is the target trajectory, therefore It is known. In summary, only the orbital period after capture control 2 is executed and the orbital period before capture control 3 is executed are unknown, and

[0220] Based on the above analysis, this application utilizes the matching characteristic of orbital period and flight time to solve for the coarse initial values ​​of the first tangential pulse for phase modulation control and the second tangential pulse for periodic control, providing a reference for the interval setting of the population in the genetic optimization algorithm. When setting the coarse initial values ​​of the first and second tangential pulses, the total flight time and the orbital period after the execution of capture control 1 can be used as a reference. Calculate the orbital period after capture control 2 is executed, i.e.: According to known and the calculated And the vitality formula for calculating the first tangential pulse Δv 2t Furthermore, according to and known Calculate the second tangential pulse Δv 3t .

[0221] Furthermore, based on the preset boundary setting value δv 2ts ,δv 2te ,δv 3ts ,δv 3te Determine Δv 2t and Δv 3tThe range of values ​​for the initial coarse value: Δv 2i ∈[|Δv 2i |-δv 2ts ,|Δv 2i |+δv 2te ], Δv 3t ∈[|Δv 3t |-δv 3ts ,|Δv 3t |+δv 3te ].

[0222] After setting the initial values ​​of the planning variables, a genetic optimization algorithm can be used to solve the problem. For example, the solution process can be found in [link to documentation]. Figure 5 Specifically, it includes:

[0223] Step 501: Randomly generate the initial population.

[0224] For example, Δv can be derived from the embodiments described above. 2t and Δv 3t The range of initial values ​​is arbitrarily selected from a certain number of individuals, each containing a Δv. 2t The rough initial value and a Δv 3t The initial values ​​are then used. Further, the selected individuals are modified to generate the initial population.

[0225] Step 502: Calculate the fitness value for each individual.

[0226] For example, each individual can be decoded and input into the fitness function to obtain the fitness value corresponding to each individual. Furthermore, the individual with the smallest fitness value can be recorded and a fitness preservation operation can be performed.

[0227] Step 503: Determine the parent individuals based on the fitness values ​​of each individual.

[0228] For example, the parent individual can be an individual whose fitness value is less than a set threshold.

[0229] Step 504: Genetic operations are performed on the parent individuals using crossover and mutation operators to generate the offspring population.

[0230] Step 505: Determine whether the number of genetic operations has reached the threshold.

[0231] If so, proceed to step 506.

[0232] If not, return to step 502.

[0233] Step 506: End evolution and determine the optimal individual with the smallest fitness value.

[0234] In one possible implementation, when calculating the fitness value in step 502 above, see below. Figure 6 The process shown specifically includes:

[0235] Step 601: Obtain any individual.

[0236] Each individual includes the first tangential pulse Δv 2t Second tangential pulse Δv 3t .

[0237] Step 602: Calculate the first intermediate orbit based on any individual.

[0238] For example, first, the target trajectory is inversely integrated, and then based on Δv 3t Calculate the second intermediate trajectory before the execution of capture control 3. Further, perform inverse integration on the second intermediate trajectory and, based on Δv... 2t Calculate the first intermediate track before the execution of capture control 2.

[0239] Step 603: Determine the initial orbit.

[0240] For example, the hyperbolic orbit can be integrated to the lunar perihelion time, and the initial orbit after the execution of capture control 1 can be calculated based on the known orbital period.

[0241] Step 604: Determine the target intersection point between the first intermediate track and the initial track.

[0242] For example, two intersection points of the orbital planes of the initial orbit and the first intermediate orbit can be determined first. Further, the target intersection point that is closest to the apogee of the initial orbit can be determined from the two intersection points.

[0243] Step 605: Adjust the first intermediate track and the initial track according to the latitude argument of the target intersection point, and determine the position vector deviation corresponding to the adjusted intersection point.

[0244] For example, the process of adjusting the track (integral calculation) can be referred to the description in the above embodiments, and will not be repeated here.

[0245] Step 606: Determine if the required number of iterations has been met.

[0246] If so, proceed to step 607.

[0247] If not, return to step 604.

[0248] Step 607: Output the fitness value for any individual.

[0249] In some embodiments, after optimizing the first and second tangential pulses using a genetic optimization algorithm, the omnidirectional control pulses and control timing for apolunar maneuvering control can be further determined. For example, when determining the control timing for apolunar maneuvering control, the target orbit can be re-integrated based on the optimized first and second tangential pulses to obtain a first intermediate orbit, and the orbital intersection point between the initial orbit and the recalculated first intermediate orbit can be calculated. The time when the orbital intersection point is reached is taken as the control timing for apolunar maneuvering control.

[0250] Furthermore, when determining the omnidirectional control pulse, the velocity vectors at the intersection points of the recalculated first intermediate orbit and the initial orbit can be determined, and the difference between the two velocity vectors can be used as the omnidirectional control pulse for lunar maneuver control.

[0251] After determining the control time, omnidirectional control pulse, first tangential pulse, and second tangential pulse for apolunar maneuvering control using a genetic optimization algorithm, a 6-to-6 differential correction algorithm can be used to accurately calculate the orbit control time and pulses for the three-pulse joint control based on the target orbit's orbital parameters. The variables to be corrected include: the control time of apolunar maneuvering control, the radial pulse of apolunar maneuvering control, the tangential pulse of apolunar maneuvering control, the normal pulse of apolunar maneuvering control, the first tangential pulse of capture control 2, and the second tangential pulse of capture control 3. The control times of capture control 2 and capture control 3 are both at the perilune. Further, the target parameters for correction are the orbital parameters of the target orbit, including: the perilune altitude, semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, and the perilune argument and mean anomaly (or true anomaly) at the target perilune (i.e., the perilune of capture control 3). Further, a 6-to-6 differential correction is performed on the variables based on the target parameters.

[0252] To demonstrate the feasibility of the solution proposed in the embodiments of this application, a simulation verification process for trajectory change control is presented as follows:

[0253] (1) Input the initial trajectory

[0254] The orbit is set as a highly elliptical orbit with a perigee altitude of 200 km and an orbital period of 3 days. Orbital epoch time: T1 = 2024-03-25T02:00:00.000, semi-major axis a = 20286.120 km, eccentricity e = 0.904447, inclination i = 99.0°, right ascension of ascending node Ω = 273.0°, argument of perigee ω = 103.0°, and mean anterior angle M = 0°.

[0255] (2) Determine the target trajectory

[0256] Orbital epoch time: TE = 2024-03-29T20:00:00.000. Periphery altitude hp = 800.000km, semi-major axis a = 9750.750km, inclination i = 95.0°, right ascension of ascending node Ω = 200.0°, argument of periphery ω = 85.0°, mean anterior angle M = 0°.

[0257] (3) Solving for coarse initial values ​​of tangential pulse

[0258] Based on the initial orbit and the target orbit, the following can be calculated:

[0259] The total flight time from the first perigee to the target time is 4.75 days = 410,400 seconds. The orbital period after the execution of capture control 1 is equal to the orbital period before the execution of capture control 2. The orbital period after capture control 2 is equal to the orbital period before capture control 3 is executed. The initial coarse value Δv of the first tangential control pulse of capture control 2 is calculated. 2t = -27.704 m / s, the coarse initial value Δv of the second tangential pulse of capture control 3 3t = -42.191m / s.

[0260] (4) Genetic optimization algorithm to solve for the initial value of the cubic pulse

[0261] Based on the initial values ​​of the tangential control pulses of capture control 2 and capture control 3, the value range of the tangential pulse of capture control 2 in the genetic optimization algorithm is set to [0, -60], and the value range of the tangential pulse of capture control 3 is set to [0, -40]. The population size of the genetic optimization algorithm is 80, the number of generations is 10, the crossover coefficient is 0.5, and the mutation coefficient is 0.8.

[0262] Table 2 shows the results of the genetic optimization algorithm solution process. In Table 2, population 1 and population 2 are the tangential pulse solutions of capture control 3 and capture control 2, respectively, and they are normalized results in the interval [0, 1].

[0263] Table 2. Results of the Genetic Optimization Algorithm

[0264]

[0265] The results of the 10th generation of the genetic optimization algorithm were converted into actual velocity increments, and the results of solving the initial values ​​of the three-pulse orbital control were output as shown in Table 3.

[0266] Table 3 Results of solving the initial value of the three pulses using the genetic optimization algorithm.

[0267] Radial pulse (m / s) Tangential pulse (m / s) Normal pulse (m / s) LAM -20.709 -97.893 113.827 LOI2 - -12.199 - LOI3 - -53.645 -

[0268] (5) Differential correction algorithm accurately calculates the three-pulse control parameters

[0269] Table 4 shows the initial values ​​of the three pulses obtained by the genetic optimization algorithm and the accurate calculation of the three pulse control parameters using the differential correction algorithm. The calculation results are shown in Tables 4 and 5.

[0270] Table 4. Precise Calculation of Three Pulses Using Differential Correction Algorithm

[0271] Orbit control time Radial pulse (m / s) Tangential pulse (m / s) Normal pulse (m / s) LAM 2024-03-25T22：29∶25 -19.601 -97.768 113.582 LOI2 2024-03-27T18：04：39 - -12.614 - LOI3 2024-03-29T20：00：00 - -53.346 -

[0272] Table 5. Pre- and post-control track parameters for precise three-pulse control.

[0273]

[0274] Simulation results show that utilizing the matching characteristic of orbital period and flight time to solve for the coarse initial values ​​of the tangential pulses of capture control 2 and capture control 3 provides a reference for the interval setting of the tangential pulse population in the genetic optimization algorithm. The coarse initial values ​​for capture control 2 and capture control 3 are -27.704 m / s and -42.191 m / s, respectively, while the results of the genetic optimization algorithm are -12.199 m / s and -53.645 m / s, respectively. The sum of the velocity increments of the two are basically consistent, and their magnitude distribution direction is consistent. The genetic optimization algorithm population intervals [0, -40] and [0, -60] preset based on the coarse initial values ​​have a good preset effect. In addition, the tangential pulse Δv of capture control 2... 2t and the tangential pulse Δv of capture control 3 3t Using the population variable and the deviation of the orbital splicing position vector at the far-lunar maneuvering control point as the fitness value, a good and effective model was constructed for the genetic optimization algorithm to solve the three-pulse full-parameter orbital control problem. During simulation, the genetic optimization algorithm, after 10 generations of iterative calculations, yielded good initial values ​​for the three-pulse orbital control timing and velocity increments, and showed rapid convergence of the fitness value as the genetic generations progressed. Finally, a differential correction algorithm was used to accurately solve the problem, obtaining the parameters for each orbital change control, achieving precise entry into the target orbit.

[0275] Based on the same concept as the method described above, see [link to relevant documentation]. Figure 7 This application provides an embodiment of a lunar orbiter orbit control device 700. Device 700 is used to implement the various steps in the above method embodiments; to avoid repetition, these will not be described again here. Device 700 includes: a lunar maneuver control unit 701, a phase adjustment control unit 702, a period control unit 703, an optimization unit 704, and a correction unit 705.

[0276] The lunar maneuver control unit 701 is used to determine the control timing and omnidirectional control pulses for the spacecraft to change its orbit from the initial lunar orbit to the first intermediate orbit based on the orbital plane, orbital camber, and perilune altitude of the first intermediate orbit of the spacecraft's lunar orbit. The omnidirectional control pulses include the spacecraft's velocity increments in the radial, normal, and tangential directions.

[0277] The phase control unit 702 is used to determine the control time and the first tangential pulse for the spacecraft to change its orbit from the first intermediate orbit to the second intermediate orbit based on the orbital phase of the second intermediate orbit;

[0278] The period control unit 703 is used to determine the control time and the second tangential pulse for the spacecraft to change its orbit from the second intermediate orbit to the target orbit based on the orbital period of the target orbit of the lunar flight.

[0279] In some embodiments, the apparatus further includes an optimization unit 704, configured to:

[0280] The first intermediate track is obtained by inverse integration of the target track based on the first tangential pulse and the second tangential pulse;

[0281] Determine two intersection points between the orbital plane of the first intermediate orbit and the orbital plane of the initial orbit, and determine the target intersection point that is closest to the apogee of the two intersection points;

[0282] Based on the latitude argument of the target intersection point, the first intermediate track is inversely integrated and the initial track is integrated in the forward direction to determine the first position vector of the first intermediate track and the second position vector of the initial track;

[0283] The first tangential pulse and the second tangential pulse are optimized and solved based on the first position vector and the second position vector.

[0284] In some embodiments, the control time for the spacecraft to change its orbit from the first intermediate orbit to the second intermediate orbit is the time when the spacecraft reaches the lunar perigee of the first intermediate orbit; the control time for the spacecraft to change its orbit from the second intermediate orbit to the target orbit is the time when the spacecraft reaches the lunar perigee of the second intermediate orbit.

[0285] The optimization unit 704 is specifically used for:

[0286] The second intermediate trajectory is determined based on the target trajectory and the second tangential pulse;

[0287] Based on the orbit extrapolation algorithm, the second intermediate orbit is inversely integrated to the moment when it reaches the previous perigee;

[0288] The first intermediate track is determined based on the integrated second intermediate track and the first tangential pulse.

[0289] In some embodiments, the optimization unit 704 is specifically used for:

[0290] The first time point is obtained by inversely integrating the first intermediate trajectory to the latitude argument of the target intersection point, and the second time point is obtained by inversely integrating the initial trajectory to the latitude argument of the target intersection point.

[0291] Determine whether the time difference between the first moment and the second moment is less than a preset time deviation;

[0292] If it is less than, then determine the first position vector of the first intermediate track at the first time and the position vector of the initial track at the second time;

[0293] If the time difference is not less than the preset time deviation, then based on the integrated first intermediate track and the integrated initial track, return to the step of determining the intersection point of the track plane until the time difference is less than the preset time deviation.

[0294] In some embodiments, the optimization unit 704 is specifically used for:

[0295] Determine the deviation between the first position vector and the second position vector;

[0296] A genetic optimization algorithm is used to optimize the first tangential pulse and the second tangential pulse by using the deviation as the fitness value of the genetic optimization algorithm.

[0297] In some embodiments, the optimization unit 704 is specifically used for:

[0298] An initial population is generated based on preset boundary settings for the first tangential pulse and the second tangential pulse, respectively; wherein, the initial population contains multiple individuals, and each individual contains a first tangential pulse and a second tangential pulse;

[0299] For each individual in the initial population, the deviation is calculated, and individuals whose calculated deviation is less than a set threshold are designated as parent individuals.

[0300] Genetic optimization algorithms employ crossover and mutation operators to perform genetic operations on the parent individuals to generate a new population. Based on the new population, the steps for calculating the deviation are repeated until the number of genetic operations reaches a threshold. Finally, the optimal individual with the smallest deviation is output.

[0301] Determine the optimal first tangential pulse and the optimal second tangential pulse contained in the optimal individual.

[0302] In some embodiments, the optimization unit 704 is further configured to:

[0303] The first intermediate track is recalculated based on the first and second tangential pulses obtained from the optimized solution, and the intersection point of the initial track and the recalculated first intermediate track is calculated.

[0304] Based on the velocity vector of the initial track at the track intersection and the velocity vector of the re-solved first intermediate track at the track intersection, the omnidirectional control pulse for changing track from the initial track to the re-solved first intermediate track is determined;

[0305] The time when the orbital intersection is reached is taken as the control time for the orbital change from the initial orbit to the first intermediate orbit obtained by resolving the problem.

[0306] In some embodiments, the apparatus further includes a correction unit 705, configured to:

[0307] Based on the orbital parameters of the target orbit, a differential correction algorithm is used to solve for the control time, the omnidirectional control pulse, the optimized first tangential pulse, and the optimized second tangential pulse of the spacecraft changing orbit from the initial orbit to the first intermediate orbit obtained by resolving the algorithm. This is to ensure that the initial orbit reaches the target orbit after three orbital controls by the omnidirectional control pulse, the first tangential pulse, and the second tangential pulse. The total orbital parameters of the target orbit include the semi-major axis, perilune altitude, inclination, right ascension of the ascending node, argument of the perilune, and true anomaly at the target orbit time.

[0308] Figure 8 A schematic diagram of the structure of an electronic device 800 provided in an embodiment of this application is shown. The electronic device 800 in this embodiment may also include a communication interface 803, such as a network port, through which the electronic device can transmit data.

[0309] In this embodiment, the memory 802 stores instructions that can be executed by at least one controller 801. By executing the instructions stored in the memory 802, the at least one controller 801 can perform various steps in the above-described method. For example, the controller 801 can implement the above-described... Figure 7 The functions of each unit.

[0310] The controller 801 is the control center of the electronic device, capable of connecting various parts of the device via various interfaces and lines. It executes instructions stored in the memory 802 and retrieves data stored in the memory 802. Optionally, the controller 801 may include one or more processing units. The controller 801 may integrate an application controller and a modem controller. The application controller primarily handles the operating system and applications, while the modem controller primarily handles wireless communication. It is understood that the modem controller may not be integrated into the controller 801. In some embodiments, the controller 801 and the memory 802 may be implemented on the same chip; in other embodiments, they may be implemented on separate chips.

[0311] The controller 801 can be a general-purpose controller, such as a central controller (CPU), digital signal controller, application-specific integrated circuit, field-programmable gate array or other programmable logic device, discrete gate or transistor logic device, or discrete hardware component, capable of implementing or executing the methods, steps, and logic block diagrams disclosed in the embodiments of this application. The general-purpose controller can be a microcontroller or any conventional controller. The steps performed by the data statistics platform disclosed in the embodiments of this application can be directly executed by the hardware controller, or executed by a combination of hardware and software modules within the controller.

[0312] Memory 802, as a non-volatile computer-readable storage medium, can be used to store non-volatile software programs, non-volatile computer-executable programs, and modules. Memory 802 may include at least one type of storage medium, such as flash memory, hard disk, multimedia card, card-type memory, random access memory (RAM), static random access memory (SRAM), programmable read-only memory (PROM), read-only memory (ROM), electrically erasable programmable read-only memory (EEPROM), magnetic storage, magnetic disk, optical disk, etc. Memory 802 can be any other medium capable of carrying or storing desired program code in the form of instructions or data structures that can be accessed by a computer, but is not limited thereto. In the embodiments of this application, memory 802 can also be a circuit or any other device capable of implementing storage functions for storing program instructions and / or data.

[0313] By designing and programming the controller 801, for example, the code corresponding to the method described in the foregoing embodiment can be embedded into the chip, so that the chip can execute the steps of the foregoing method when running. How to design and program the controller 801 is a well-known technique to those skilled in the art, and will not be described in detail here.

[0314] 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.

[0315] 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 controller of a general-purpose computer, special-purpose computer, embedded processor, or other programmable data processing device to produce a machine, such that the instructions, which execute via the controller of the computer or other programmable data processing device, generate instructions for implementing the flowchart illustrations. Figure 1 One or more processes and / or boxes Figure 1 A device that provides the functions specified in one or more boxes.

[0316] These computer program instructions may also be stored in a computer-readable storage medium that can direct a computer or other programmable data processing device to function in a particular manner, such that the instructions stored in the computer-readable storage medium produce an article of manufacture including instruction means, which are implemented in a process Figure 1 One or more processes and / or boxes Figure 1 The function specified in one or more boxes.

[0317] These computer program instructions may also be loaded onto a computer or other programmable data processing equipment to cause a series of operational steps to be performed on the computer or other programmable equipment to produce a computer-implemented process, thereby providing instructions that execute on the computer or other programmable equipment for implementing the process. Figure 1 One or more processes and / or boxes Figure 1 The steps for the functions specified in one or more boxes.

[0318] Although preferred embodiments of this application have been described, those skilled in the art, upon learning the basic inventive concept, can make other changes and modifications to these embodiments. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments as well as all changes and modifications falling within the scope of this application.

[0319] 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 method for orbital control of a lunar orbiter, characterized in that, The method includes: Based on the orbital plane, orbital camber, and perigee altitude of the spacecraft's first intermediate orbit around the moon, the control timing and omnidirectional control pulses for the spacecraft to change its orbit from the initial orbit around the moon to the first intermediate orbit are determined; wherein, the omnidirectional control pulses include the spacecraft's velocity increments in the radial, normal, and tangential directions; Based on the orbital phase of the second intermediate orbit, the control time and the first tangential pulse for the spacecraft to change orbit from the first intermediate orbit to the second intermediate orbit are determined; Based on the orbital period of the target orbit for lunar orbiting, the control time and second tangential pulse for the spacecraft to change its orbit from the second intermediate orbit to the target orbit are determined; After determining the control time and tangential pulse for the spacecraft to change its orbit from the second intermediate orbit to the target orbit based on the orbital period of the target lunar orbit, the method further includes: The first intermediate track is obtained by inverse integration of the target track based on the first tangential pulse and the second tangential pulse; Determine the line of intersection between the orbital plane of the first intermediate orbit and the orbital plane of the initial orbit, and determine the target intersection point among the intersection points of the intersection line and the initial orbit that is closest to the apogee; Based on the latitude argument of the target intersection point, the first intermediate track is inversely integrated and the initial track is integrated in the forward direction to determine the first position vector of the first intermediate track and the second position vector of the initial track; The first tangential pulse and the second tangential pulse are optimized and solved based on the first position vector and the second position vector; The control time for the spacecraft to change its orbit from the first intermediate orbit to the second intermediate orbit is the time when the spacecraft reaches the lunar perigee of the first intermediate orbit; the control time for the spacecraft to change its orbit from the second intermediate orbit to the target orbit is the time when the spacecraft reaches the lunar perigee of the second intermediate orbit; the step of obtaining the first intermediate orbit by inverse integration of the target orbit based on the first tangential pulse and the second tangential pulse specifically includes: The second intermediate trajectory is determined based on the target trajectory and the second tangential pulse; Based on the orbit extrapolation algorithm, the second intermediate orbit is inversely integrated to the moment when it reaches the previous perigee; The first intermediate track is determined based on the integrated second intermediate track and the first tangential pulse; The optimization solution for the first tangential pulse and the second tangential pulse based on the first position vector and the second position vector specifically includes: Determine the deviation between the first position vector and the second position vector; A genetic optimization algorithm is used to optimize the first tangential pulse and the second tangential pulse by using the deviation as the fitness value of the genetic optimization algorithm.

2. The method according to claim 1, characterized in that, The step of determining the first position vector of the first intermediate track and the second position vector of the initial track by performing inverse integration on the first intermediate track and forward integration on the initial track based on the latitude argument of the target intersection point specifically includes: The first time point is obtained by inversely integrating the first intermediate trajectory to the latitude argument of the target intersection point, and the second time point is obtained by inversely integrating the initial trajectory to the latitude argument of the target intersection point. Determine whether the time difference between the first moment and the second moment is less than a preset time deviation; If it is less than, then determine the first position vector of the first intermediate track at the first time and the position vector of the initial track at the second time; If the time difference is not less than the preset time deviation, then based on the first intermediate track after inverse integration and the initial track after forward integration, return to the step of determining the intersection line of the track planes until the time difference is less than the preset time deviation.

3. The method according to claim 1, characterized in that, The step of employing a genetic optimization algorithm, using the deviation as the fitness value of the genetic optimization algorithm, to optimize the first tangential pulse and the second tangential pulse, specifically includes: An initial population is generated based on preset boundary settings for the first tangential pulse and the second tangential pulse, respectively; wherein, the initial population contains multiple individuals, and each individual contains a first tangential pulse and a second tangential pulse; For each individual in the initial population, the deviation is calculated, and individuals whose calculated deviation is less than a set threshold are designated as parent individuals. The genetic optimization algorithm uses crossover and mutation operators to perform genetic operations on the parent individuals to generate a new population. Based on the new population, the algorithm returns to the step of calculating the deviation until the number of genetic operations reaches a threshold. Finally, the optimal individual with the smallest deviation is output. Determine the optimal first tangential pulse and the optimal second tangential pulse contained in the optimal individual.

4. The method according to claim 1, characterized in that, After optimizing the solution for the first tangential pulse and the second tangential pulse based on the first position vector and the second position vector, the method further includes: The first intermediate track is recalculated based on the first and second tangential pulses obtained from the optimized solution, and the intersection point of the initial track and the recalculated first intermediate track is calculated. Based on the velocity vector of the initial track at the track intersection and the velocity vector of the re-solved first intermediate track at the track intersection, the omnidirectional control pulse for changing track from the initial track to the re-solved first intermediate track is determined; The time when the orbital intersection is reached is taken as the control time for the orbital change from the initial orbit to the first intermediate orbit obtained by resolving the problem.

5. The method according to claim 4, characterized in that, After setting the moment of reaching the orbital intersection as the control moment for the orbital change from the initial orbit to the first intermediate orbit obtained through re-solution, the method further includes: Based on all the orbital parameters of the target orbit, a differential correction algorithm is used to solve for the control time of the spacecraft's orbit change from the initial orbit to the first intermediate orbit obtained by resolving the algorithm, the omnidirectional control pulse, and the first and second tangential pulses obtained by optimization. This allows the initial orbit to reach the target orbit after three orbital controls: the omnidirectional control pulse, the first tangential pulse, and the second tangential pulse. The total orbital parameters of the target orbit include the semi-major axis, lunar altitude, inclination, right ascension of the ascending node, argument of the lunar perigee, and true anomaly at the target orbit time.

6. A lunar orbiter orbit control device, characterized in that, The device includes: The lunar maneuver control unit is used to determine the control timing and omnidirectional control pulses for the spacecraft to change its orbit from the initial lunar orbit to the first intermediate orbit, based on the orbital plane, orbital camber, and perilune altitude of the spacecraft's first intermediate orbit around the moon; wherein the omnidirectional control pulses include the spacecraft's velocity increments in the radial, normal, and tangential directions; The phase control unit is used to determine the control time and the first tangential pulse for the spacecraft to change its orbit from the first intermediate orbit to the second intermediate orbit based on the orbital phase of the second intermediate orbit; A period control unit is used to determine the control time and the second tangential pulse for the spacecraft to change its orbit from the second intermediate orbit to the target orbit based on the orbital period of the target orbit for lunar orbiting. The device further includes an optimization unit for: The first intermediate track is obtained by inverse integration of the target track based on the first tangential pulse and the second tangential pulse; Determine the line of intersection between the orbital plane of the first intermediate orbit and the orbital plane of the initial orbit, and determine the target intersection point among the intersection points of the intersection line and the initial orbit that is closest to the apogee; Based on the latitude argument of the target intersection point, the first intermediate track is inversely integrated and the initial track is integrated in the forward direction to determine the first position vector of the first intermediate track and the second position vector of the initial track; The first tangential pulse and the second tangential pulse are optimized and solved based on the first position vector and the second position vector; The control time for the spacecraft to change its orbit from the first intermediate orbit to the second intermediate orbit is the time when the spacecraft reaches the lunar perigee of the first intermediate orbit; the control time for the spacecraft to change its orbit from the second intermediate orbit to the target orbit is the time when the spacecraft reaches the lunar perigee of the second intermediate orbit; the optimization unit is further configured to: The second intermediate trajectory is determined based on the target trajectory and the second tangential pulse; Based on the orbit extrapolation algorithm, the second intermediate orbit is inversely integrated to the moment when it reaches the previous perigee; The first intermediate track is determined based on the integrated second intermediate track and the first tangential pulse; The optimization unit is further configured to: Determine the deviation between the first position vector and the second position vector; A genetic optimization algorithm is used to optimize the first tangential pulse and the second tangential pulse by using the deviation as the fitness value of the genetic optimization algorithm.

7. The apparatus according to claim 6, characterized in that, The optimization unit is specifically used for: The first time point is obtained by inversely integrating the first intermediate trajectory to the latitude argument of the target intersection point, and the second time point is obtained by inversely integrating the initial trajectory to the latitude argument of the target intersection point. Determine whether the time difference between the first moment and the second moment is less than a preset time deviation; If it is less than, then determine the first position vector of the first intermediate track at the first time and the position vector of the initial track at the second time; If the time difference is not less than the preset time deviation, then based on the first intermediate track after inverse integration and the initial track after forward integration, return to the step of determining the intersection line of the track planes until the time difference is less than the preset time deviation.

8. The apparatus according to claim 6, characterized in that, The optimization unit is specifically used for: An initial population is generated based on preset boundary settings for the first tangential pulse and the second tangential pulse, respectively; wherein, the initial population contains multiple individuals, and each individual contains a first tangential pulse and a second tangential pulse; For each individual in the initial population, the deviation is calculated, and individuals whose calculated deviation is less than a set threshold are designated as parent individuals. The genetic optimization algorithm uses crossover and mutation operators to perform genetic operations on the parent individuals to generate a new population. Based on the new population, the algorithm returns to the step of calculating the deviation until the number of genetic operations reaches a threshold. Finally, the optimal individual with the smallest deviation is output. Determine the optimal first tangential pulse and the optimal second tangential pulse contained in the optimal individual.

9. The apparatus according to claim 6, characterized in that, The optimization unit is further configured to: The first intermediate track is recalculated based on the first and second tangential pulses obtained from the optimized solution, and the intersection point of the initial track and the recalculated first intermediate track is calculated. Based on the velocity vector of the initial track at the track intersection and the velocity vector of the re-solved first intermediate track at the track intersection, the omnidirectional control pulse for changing track from the initial track to the re-solved first intermediate track is determined; The time when the orbital intersection is reached is taken as the control time for the orbital change from the initial orbit to the first intermediate orbit obtained by resolving the problem.

10. The apparatus according to claim 9, characterized in that, The device further includes a correction unit for: Based on all the orbital parameters of the target orbit, a differential correction algorithm is used to solve for the control time of the spacecraft's orbit change from the initial orbit to the first intermediate orbit obtained by resolving the algorithm, the omnidirectional control pulse, and the first and second tangential pulses obtained by optimization. This allows the initial orbit to reach the target orbit after three orbital controls: the omnidirectional control pulse, the first tangential pulse, and the second tangential pulse. The total orbital parameters of the target orbit include the semi-major axis, lunar altitude, inclination, right ascension of the ascending node, argument of the lunar perigee, and true anomaly at the target orbit time.

11. An electronic device, characterized in that, The electronic device includes a controller and a memory. The memory is used to store computer programs or instructions; The controller is configured to execute a computer program or instructions in a memory, such that the method of any one of claims 1-5 is performed.

12. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores computer-executable instructions that, when invoked by a computer, cause the computer to perform the method as described in any one of claims 1-5.

13. A computer program product, characterized in that, The computer program product includes: computer program code, which, when run on a computer, causes the computer to perform the method described in any one of claims 1-5.

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