36 results about "Orbit+Face" patented technology

The invention discloses a satellite network laser inter-satellite link establishment method and system, electronic equipment and a storage medium, and mainly solves the problem of low availability of a satellite network established in the prior art due to sun-sun influence. According to the scheme, the method comprises the following steps: inputting constellation configuration parameters of a satellite network, and calculating satellites which are uniformly distributed on each orbit surface; according to the input constellation configuration parameters, establishing a same-orbit inter-satellite laser link between all satellites in the constellation and adjacent satellites in the same orbit plane where the satellites are located; according to the constellation configuration parameters, calculating boundary phase factors used for judging different-orbit link connection modes; and according to the phase factor of the constellation and the boundary phase factor, establishing a different-orbit laser inter-satellite link between satellites on all adjacent orbits, and completing the construction of the whole-network inter-satellite laser link. According to the method, the optimal link connection configuration can be screened out in the satellite network construction stage, the availability of the satellite network is improved, and the method can be used for inter-satellite link sun-sun avoidance.

The application aims to provide a remote sensing satellite descending node local time control method based on orbit height adjustment, which has simple control and high control precision. The method comprises the following steps: S1, acquiring orbit elements of a remote sensing satellite, wherein the orbit elements are instantaneous elements or two-line elements; S2, calculating a current descending node local time of the satellite based on the orbit elements; S3, calculating a drift deviation of the current descending node local time from a nominal descending node local time △T ; S4, calculating a current orbit plane rotation rate based on a J2 perturbation parameter of the earth; S5, calculating a difference between the current orbit plane rotation rate and an earth revolution rate; S6, designing an expected orbit plane rotation rate based on the difference calculated in the step S5; S7, calculating an expected orbit height according to the expected orbit plane rotation rate; and S8, adjusting the orbit height of the satellite to the expected orbit height. The application is applied to the technical field of remote sensing satellite measurement and control.

The application relates to a rotation error correction method of a distributed autonomous navigation constellation, and comprises the following steps: S1, each satellite calculates the orbit plane orientation error of the satellite in orbit and distributes the result to all visible satellites in the constellation in real time; S2, each satellite fits a local constellation rotation quantity according to the orbit plane orientation error of the satellite and the collected orbit plane orientation error; S3, each satellite determines a virtual satellite orbit observation quantity according to the local constellation rotation quantity of the satellite; S4, each satellite constructs a conditional constraint equation of virtual observation according to the virtual satellite orbit observation quantity and the second type of non-singularity orbit root number; and S5, each satellite solves the satellite orbit through filtering according to the conditional constraint equation of virtual observation, so that the overall rotation effect of the constellation is inhibited and corrected. The application can inhibit the rotation effect of the constellation, improve the distributed autonomous orbit determination precision of the satellite, does not increase the inter-satellite data interaction communication pressure, and can stably maintain the constellation configuration for a long time.

The application discloses an initialization method of a conic surface under a projection linearization track surface, and belongs to the technical field of calculation, reckoning or counting. The method first obtains a preliminary estimation of a target track through three groups of observation data; determines a conic surface containing an initial track surface according to initial track information and geometric knowledge; estimates the missing height information in combination with the conic surface, and obtains more position information of the target in combination with the existing angle information; projects the position information of the target into a two-dimensional section, re-estimates the track parameters by using a least square algorithm, and more accurate track parameters can be obtained; then the iteration process is repeated by using the newly obtained track information until a certain initial orbit determination accuracy requirement is met. The algorithm firstly fits the height information of the target by using a conic surface coordinate system, and realizes linear fitting of the position vector by using a track surface projection method, so that the calculation amount is small and the accuracy is high.

The application discloses a heterogeneous satellite constellation configuration correction method based on stable earth coverage performance. The method comprises the following steps: constructing a first query table and a second query table; for each orbit plane, using a working orbit parameter optimization method, combining the first query table and the second query table, analyzing and calculating dynamic visibility conditions and a visibility list, and then correcting the working semi-major axis and the latitude argument AoL of each satellite on the orbit plane; for each orbit plane, according to the working semi-major axis correction value of the satellite on the orbit plane, combining the first query table, calculating the right ascension of the ascending node RAAN correction value of the satellite; and the working semi-major axis correction value, the latitude argument AoL correction value and the right ascension of the ascending node RAAN correction value of each satellite on all orbit planes constitute the configuration parameters of the heterogeneous satellite constellation. The application restrains the drift of the earth coverage index of the heterogeneous satellite constellation, reduces the constellation control frequency, and avoids the additional increase of the number of satellites due to the drift of the coverage index and the multiple artificial iterative design.

The invention discloses a two-way rocket launching window searching method, system and device and a medium, and the method comprises the steps: rapidly calculating two nearest launching windows of an ascending orbit and a descending orbit of a rocket according to the attributes of a central body, target orbit plane parameters, the geographic coordinates of launching points and the flight duration of the rocket; according to the method, the problem is converted into numerical calculation based on space geometry, the crossing moment is directly solved by using the geometrical relationship between the normal vector of the orbital plane and the inertial position vector of the launching point location and adopting efficient numerical methods such as dichotomy and the like, and complex dynamic iteration is avoided. According to the method, the calculated amount is greatly reduced, rapid searching of the launching window is achieved, the method is suitable for an interstellar detection task scene needing rapid response or multiple times of calculation, and the real-time performance of task planning is remarkably improved.

The application discloses an antenna switching method for inter-satellite networking. When the included angle between the direction vector of any antenna in a closed state and the direction vector of the opposite satellite is less than or equal to a preset value within a continuous specified time length, the antenna is opened, and the currently opened antenna is closed.

The application discloses a kind of unloading method and device of asymmetric configuration satellite angular momentum, the method includes: by solar sensor, the solar vector under satellite body system is acquired, and the solar vector is converted into the solar vector under orbit system;The included angle β of the solar vector under the orbit system and orbit plane is calculated, and the rotation angle of solar wing rotation axis and the preset angle of biasing axis are set according to the included angle β of the solar vector under the orbit system and orbit plane;The angular momentum of flywheel system is calculated and is converted into the angular momentum under satellite body system;According to the modulus of satellite to be unloaded angular momentum and preset unloading threshold value are compared, according to the comparison result, it is judged whether magnetic moment device unloading needs to be started, and according to the comparison result, the control instruction of magnetic moment device is output.The application effectively reduces angular momentum accumulation and space moment accumulation;With the advantages of saving fuel, low power consumption requirement and high attitude stability, it provides a theoretical basis for angular momentum unloading of weak magnetic field planet explorer.

The application discloses a method for calibrating a roll angle of an emergency anti-solar time of a satellite on the earth, which is beneficial to reestablishing a three-axis attitude of the satellite on the earth. Firstly, an angle between a projection of a sun vector on an orbit plane and an orbit system-X axis is calculated; then, whether a current orbit position meets a requirement is determined according to the angle between the projection of the sun vector on the orbit plane and the orbit system-X axis, and whether an absolute attitude reference is available at present is determined, if the current orbit position meets the requirement and the absolute attitude reference is available at present, the roll angle is calibrated at a current control beat; otherwise, the roll angle is not calibrated at the current control beat; if the roll angle is calibrated at the current control beat, the roll angle is calculated based on the absolute attitude reference; if the roll angle is not calibrated at the current control beat, the roll angle is calculated by using gyro inertia angular velocity integration. The application can provide high-precision and reliable attitude information for reestablishing the attitude of the satellite on the earth.

The invention discloses an automatic division method and system for a satellite constellation orbital plane, and relates to the technical field of spaceflight measurement and control. The method comprises the following steps: firstly, calculating an orbital plane unit normal vector based on a satellite orbit inclination angle and an ascending node right ascension, and performing first-level spatial pointing clustering by adopting a density clustering algorithm to form an orbital plane group; secondly, in the same orbital plane group, by taking a semi-major axis and an eccentricity rate as characteristics, automatically determining an optimal shell layer number by utilizing a Gaussian mixture model and a Bayesian information criterion, and realizing second-stage orbital energy shell layer division; and finally, in the same shell layer, sorting and interval statistical analysis are performed on the right ascension of the ascending node, recursive subdivision is performed through dynamic threshold segmentation in combination with a range-based dual-condition verification mechanism, and fine division of a third-level orbital plane is completed. The method realizes full-process automation from orbit data to division results, has the advantages of high division precision, strong adaptability, good engineering practicability and the like, and is suitable for operation and maintenance management of large-scale heterogeneous satellite constellations.

The application discloses a method for quickly changing orbit of an elliptical orbit out-of-plane target to the same orbit with the least fuel control on a satellite, and a method for independently solving the least fuel control on a satellite for quickly changing orbit from an out-of-plane orbit to the same orbit, which comprises the following steps: judging whether to perform normal orbit plane correction or orbit shape correction according to orbit elements of an initial satellite and a target; solving a least fuel normal orbit plane correction control strategy according to a normal orbit plane element difference between the satellite and the target; solving a least fuel orbit shape correction control strategy according to an in-orbit plane element difference between the satellite and the target; and iteratively correcting the least fuel control strategy by using a high-precision orbit extrapolation model. The method is suitable for orbit changing of any elliptical orbit satellite in near-earth space to any elliptical orbit target, and the control precision can reach ten meters; and the calculation amount is small, and the method is suitable for solving a task planning on a satellite.

This invention discloses a method for initial orbit determination of a non-cooperative target under space-based angle-only extremely short arc observation, belonging to the technical field of initial orbit determination. This invention addresses the failure of existing initial orbit determination methods under extremely short arc conditions. The method includes: using the unit normal vector of the target orbital plane as the variable to be optimized; obtaining the relationship between the slant distance and the unit normal vector based on the geometric relationship between the target position vector, the slant distance between the space-based platform and the target position, the target azimuth unit vector, and the space-based platform position vector; transforming the solution of the unit normal vector into the solution of the X-axis and Y-axis components of the unit normal vector based on the normalization characteristics of the unit normal vector, and determining the constraints of the solution; and employing a heuristic algorithm to solve for the X-axis and Y-axis components of the unit normal vector, thereby determining the unit normal vector and achieving initial orbit determination. This invention improves the accuracy and robustness of initial orbit determination.

The invention provides a sun-synchronous orbit inclination angle adjusting method and device, and the method comprises the steps: determining a position relation between an initial orbit and a target orbit, enabling two intersection points of the initial orbit and the target orbit to be located above an equator, and taking one of the intersection points as a satellite orbit transfer point; determining an orbital transfer angle required for adjusting the satellite from the initial orbital plane to the target orbital plane; according to the system state of the satellite and the task target, the single ignition duration of the satellite and the ignition moment and shutdown moment of the satellite are comprehensively determined; determining a satellite ignition direction; the speed increment of satellite monopulse ignition is analyzed and calculated, and the total impulse required for changing the inclination angle of the satellite orbit is determined according to the speed increment; the total satellite ignition orbit number is analyzed and calculated according to the total impulse and the actual impulse generated by each time of ignition; and based on the satellite orbital transfer point, the satellite ignition direction, the satellite ignition time and the shutdown time, the satellite is controlled to carry out secondary ignition orbital transfer, and then inclination angle adjustment is completed.

This invention provides a method and apparatus for adjusting the inclination of a sun-synchronous orbit. The method includes: determining the positional relationship between an initial orbit and a target orbit, such that the two intersection points of the initial orbit and the target orbit are located above the equator, and using one of the intersection points as the satellite's orbit change point; determining the orbit change angle required for the satellite to adjust from the initial orbital plane to the target orbital plane; comprehensively determining the satellite's single ignition duration, ignition time, and shutdown time based on the satellite's system status and mission objectives; determining the satellite's ignition direction; analytically calculating the velocity increment of the satellite's single pulse ignition, and determining the total impulse required to change the satellite's orbit inclination based on the velocity increment; analytically calculating the total number of satellite ignitions based on the total impulse and the actual impulse generated by each ignition; and controlling the satellite to perform a single ignition orbit change to complete the inclination adjustment based on the satellite's orbit change point, ignition direction, ignition time, and shutdown time.

The invention discloses a three-dimensional deduction table supporting space orbit control in a space-ground cooperation mode. The three-dimensional deduction table is characterized in that a map placing area, an arc-shaped spacing line used for representing the orbit height of 200-36000 km, an inner annular pulley orbit and an outer annular pulley orbit are arranged on a table top of the three-dimensional deduction table; the space plate frame comprises an inner ring plate frame and an outer ring plate frame which are used for representing a top-crossing track surface, the inner ring plate frame is erected on the inner ring-shaped pulley track and can rotate along the inner ring-shaped pulley track, and an arc-shaped spacing line used for representing the track height of 200-2000 km is marked on a transparent plate surface; the outer ring plate frame is erected on the outer ring-shaped pulley track and can rotate along the outer ring-shaped pulley track, and an arc-shaped spacing line for representing the track height of 2000-36000 km is marked on a transparent plate surface; the operators are used for space-ground collaborative deduction and comprise a ground operator used for a map placement area and a space operator used for a horizontal track area and a space grillage.

This invention discloses a large-scale space target collision early warning combination screening method, belonging to the field of space situational awareness. The implementation method is as follows: Preliminary screening of similar targets is performed based on the distance between intersection points on the orbital plane intersection line to quickly reduce the number of satellite combinations to be analyzed; by decomposing the analytical solution considering the minimum relative distance between satellites under J2 perturbation, it is divided into long-term terms and exact solutions; based on the evolution characteristics of the long-term terms, a second screening is performed using the trend characteristics of the long-term terms; a third screening is performed using the analytical exact solution considering the relative distance between satellites under J2 perturbation; based on the results of the third screening, satellite combinations with collision risk are determined, realizing large-scale space target collision early warning combination screening, thereby protecting the safety of space assets. This invention fully considers the long-term evolution characteristics of orbits under perturbation, significantly reducing the calculation time for early warning, and is suitable for rapid screening of long-term collision early warnings for large-scale space targets.

The invention discloses a multi-satellite cooperative computing and task scheduling method and system suitable for constellation scale evolution, and the method comprises the steps: obtaining heterogeneous state data of each satellite based on an orbital plane state view under a current time slot, carrying out the unified feature coding, obtaining the feature representation of each satellite, and aggregating the feature representation into an orbital plane-level global state feature; segmenting and coding each to-be-decomposed task according to the global state characteristics to obtain a sub-task decomposition frequency, a task decomposition scheme and an action vector for executing a satellite allocation scheme of each to-be-decomposed task; calculating the task completion rate of each priority group according to the action vector and the task priority to determine a system-level reward value; and updating each satellite state according to the computing resources correspondingly used by each to-be-decomposed task under the current time slot to generate a next time slot orbital plane state view, returning to execute the step of obtaining the heterogeneous state data until all time slots are iterated, and obtaining a task scheduling scheme based on the action vector corresponding to each to-be-decomposed task under each time slot.

The application discloses a high-precision space target orbit prediction method based on numerical fitting, and relates to the technical field of space target orbit prediction.The frequency domain decomposition of the application is combined with a harmonic compensation model to completely eliminate seasonal system drift caused by the gravitation of the sun and the moon; high-frequency components are stripped of attitude noise through light pressure parameterization and smooth difference to avoid the transmission of perturbation coupling error; a harmonic order adaptive mechanism captures the main disturbance term with the least amount of calculation; least square batch estimation realizes one-time solution of coefficients; template library preloading technology effectively compresses the time consumption of new target initialization; in addition, the application introduces a geomagnetic modulation factor to dynamically suppress compensation distortion during magnetic storms, and a sliding window mechanism guarantees real-time updating of coefficients, and stable prediction is still maintained under extreme events; photometric curve inversion is adopted instead of attitude measurement to make non-cooperative target prediction possible; tangential component directional compensation ensures the stability of the orbit plane and reduces satellite position fuel consumption.

The invention relates to the technical field of atmospheric science and remote sensing, in particular to a space-based distributed near space atmosphere detection method, system, device and product, and the method comprises the steps: designing a plurality of satellites for detecting near space atmosphere based on the number of orbital planes and the number of satellites included in each orbital plane; detecting the near space atmosphere at a target moment based on each satellite, and obtaining state parameters of the near space atmosphere from different angles; obtaining a tomographic equation based on the global three-dimensional grid diagram and the obtained state parameters; and solving the tomographic equation to obtain the three-dimensional distribution characteristics of the near space atmosphere at the target moment. A plurality of satellites are designed through an orbital plane, various observation modes can be effectively compatible, and multi-dimensional data of near space atmosphere can be obtained; the state parameters are processed by using the global three-dimensional grid diagram so as to extract key three-dimensional parameter information, and the state and the change process of the near space atmosphere can be accurately monitored and predicted.

The application relates to a minimum relative image movement imaging method of a low-orbit remote sensing satellite on a high-orbit target, and solves the problem of poor imaging effect on the high-orbit target. The application is convenient for the low-orbit remote sensing satellite to perform long-exposure imaging on the high-orbit target under the condition that the low-orbit remote sensing satellite is in an inertial space stable attitude. The method is as follows: the low-orbit remote sensing satellite observes a high-orbit target to be observed at an angle alpha, the relative speed of the high-orbit target to be observed relative to the low-orbit remote sensing satellite is zero in the projection of the relative direction normal; the high-orbit target to be observed is located on the orbit plane of the low-orbit remote sensing satellite, and the normal is located on the orbit plane; the angle alpha satisfies:

This invention discloses a multi-satellite collaborative computing and task scheduling method and system applicable to constellation-scale evolution. The method includes: acquiring heterogeneous state data of each satellite based on the orbital plane state view of the current time slot, performing unified feature encoding to obtain the feature representation of each satellite, and aggregating them into a global state feature at the orbital plane level; segmenting and encoding each task to be decomposed according to the global state feature to obtain an action vector including the number of sub-task decompositions, the task decomposition scheme, and the satellite allocation scheme for each task to be decomposed; calculating the task completion rate of each priority group based on the action vector and task priority to determine the system-level reward value; updating the state of each satellite according to the computing resources used by each task to be decomposed in the current time slot to generate the orbital plane state view of the next time slot, and then returning to the step of acquiring heterogeneous state data, until all time slots are iterated, and obtaining a task scheduling scheme based on the action vector corresponding to each task to be decomposed in each time slot.

A satellite constellation (110) includes a plurality of artificial satellites flying on an inclined circular orbit for each of six or more orbital planes having a common orbital inclination and arranged so that south and north axes are mutually shifted in an east-west direction. The plurality of artificial satellites include eight or more artificial satellites for each orbital plane. Each of the artificial satellites includes a fore-and-aft communication device. For each orbital plane, each of the artificial satellites forms a communication network with front and rear artificial satellites by the fore-and-aft communication device. Each of the artificial satellites on each orbital plane crosses southern and northern edges of the orbital plane in synchronization with artificial satellites on other orbital planes, and forms a communication network with that artificial satellites by the fore-and-aft communication device.

The application provides a collision risk satellite combination processing sequencing optimization method, steps are as follows: according to TLE orbit root number data, orbit pre-division is carried out, satellites located on the same orbit plane are divided together; a perigee-apogee pre-screening method is used to carry out pre-screening of satellite collision combination, and satellite and debris combinations generating collision risks are obtained; TLE orbit root number data is combined with an orbit prediction model and a fast screening trajectory space intersection algorithm to determine the position of trajectory intersection and overlap in space; high collision risk satellite combinations and high collision risk regions are determined; a collision combination priority processing sequencing model is constructed in combination with satellite states and satellite characteristics; and a gradient descent method is used to optimize parameters of the multi-factor collision combination priority processing sequencing model. The application improves the screening speed of high collision risk combinations and improves the overall safety of a satellite constellation by using the geometric screening method of the intersection angle between two orbits and the gradient descent method to optimize multi-factor parameters.

The application discloses a LEO orbit single-particle effect inversion method, device, medium and product, relates to the field of space environment monitoring and evaluation, and comprises the following steps: training and optimizing the first and second LSTM neural networks simultaneously by using a training data set to obtain a classification model and a regression model; inputting the average value of the space weather parameters of each hour into the classification model and the regression model simultaneously by pushing the current position and the current time forward by a set number of hours; multiplying the value output by the classification model and the particle flux prediction value of the current position output by the regression model to obtain the final particle flux prediction value of the current position; obtaining the LEO orbit plane particle flux according to the final particle flux prediction value of each position on the LEO orbit plane to be predicted; and obtaining the LEO orbit plane single-particle flip rate according to the LEO orbit plane particle flux. The application can accurately predict the LEO orbit plane particle flux and the LEO orbit plane single-particle flip rate.

The patent discloses a kind of to ellipse orbit same plane target fast companion fly around fly the most fuel-efficient control on-board autonomous planning method, from the same orbit plane fast approach, high-precision companion fly / around fly / sweep fly fuel-saving control on-board autonomous solution method at long distance, comprising: a kind of 3 planning, 5 maneuvering long-distance approach companion fly / around fly / sweep fly configuration formation method is proposed;Propose a kind of two-body model is solved the most fuel-efficient control analytical solution initial value suitable for on-board autonomous solution, high-precision orbit extrapolation model iterative accurate solution final value efficient iteration method.The method is suitable for the fast approach companion fly / around fly / sweep fly of any elliptical orbit satellite to long-distance target on approximate same orbit in near-earth space, and control accuracy can reach ten meters level;And small amount of calculation is suitable for on-board autonomous task planning solution.

The invention relates to a satellite formation configuration control method based on inter-satellite measurement information, and the method comprises the steps: obtaining the distance and direction information of all slave satellites in a local coordinate system of a master satellite, and directly solving the position vector information of the slave satellites relative to the master satellite according to a geometrical relationship; performing data fitting processing on the position vector information of the slave satellite relative to the master satellite, and calculating an average relative orbital element in combination with a relative motion equation of the slave satellite relative to the master satellite; determining a target average relative orbit element number according to the target satellite formation configuration, and determining an average relative orbit element number deviation based on the average relative orbit element number and the target average relative orbit element number; and the average relative orbital element deviation is corrected, so that the relative motion amplitude inside and outside the orbital plane of the slave satellite relative to the master satellite and the long-term relative motion in the orbital trace direction are regulated and controlled, and the configuration maintenance and reconstruction control of the satellite formation are carried out. The satellite formation configuration control method can be suitable for the satellite formation without a navigation system.

The invention relates to an orbit control method for different orbit plane low-thrust satellite formation rapid networking. Relates to the field of near-earth orbit satellite formation, in particular to an orbit control method for different-orbit-plane low-thrust satellite formation rapid networking. The control quantity which must be completed through thrust is converted into the drift distance which can be naturally completed through the orbit dynamics process, correction is carried out with the assistance of a small amount of direct control, and therefore the total control frequency and fuel consumption are remarkably reduced. The method comprises the following steps: acquiring a formation flying target; relevant parameters required by orbit control are calculated; constructing an ascending node right ascension control change model, and establishing an ascending node right ascension change equation; solving an ascending node right ascension change equation to obtain a minimum control circle number and a standard control vector; calculating a phase drift rate according to the adjustment time required by the right ascension of the ascending node; and combining the control vectors to form a final control plan.

The application discloses a method for constructing a cross-orbit chain topology of a mega-constellation network, and the method comprises the following steps: S1, acquiring the number of uniformly distributed orbit planes in the mega-constellation network P and the number of uniformly distributed satellites in each orbit plane S ; S2, each satellite on each orbit plane establishes an inter-orbit satellite interlink with the satellites adjacent to the orbit plane, meanwhile, each satellite on each orbit plane establishes an inter-orbit satellite interlink with the corresponding satellite on the corresponding orbit plane by selecting at least two cross numbers to form a cross-orbit chain topology of the mega-constellation network, and the greatest common divisor of all cross numbers is 1. The cross-orbit chain topology of the application makes the inter-orbit satellite interlink span the satellites of several orbit planes, under the premise of ensuring global connectivity, significantly improves the long-distance transmission efficiency, breaks through the scale bottleneck of a single-layer constellation, reduces the dependence on the ground gateway station, and makes the mega-constellation network become a high-performance infrastructure complementary to the ground network.

This invention provides a polar-orbiting satellite navigation constellation system, a navigation method, and a user terminal. The polar-orbiting satellite navigation constellation system includes: four orbital planes, each with its intersection point with the equator evenly distributed along the equator, with a longitude interval of 45 degrees between adjacent orbital planes; nine satellites evenly distributed on each orbital plane, with a phase interval of 40 degrees between adjacent satellites; each satellite orbits at an altitude of 80,000–90,000 km and an orbital inclination of 90 degrees. It also employs an isotope power supply system and a redundant cold / hot backup system in the attitude and orbit control system. This invention utilizes a high-orbit scheme that avoids the influence of the Van Allen radiation belts, effectively solving the radiation effects caused by the Earth's radiation belts and increasing the reliability of the navigation system; the use of an isotope power supply system improves the operational lifespan of the navigation system; and the cold / hot redundant backup of the attitude and orbit control system enhances the reliability of the control system.

The application discloses a transfer orbit fast optimization method based on spatial conic curve splicing, and belongs to the field of deep space exploration. The application expands the plane conic curve splicing method to the optimization of spatial earth-moon transfer orbit on the basis of the plane conic curve splicing method used for orbit transfer, by using coordinate conversion and geometric analysis method to increase the azimuth angle parameters theta and psi used for describing the spatial orientation of the earth-centered launch orbit plane; the application can improve the orbit analysis and optimization efficiency by using the two-body orbit theory to deduce the two-body model and based on the double two-body model splicing with analytical characteristics; the application can realize the fast optimization and solution of the transfer orbit based on the spatial conic curve splicing by establishing the penalty function iterative optimization of four independent optimization quantities, i.e., the earth-moon transfer launch speed increment, the moon influence ball entry point phase, the azimuth angle parameters theta and psi, according to the orbit inclination and the perigee height constraints. The application has the advantages of good convergence, high optimization efficiency and wide application range. The application is suitable for various earth-moon transfer missions and moon flyby orbit optimization.