42 results about "Right ascension" patented technology

The invention provides a star catalogue file generation method and device, a celestial body data query method and device, computer equipment and a storage medium. The star catalogue file generation method comprises the steps of obtaining star catalogue data; dividing the celestial sphere surface area corresponding to the plurality of celestial bodies into a plurality of local areas, and determining an identifier of each local area; determining an identifier matched with each celestial body data according to the right ascension and declination included in the celestial body data and the identifier of each local area in the celestial sphere surface area; sorting the celestial body data of the plurality of celestial bodies based on the identifiers corresponding to the celestial body data; and writing the sequenced celestial body data of the plurality of celestial bodies into an initial column storage file matched with the identifier to obtain a target file stored with the celestial body data, and generating file index information corresponding to the target file. The celestial body data is stored by utilizing the column type storage file, so that the problems of high maintenance cost and high use threshold caused by utilizing a database to store data are relieved while the data query efficiency is guaranteed.

The invention discloses an astronomical cross identification method based on spherical trapezoidal division, and the method comprises the steps: S1, taking a unit spherical surface as a celestial sphere, dividing the celestial sphere into a plurality of strips parallel to an equatorial line, and determining the right ascension span of each strip for the transverse segmentation of the strip; s2, dividing each strip into a same spherical isosceles trapezoid with two waists perpendicular to the equator, and taking the same spherical isosceles trapezoid as a sky area; s3, determining an index of each sky area; s4, dividing a star catalogue A and a star catalogue B, and creating a zipper type hash table based on the row number of the star catalogue B; s5, mapping each celestial body in the star catalogue B into a linked list in a hash table in a multi-thread parallel mode, and inserting celestial body data into the linked list by means of a two-stage insertion strategy; s6, mapping each celestial body in the star catalogue A into a linked list in a group of hash tables in a multi-thread parallel mode, and traversing to obtain celestial body IDs of which the spherical distance is smaller than the error radius; the method not only can completely avoid the problem of result omission, but also can greatly improve the astronomical cross identification calculation performance.

The invention provides a fixed star sight atmospheric refraction correction method for optical satellite limb observation, and the method comprises the steps: constructing an atmospheric ellipsoid model according to an earth atmosphere parameter set and earth ellipsoid parameters; according to the satellite orbit, the attitude measurement value, the image imaging time, the fixed star image plane coordinate and the satellite camera parameter, calculating the fixed star sight line; performing penetration layer atmospheric refraction correction on the fixed star sight line based on the atmospheric ellipsoid model to obtain a corrected first fixed star sight line; based on the atmosphere ellipsoid model, performing penetrating layer atmosphere refraction correction on the corrected first fixed star sight line to obtain a corrected second fixed star sight line; and calculating the right ascension and declination of the fixed star by using the corrected second fixed star sight line. According to the method, fixed star sight atmospheric refraction correction processing of optical satellite limb observation can be realized, so that support is provided for high-precision geometric processing of optical satellite limb observation images.

The invention discloses a strict regression frozen orbit design method. The method comprises the following steps: calculating an initial estimated value of a target strict regression frozen orbit mean element number based on an analytic formula of a regression orbit and a frozen orbit under a third-order gravity field; under the high-order gravity field dynamic model, iteratively correcting the orbit semi-major axis and the orbit inclination angle based on the right ascension change rate of the ascending node and the deviation value of the longitude of the revisit ascending node; taking the geocentric vector difference of the revisit ascending node as an optimization target, taking the eccentricity rate as an optimization variable, and adopting an optimization algorithm to obtain a current optimal frozen eccentricity rate; iteratively correcting the semi-major axis of the orbit and the inclination angle of the orbit, and optimizing the eccentricity ratio to obtain the orbital number of the strictly frozen regression orbit.

Disclosed are a method and an astrophotographic apparatus for acquiring images of targets in a sky area. The method includes: step 1, an image acquisition device is driven through a controlled rotatable component to point to the position near a target in a sky area, and images of the sky area are acquired through the image acquisition device; step 2, the images of the sky area are analyzed to obtain the right ascension coordinates and the declination coordinates of the center point of images for synchronizing the coordinates to the controlled rotatable component; step 3, the image acquisition device is driven by the controlled rotatable component to point to a target position of the target right ascension coordinates and the target declination coordinates corresponding to the target celestial body to acquire the images; and step 4, an image processing is performed on the images to obtain the processed images.

The present application relates to the technical field of altazimuth telescope, in particular to a pointing model calibration method of altazimuth telescope, the error function of the telescope is obtained by two-dimensional surface fitting, the star image is shot every 10 degrees or so in the azimuth and elevation directions, the right ascension and declination of the image center are obtained by comparing with the standard star map, the right ascension and declination are converted into the real azimuth and real elevation in the geodetic horizontal coordinate according to the shooting time, the geographical longitude and latitude of the telescope, the altitude, the air temperature and air pressure; the real azimuth and real elevation are subtracted by the azimuth and elevation read by the angle sensor of the telescope respectively, the azimuth error matrix and the elevation error matrix are obtained respectively; the error matrix is fitted to obtain the azimuth error function, the real azimuth and real elevation of any target are subtracted by the azimuth pointing error and the elevation pointing error respectively, and the accurate pointing of the telescope can be obtained.

The present invention relates to the field of celestial navigation technology, and specifically discloses a star map simulation method, system, computer device and computer-readable storage medium, the method comprising: partitioning the celestial sphere by right ascension and declination to obtain a plurality of first partitions of the celestial sphere; determining a second partition from all first partitions according to the optical axis direction and rectangular field of view of a star sensor, and constructing a minimized partition corresponding to each second partition; determining the stars in the imaging range of the star sensor in the minimized partition as the observed stars of the celestial sphere, and performing star map simulation based on all the observed stars. The present invention can narrow the discrimination range during star map simulation, reduce the number of discriminations, and thus improve generation efficiency.

The invention discloses an inter-satellite link fault avoidance method and system based on time quantitative calculation under the sun-sun effect, and mainly solves the problems that the calculation complexity of inter-satellite link faults of a satellite network is high and the link fault duration cannot be directly calculated under the influence of the sun-sun effect in the prior art. According to the implementation scheme, the method comprises the following steps: initializing constellation parameters, establishing a J2000.0 geocentric inertial coordinate system, and determining a satellite position at any moment; the right ascension and declination of the sun are calculated, the included angle between the sunlight and each satellite orbit plane is calculated according to the parameters, the link fault position existing in the orbit is determined, and the link fault range of each orbit is calculated; calculating positions and fault duration of all fault links in each orbit according to the link fault range and the positions of satellites; and the satellites autonomously perform link avoidance according to the link fault range and the fault duration. According to the method, the calculated amount can be effectively reduced, the response speed is improved, the duration of the link fault is directly obtained, and the method can be used for analyzing the inter-satellite link fault condition of the satellite network under the sun-sun effect.

The invention provides a fixed star map simulation method based on in-orbit measured data of a low-orbit moonlet, which comprises the following steps of: calculating orbit and attitude data of four corners of a star map CCD (Charge Coupled Device) according to row and column number interpolation of a to-be-simulated satellite star map image on the basis of satellite orbit, attitude and row time data measured in-orbit by the low-orbit moonlet in combination with a fixed star catalog; calculating right ascension and declination of angular points by combining internal and external orientation element information of a satellite, and selecting fixed star points falling into a star map in a star catalogue; the row and column numbers of the fixed star points in the star map are obtained by adopting a grid point bilinear interpolation method, and the gray scale information of the fixed star points is obtained by adopting a two-dimensional Gaussian distribution method. The star sky environment is further simulated by adding noise, and a simulation star map is output according to detailed information of result star points. According to the method, the simulation star map can be obtained quickly and efficiently, and the method has positive significance in improving the geometric calibration precision of the visible light load of the low-orbit moonlet through fixed star calibration.

The invention provides a star catalogue data merging method and device, computer equipment and a storage medium. The method comprises the steps of obtaining a target star catalogue file; according to right ascension and declination included in the celestial body data, determining serial number information of each piece of celestial body data under the target hierarchy, and dividing the celestial body data with the same serial number information in the target star catalogue file into the same data group to obtain a plurality of data groups; for each data group, determining an adjacent data group corresponding to the data group according to the number information corresponding to the data group; for each piece of celestial body data in the data group, determining a matching result between the celestial body data and other celestial body data in the data group, and a matching result between the celestial body data and each piece of celestial body data in the adjacent data group; and generating an index file corresponding to the target star catalogue file according to a matching result among the multiple pieces of celestial body data. According to the method, the whole table crossing and merging processing of the star catalogue file is realized, the number of repeated crossing times is reduced, and the star catalogue data crossing and merging efficiency is improved.

The application discloses an astronomical cross-identification method for a heterogeneous many-core processor, and comprises the following steps: S1, obtaining the strip id of two star catalogs based on declination and the multi-node data distribution in strip units; S2, determining the number of heterogeneous many-core processors of each computing node and the maximum number of threads that can be simultaneously run by each heterogeneous many-core processor; S3, rearranging the star catalog data in each strip in strip units, and dividing the star catalog data into grids based on right ascension; S4, judging whether the right ascension span of each strip is too large during internal identification, so as to select the identification calculation of the CPU end in step S5 in a multi-thread parallel mode or the identification calculation on the heterogeneous many-core processor in step S6; S7, determining the identification selected result; and S8, obtaining the ID of a star body in two star catalogs respectively. The method can guarantee the calculation accuracy and integrity, improve the parallelism of the cross-identification algorithm and the use efficiency of the heterogeneous many-core processor, and realize the improvement of the overall calculation performance of the astronomical cross-identification.

The utility model relates to a kind of hybrid reduction drive equatorial telescope based on cross roller bearing, including shell, right ascension shaft hybrid reduction drive assembly, declination shaft reduction drive assembly, cross roller bearing, laser beam parallelism adjusting mechanism, pedestal;Right ascension shaft hybrid reduction drive assembly includes right ascension shaft motor, right ascension primary reducer and right ascension secondary reducer, and right ascension primary reducer is harmonic reducer, and right ascension secondary reducer is worm gear transmission pair;Declination shaft reduction drive assembly includes declination shaft motor, declination shaft reducer;The outer ring of cross roller bearing is fixed on rear shell, the end surface of inner ring is fixed with the output end of right ascension shaft hybrid reduction drive assembly, and the other end surface of inner ring is fixed with right ascension output seat;Declination shaft reduction drive assembly is fixed on right ascension output seat.The utility model, with high degree of freedom, small, light weight, the characteristics of strong carrying capacity, and the reduction ratio of this equatorial telescope is high, tracking is stable, tracking precision is high, and power consumption is low.

The invention provides an observation constellation configuration control method and device based on a maximum allowable drift domain, and belongs to the technical field of universe navigation. According to the method, based on the observation performance of a target constellation on the surface of a central celestial body, a critical equivalent view field is obtained through geometric equivalence, and a maximum allowable drift domain is obtained through analysis and derivation. And determining a first maximum permissible drift distance (namely, a phase maximum permissible drift distance) and a second maximum permissible drift distance (namely, an ascending node right ascension maximum permissible drift distance) from the maximum permissible drift domain in an inscribed interception mode, and taking the first maximum permissible drift distance and the second maximum permissible drift distance as target constellation configuration control threshold values, so as to realize configuration control on the target constellation. According to the method, the calculation amount required for determining the maximum allowable drift distance is reduced, so that the precision of the maximum allowable drift distance is improved under the same calculation power, and the configuration control effect of the target constellation is improved.

The present invention discloses a method and system for automatically extracting very long baseline interferometry (VLBI) schedules and generating trajectory files. The method comprises the following steps: extracting information of specific stations from the observation schedule files issued in VLBI observations, including the coordinates of the source and the start and end times corresponding to each scan; calculating the right ascension and declination of the telescope between every two scans under the conditions of telescope speed and acceleration limits; combining the right ascension and declination of the observed source in each scan to generate the azimuth and elevation pointing of the specified station with an accuracy of 0.1 second; detecting data and generating an observation trajectory file. This method can achieve one-key generation from VLBI observation schedules to observation trajectory files. On the one hand, it reduces errors caused by human factors; on the other hand, it enables the telescope to be unrestricted by the observation mode when participating in VLBI observations.

A method for target attitude planning of a multi-field-of-view star observation satellite, comprising: Step 1, obtaining the season or month when the satellite conducts star observation, the satellite orbit altitude, and the stray light suppression angle of the imaging optical device; Step 2, giving the optional sky areas of each optical device according to the projection relationship between the optional fields of view of the optical device and the observable sky area; Step 3, selecting a set of observation points according to the star distribution in the optional sky area, obtaining the pointing of each optical device in the celestial coordinate system under this observation point, judging whether it is in its available sky area, and deleting the unavailable observation points; Step 4, counting the number of stars in the camera field of view under the available observation points and optimizing the available observation points; Step 5, repeating Steps 3 and 4 until a satisfactory set of observation points is found, and outputting the right ascension and declination of the optimal observation point and the satellite attitude transformation matrix. The present invention provides an ideal observation sky area and target attitude for remote sensing satellites and astronomical observation satellites when performing multi-field-of-view star imaging.

The present invention provides a method, device and electronic device for determining the inter-satellite distance of satellites, which obtains the first orbital inclination, the first latitude argument and the first ascending node right ascension of the first satellite, and obtains the second orbital inclination, the second latitude argument and the second ascending node right ascension of the second satellite; wherein the first satellite and the second satellite belong to the same satellite shell; based on the obtained information and the pre-constructed target distance expression, the maximum distance threshold and the minimum distance threshold between the first satellite and the second satellite are determined. In this method, it is only necessary to obtain the orbital inclination, latitude argument and ascending node right ascension corresponding to the first satellite and the second satellite, and the maximum distance threshold and the minimum distance threshold corresponding to the first satellite and the second satellite can be calculated by the target distance expression, and the calculation is simpler and more efficient.

The invention provides an ultra-low orbit communication constellation design method based on a fixed sub-satellite point trajectory constellation, which comprises the following steps: (1) task demand description, namely designing an ultra-low orbit constellation, so that a target location T can realize continuous communication with duration of T every day; if the view field half cone angle of the satellite sensor is that the target location is in the sensor coverage range of any satellite, the communication is successful; (2) deriving a sub-satellite point regression condition; (3) acquiring the orbit semi-major axis and the inclination angle of the constellation, and further determining the satellite orbit height; (4) determining the constellation scale to ensure that the target location is within the sensor duration range of the satellite at any moment; and (5) determining the constellation phase, namely determining the right ascension and the latitude argument of the ascending node of each satellite. The method is reasonable in conception, can ensure that a target site can realize continuous uninterrupted communication with fixed duration every day, adopts a sub-satellite point sharing design to facilitate installation of ground equipment, and can flexibly adjust constellation configuration.

The application relates to a steady-state gain correction GNSS autonomous navigation system based on normalized orbit parameters, which comprises a GNSS information conversion module, an orbit extrapolation updating module, a steady-state gain Kalman filter correction module and a steady-state gain offline calculation module. The application adopts a normalized orbit semi-major axis, an orbit inclination, an ascending node right ascension, an eccentricity vector and a flat latitude amplitude angle as state variables of navigation filter estimation, adopts a Cowell method to perform one-step prediction of the above-mentioned orbit element information, converts position in a WGS84 system output by a GNSS receiver into orbit element information in combination with one-step predicted speed information, and adopts a steady-state gain Kalman filter algorithm to complete filter correction of the above-mentioned orbit element information.

The invention provides an orbit design method for space-based observation of a low-orbit satellite constellation, belongs to the field of spacecraft orbit design, and solves the problem that a space-based observation platform is low in observation efficiency of the low-orbit satellite constellation. Comprising the following steps: combining an earth gravitational potential function with an earth non-spherical perturbative force function to obtain a J2 perturbative force function; a Kepler orbital element is used for representing a J2 perturbation function, and the J2 perturbation function is combined with a Lagrange type perturbation motion equation to construct a Lagrange perturbation model; a lagrangian perturbation model, the influence of J2 perturbation force on the number of low-orbit satellite orbits and the drift law of the right ascension of an ascending node are utilized to construct complementary angles of the orbit inclination angle of a space-based observation platform and the orbit inclination angle of a low-orbit satellite; the space-based observation platform is a retrograde orbit configuration or a plurality of retrograde orbit configurations, and the eccentricity rate and the semi-major axis of the space-based observation platform are the same as those of the low-orbit satellites, so that observation of all the low-orbit satellites in the constellation is completed; according to the invention, the observation efficiency of the space-based observation platform on the low-orbit satellite constellation is improved.

The current disclosure provides systems and methods of directly measuring angular velocity and angular acceleration of space objects and using the measured angular velocity and angular acceleration as inputs into new and unique algorithms for initial orbit determination. Sensors measuring locations and times of light events may be used to generate a virtual rate track image for identification of space objects. Right ascension and declination of the space object events versus time may be fit to polynomials or splines to determine associated angle, angular rate, and angular acceleration of the space objects. New and unique initial orbit algorithms may then be applied to estimate orbital elements of the space objects.

A high-precision lunar angle prediction method for a space remote sensing instrument includes the following steps: according to the forecast start time broadcast by a satellite platform and the required forecast time interval, the lunar geocentric apparent right ascension and apparent declination at the current forecast time are calculated; according to the satellite orbit instant root and perturbation motion equation at the forecast start time broadcast by the satellite platform, the satellite orbit instant root at the current forecast time is calculated; according to the satellite orbit instant root at the current forecast time, the transformation matrix from the geocentric inertial coordinate system to the satellite orbit coordinate system is calculated; according to the satellite attitude angle at the forecast start time broadcast by the satellite platform, the satellite attitude correction matrix at the current forecast time is calculated; according to the above results and the calibrated satellite body coordinate system to the instrument body coordinate system conversion matrix, the lunar angle in the instrument body coordinate system at the current forecast time is calculated. The present invention simplifies the forecast process of the lunar angle, is simple to calculate, and can perform long-term and high-precision lunar angle forecasts.

The invention discloses a space target observation right ascension and declination data three-dimensional simulation method and system, and the method comprises the steps: adding observation equipment and an observed target, and marking a J2000 position; calculating the J2000 position of the observation equipment and the observed target in the visible time period; converting the J2000 position of the observed target in the visible time period into a required J2000 position; the converted J2000 position data is right ascension and declination; and setting the same height according to the calculated right ascension and declination, and displaying in a three-dimensional situation. Through the three-dimensional visualization technology, the defects of two-dimensional projection are effectively avoided, the motion trail and the space relation of the space observed target can be accurately presented, and reliable support is provided for related research, task planning and the like.

The present application provides a star catalog data merging method, apparatus, computer equipment, and storage medium. The method comprises: obtaining a target star catalog file; determining the numbering information of each celestial data item at a target level based on the right ascension and declination included in the celestial data; dividing the celestial data items with the same numbering information in the target star catalog file into the same data group, thereby obtaining multiple data groups; for each data group, determining the adjacent data group corresponding to the data group based on the numbering information corresponding to the data group; and for each celestial data item in the data group, determining the matching results between the celestial data item and other celestial data items in the data group, as well as the matching results between the celestial data item and each celestial data item in the adjacent data group; and generating an index file corresponding to the target star catalog file based on the matching results between the multiple celestial data items. The present application implements cross-merging processing of the entire star catalog file, reduces the number of repeated cross-merging, and improves the efficiency of cross-merging of star catalog data.

The application discloses a spacecraft trajectory planning method for constellation multi-target traversal flyby, and belongs to the technical field of aerospace. The method comprises the following steps: introducing a flyby time constraint condition and an ideal flyby azimuth, constructing a task star flyby state optimization problem, and generating a relative motion trajectory for flying by a single target. A tangential velocity pulse is applied to an apogee of the task star to adjust a semi-major axis, and a difference in right ascension drift rate of the task star and target stars under J2 perturbation at different orbit semi-major axes is utilized to realize transfer of the task star among orbit planes of the target constellation. A correction maneuver is applied during orbit plane drift of the task star to realize matching of an amplitude angle of the task star apogee and a target star phase, so that the task star enters a resonance orbit for satellite traversal flyby in the constellation. The above-mentioned transfer trajectory optimization process of the task star among adjacent orbit planes is repeated to realize flyby traversal of a single task star to all satellites of the target constellation, and further realize spacecraft trajectory planning.

The application discloses a GNSS-based extrapolation method for a large-ellipse frozen orbit, which comprises the following steps: step one: acquiring the position and speed of a navigation receiver orbit according to the effective state of navigation data; step two: calculating satellite orbit elements (distance and speed values, moment of momentum, semi-major axis, orbital angular velocity, eccentricity, orbital inclination, ascending node right ascension, latitude amplitude angle, and perigee amplitude angle) by using t GNSS moment of momentum, semi-major axis, orbital angular velocity, eccentricity, orbital inclination, ascending node right ascension, latitude amplitude angle, and perigee amplitude angle); step three: judging the effectiveness of navigation data according to the calculation results of the satellite orbit elements; step four: extrapolating and calculating the orbit parameter true anomaly and latitude amplitude angle at the current satellite time; and step five: calculating the instantaneous orbital angular velocity, thereby completing the GNSS-based orbit extrapolation design for a large-ellipse frozen orbit. The application only uses the orbiting information of the GNSS time to extrapolate the orbit elements of the satellite, thereby achieving satellite autonomous positioning and autonomous attitude control, reducing the positioning requirements of the satellite to the ground station, and improving the autonomous ability of the satellite.

The integrated equatorial telescope comprises the integrated equatorial telescope with a multi-rotating-shaft structure, the right ascension angle is adjusted through a first rotating shaft, the declination angle is adjusted through a second rotating shaft, the azimuth angle is adjusted through a third rotating shaft, the pitch angle is adjusted through a fourth rotating shaft, and system balance is adjusted in combination with a balance part. Meanwhile, the invention provides a full-automatic pole alignment method based on the telescope. The north pole or the south pole can be accurately positioned, and synchronous correction of coordinates is achieved. According to the invention, the effects of improving the operation convenience of the telescope, improving the antipode precision and stability and optimizing the balance performance are achieved.

The invention relates to a right ascension axis visual polar axis calibration method based on laser beam indication and an equatorial telescope, and the method comprises the following steps: adjusting a laser beam through a laser beam parallelism adjusting mechanism, and enabling the laser beam to be parallel to the right ascension axis of the equatorial telescope; the direction of an optical imaging unit is adjusted, the optical imaging unit comprises an optical lens and an imaging camera used in cooperation with the optical lens, and the tail end of the laser beam is made to be located in the view field of the optical imaging unit; a latitude adjusting seat of the equatorial telescope is adjusted, so that the laser beam points to the vicinity of the polar axis; performing exposure by using an optical imaging unit to obtain recognizable star points; the polar axis position is positioned based on the relative position of the recognizable star points; and the latitude adjusting seat is adjusted, so that the imaging point of the tail end of the laser beam on the display screen coincides with the polar axis position. According to the invention, laser beam indication is combined with an optical imaging technology, accurate polar axis alignment of the equatorial telescope before use is realized, and the equatorial telescope is suitable for astronomical observation scenes of southern and northern hemispheres.

The invention provides a heterogeneous orbit constellation configuration regression period determination method and system, and the method comprises the steps: finding a group of coprime operation turns and regression number for each satellite, and enabling the operation turns and the regression number to meet the condition that the ratio of the operation turns to a first set parameter is equal to the ratio of the regression number to a second set parameter, the regression period of the satellite can be obtained; wherein the first set parameter is the sum of the average change rate of the perigee argument and the average change rate of the mean anomaly; the second set parameter is the difference between the earth rotation angular velocity and the right ascension precession rate of the ascending node; and calculating the least common multiple of the time periods of all satellites to obtain the regression period of the constellation. The method has the advantages that compared with a traditional method, the efficiency is improved by 13.3% when compatibility analysis is carried out on the regression cycle calculated through the method, and the maximum interference level in the regression cycle is consistent with that of annual simulation.

The present application relates to the technical field of spacecraft attitude orbit control, and particularly relates to a kind of coplanar installation electric thruster autonomous position keeping control method and device.Method includes: when receiving autonomous position keeping control instruction, based on current star time and pre-determined position keeping control period, the orbit inclination vector variation in position keeping control period is determined;Based on the orbit inclination vector variation in position keeping control period, the total firing duration in position keeping control period is determined to determine the firing duration corresponding to each operating day;Based on the orbit inclination vector variation in position keeping control period, the firing point right ascension is determined;Based on the firing duration corresponding to each operating day and the firing point right ascension, the coplanar installation electric thruster is used to autonomously carry out position keeping control in each operating day.This scheme can predict the position keeping parameters of multiple days, and the calculation is simple and easy to implement.

The present application relates to a method for designing a transfer orbit from a DRO orbit to a low lunar circular orbit. The method includes: establishing a mapping data table and a three-dimensional data pair of the delta-v of orbit transfer according to the preset DRO orbit size parameters and the phase angle of the DRO vehicle at the initial moment; dividing the three-dimensional data pair of the delta-v of orbit transfer into two groups according to the positive and negative conditions of the high and low angles of the delta-v of orbit transfer, obtaining the right ascension of the ascending node and the transfer flight time corresponding to the data pair with the minimum delta-v during the transfer process in each group to calculate the orbit transfer moment, and selecting the minimum value of the two orbit transfer moments as the selected orbit transfer moment; correcting the phase of the selected orbit transfer moment, and calculating the projection of the delta-v of orbit transfer according to the corrected phase; using the Newton iteration method to perform iterative solution with the selected orbit transfer moment and the projection of the delta-v of orbit transfer as the initial iteration values to obtain the transfer orbit transfer moment and the transfer delta-v of orbit transfer. Using this method can improve the design accuracy of the transfer orbit.