82 results about "Orbital elements" patented technology

The invention relates to a laser communication system autonomous pointing method based on an inter-satellite link. The method comprises the steps that a laser communication system receives link building target satellite instantaneous orbit parameters transmitted by the inter-satellite link; the laser communication system converts the link establishment target satellite instantaneous orbit parameter transmitted by the inter-satellite link into a singular-point-free average orbit element, and the singular-point-free average orbit element serves as a link establishment target satellite orbit initial value of an inter-satellite link source; the laser communication system selects a link establishment target orbit initial value of an inter-satellite link source according to the instruction or the state, and calculates real-time orbit data of a link establishment target satellite; and the laser communication system automatically calculates a laser beam pointing angle according to the attitude and orbit data of the satellite and the real-time orbit data of the link building target satellite. According to the invention, the dependence of maintaining the inter-satellite laser link on the measurement and control conditions of the ground station can be reduced, and the laser beam pointing precision can be effectively improved.

The invention relates to the technical field of low-orbit satellite communication, in particular to an orbit optimization method and system for dynamically and autonomously tracking a low-orbit satellite based on a ground station, and the method comprises the steps: collecting TLE orbit element data, constructing an SGP4 orbit prediction model, and obtaining a prediction result; low-earth-orbit satellite orbits with different inclination angles are determined through STK software modeling based on a ground station, and corresponding standard results are obtained; comparing the prediction result with a standard result to obtain an error result; dynamically optimizing the Bstar parameter based on an error result by adopting a genetic algorithm, and determining an optimal Bstar parameter; adjusting the SGP4 orbit prediction model based on the optimal Bstar parameter, and obtaining an optimization result through the adjusted SGP4 orbit prediction model; according to the method, optimization is carried out based on an existing SGP4 model, optimal Bstar parameters are dynamically searched through a genetic algorithm, accumulative errors caused by the fact that TLE data are not updated in time are compensated, stable tracking of a ground station is guaranteed, and the precision of medium and long term orbit forecasting is improved.

The invention provides a two-line element generation method and orbit prediction method adaptive to an analytic propagation model, and the method comprises the steps: obtaining the historical osculating orbit elements at each epoch moment, and carrying out the orbit element optimization step of a plurality of iteration rounds on each historical osculating orbit element through employing the analytic propagation model, obtaining a target osculating orbital element number at each epoch moment, enabling the historical osculating orbital element number to be the same as the target osculating orbital element number, and performing inversion to obtain an average orbital element number; and extracting each average semi-major axis of each average orbital element, fitting each average semi-major axis by adopting a least square method to obtain a semi-major axis attenuation rate, inverting the semi-major axis attenuation rate to obtain an atmospheric resistance parameter, and equivalently converting the atmospheric resistance parameter into a standard resistance term in two rows of elements to obtain the two rows of elements. According to the method, the problems of inconsistent models, insensitive residual errors, unable physical explanation of parameters and the like in the two-line element generation process can be solved, and the ephemeris extrapolation orbit prediction precision is improved.

The invention provides a centimeter-level broadcast ephemeris method and device based on a GNSS precise ephemeris, and the method comprises the steps: converting a classical Kepler orbit element into a nonsingular orbit element for the near-zero eccentricity rate of a middle earth orbit satellite; performing fast Fourier transform on a satellite orbit, selecting a certain number of additional parameters from the optional additional parameter set according to a spectral analysis result of the Fourier transform, and combining the additional parameters with an original broadcast ephemeris model to construct a centimeter-level broadcast ephemeris model; interpolating the precise ephemeris to obtain dense ephemeris data points, and determining centimeter-level broadcast ephemeris model parameters; calculating an error in a geocentric inertial coordinate system, converting the error into an orbital coordinate system, and determining the precision of a centimeter-level broadcast ephemeris model by using a space signal ranging error; and broadcasting the centimeter-level broadcast ephemeris model for satellite orbit determination and user high-precision positioning. The method provides support for real-time high-precision satellite positioning application.

The invention belongs to the technical field of aerospace asteroid defense, and relates to a variable mass gravitational traction deflection asteroid orbit full-flow design method based on real working condition constraint, which comprises the following steps of: obtaining orbit elements corresponding to a target asteroid according to initial conditions; substituting into a high-precision numerical integration model, and calculating perigee moments and perigee distances under the condition that the gravitational traction action is not applied; according to the early warning time requirement, the carrying capacity and the real engineering constraint, a genetic algorithm and sequential quadratic programming are used, the maximum perigee deflection distance serves as a target, and a transmitting window is optimized and solved; and substituting the optimized emission window into the high-precision deflection efficiency numerical integration model, and outputting the asteroid perigee deflection distance. According to the method, comprehensive analysis is carried out in combination with early warning time, emission capability and actual working conditions from the perspective of a whole-process task, so that the authenticity and engineering applicability of a calculation result are ensured.

The invention discloses a spatial distributed observation method based on irregular region geometric feature sampling, and belongs to the technical field of aerospace. The implementation method comprises the following steps: establishing a coordinate system for describing the motion of the spacecraft; establishing a spacecraft dynamic model under the coordinate system; six initial orbit elements of a satellite can be obtained based on existing low-orbit constellation orbit data, the spatial position is calculated through the orbit elements, and then position components are converted into latitude and longitude coordinates, namely sub-satellite point coordinates; defining vertex coordinates of the polygon, and constructing a planar polygon by connecting the vertex coordinates end to end; discretizing the internal region of the closed polygon; on the basis of the kinetic model, on the basis of the coordinates of the sub-satellite point and the center coordinates of the discrete sampling circle, calculating the characteristic distance from the sampling point in the observation area to the sub-satellite point, and on the basis of the size relationship between the characteristic distance and the radius of the observation view field, judging whether the irregular area can be successfully observed in a preset time constraint interval; therefore, spatial distributed observation is realized.

Embodiments of the present application provide a space target orbit prediction method, a prediction device, an electronic device and a medium. The method comprises: obtaining two-line orbital elements of a space target at a current time, space environment data and perturbation coefficients, the perturbation coefficients being coefficients in a space target orbit prediction model; correcting the two-line orbital elements and / or the perturbation coefficients based on a pre-trained parameter correction model, the two-line orbital elements, the perturbation coefficients and the space environment data, to obtain corrected data, the corrected data comprising corrected two-line orbital elements and / or corrected perturbation coefficients; and predicting orbit information of the space target at a preset future time based on the space target orbit prediction model and the corrected data. The method makes the predicted orbit information more accurate.

The invention discloses a ground-based radar space object long-time energy accumulation method based on an orbital element model, and the method comprises the following implementation steps: 1, building an orbital element parameter space, and mapping the grid points of the parameter space into the distance between a target and a radar; and step 2, calculating an envelope position and a phase which need to be compensated according to the mapped distance history at each parameter space grid point. And step 3, compensating a corresponding phase for the echo signal after pulse compression along the calculated envelope position, and then performing coherent superposition, thereby realizing long-time coherent accumulation of echo energy. According to the method, the orbital elements are adopted to construct the parameter space, energy accumulation is carried out in the parameter domain, due to the fact that the orbital elements can better represent the actual motion of a space object, the energy accumulation effect can be improved, meanwhile, the orbital elements are small in parameter space dimension, and the algorithm calculation amount can be greatly reduced.

The application provides a method for analyzing a pointing angle of an inter-satellite link, comprising: obtaining orbital elements of a satellite and orbital parameters of a target satellite matched with an SGP4 model in a J2000.0 inertial coordinate system; determining a target recursion time t; obtaining positions of the satellite and the target satellite at the time t in the J2000.0 inertial coordinate system; determining a pointing vector from the satellite to the target satellite at the time t in the J2000.0 inertial coordinate system; determining the pointing vector in a satellite orbital coordinate system; determining the pointing vector in a satellite body coordinate system; determining the pointing vector in a satellite antenna coordinate system; determining a pointing angle of the satellite antenna coordinate system at the target recursion time; and determining whether a cycle number reaches a set cycle number, if the cycle number does not reach the set cycle number, continuously obtaining the pointing angle of the satellite antenna coordinate system at a next target recursion time until the pointing angle of the satellite antenna coordinate system at the set cycle number is obtained.

The invention provides an earth-moon transfer orbit iterative calculation method based on earth-moon relative motion, which comprises the following steps: firstly, calculating an ideal earth-moon transfer orbit number according to the deviation between the current flight time and the theoretical time, and taking the ideal earth-moon transfer orbit number as an iterative guidance target orbit; and then, by utilizing the characteristic that the shutdown time of the rocket in the flight section can be predicted in the iterative guidance calculation process, the influence of the flight time deviation on the earth-moon transfer orbit is compensated to the number of target orbits in real time, so that correction of a guidance program angle instruction is realized, and the rocket is controlled to fly towards a more ideal earth-moon transfer orbit all the time.

The invention provides a non-singular semi-analytical method for forecasting a low-thrust perturbation orbit. The method comprises the following steps: firstly, establishing a Gaussian variational equation for describing spring equinox orbital element evolution under small thrust control based on a periodic small thrust hypothesis; thirdly, constructing a long-term equation and a short-period term based on the average spring equinox orbital elements, performing large-step integration on the long-term equation to obtain a flat root, and then performing orbital element correction through the short-period term to finish high-precision orbital forecasting; and finally, establishing an analytic form error estimation model through a semi-major-axis second-order long-term equation, and using the analytic form error estimation model to evaluate orbit forecast errors. The semi-analytic low-thrust orbit forecasting method provided by the invention achieves good balance between calculation efficiency and precision, avoids the problem of singularity, and is suitable for high-precision forecasting of any low-thrust elliptical orbit; the error estimation model provided by the invention can effectively quantify the prediction error and enhance the reliability of the orbit prediction results of most small eccentricity rate on-orbit satellites.

The invention relates to a multi-satellite collaborative observation probability estimation method for a window type target. The method comprises the following steps: inputting a statistical time range, an orbital element number and a load observation range of a guiding constellation A, and an orbital element number and a load observation range of a guided constellation B; and determining a to-be-observed area O, based on whether the constellation A is a geostationary orbit satellite, calculating the collaborative observation probability that the constellation A guides the low-orbit constellation B, and outputting a collaborative observation probability result. According to the method, the defect that the time dynamic characteristics of the window type target are not fully considered when the collaborative observation probability of the satellite to the land target is analyzed in the existing method is overcome, and the problem that the window type target collaborative observation success probability is inaccurately estimated in the existing method is solved.

The present invention belongs to the field of aerospace asteroid defense technology and relates to a full-process design method for variable-mass gravitational traction deflection of an asteroid orbit based on real-world operating constraints. The method comprises the following steps: obtaining the orbital elements corresponding to the target asteroid based on initial conditions; substituting them into a high-precision numerical integration model to calculate the perigee moment and perigee distance without gravitational traction; optimizing the launch window using a genetic algorithm and sequential quadratic programming, with the goal of maximizing the perigee deflection distance, based on warning time requirements, launch capacity, and real-world engineering constraints; and substituting the optimized launch window into a high-precision deflection efficiency numerical integration model to output the asteroid's perigee deflection distance. This method, based on the full-process mission perspective, conducts a comprehensive analysis combining warning time, launch capacity, and actual operating conditions to ensure the authenticity and engineering applicability of the calculation results.

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.

This application relates to a radar-optical hybrid dual-arc correlation orbit determination method, comprising: acquiring radar tracking arc segments and optical tracking arc segments; calculating the initial radar orbital elements based on the radar tracking arc segments using a Lambert boundary value problem; calculating the initial optical orbital elements based on the optical tracking arc segments using a distance search method to obtain candidate solutions; determining whether the radar tracking arc segments and optical tracking arc segments could originate from the same space target; if not, the process terminates; if possible, it continues; using the radar initial orbital elements to determine the estimated semi-major axis, the candidate solutions are screened, and a position observation sequence arc segment is generated; based on the position observation sequence arc segment and the radar tracking arc segment, a Lambert boundary value problem is constructed to perform dual-arc correlation orbit determination, and the solution is obtained using the Gooding method. This invention provides a radar-optical hybrid dual-arc correlation orbit determination method that constructs a Lambert boundary value problem to solve dual-arc orbit determination, achieving excellent radar-optical hybrid dual-arc correlation accuracy and dual-arc orbit determination precision.

The invention relates to an arbitrary polyhedron spinning space debris space-based observation light variation curve simulation method. The method comprises the following steps: calculating position vectors of debris and an observation satellite in a geocentric inertial system at the current moment according to orbital elements of the debris and the observation satellite; calculating a sun vector and an observation satellite vector under the debris body system according to the attitude of the debris at the current moment; calculating a sun incident zenith angle and a satellite observation zenith angle of each surface of the debris; calculating the actual effective reflection area of each surface of the fragment; calculating the optical reflectivity of each surface of the fragments based on a BRDF scattering model; calculating the total reflection irradiance of each surface of the fragments; calculating to obtain the apparent magnitude of the fragments at the observation satellite according to an apparent magnitude calculation formula; and on the basis of a preset fragment attitude sequence, repeating the steps to calculate and obtain the light variation curve of the fragments. According to the method, the attitude spinning condition of the debris target is fully considered, and a basis can be provided for designing a space-based debris observation satellite remote sensor and inverting the debris attitude by using the light variation curve.

The invention provides a satellite chip thermal management optimization method and device, and the method comprises the steps: carrying out the thermal load time sequence prediction of orbit elements and solar calendar data, obtaining the thermal load characteristics of an orbit, and calculating a heat dissipation control parameter according to a characteristic coefficient; according to the parameter matrix, fatigue accumulation characteristics of the heat dissipation part are analyzed, a health state evaluation result is obtained, a working mode is switched, and a time division multiplexing control strategy is obtained; performing matching optimization on chip thermal resistance and heat dissipation thermal resistance according to a control strategy, and coupling chip power consumption and a heat dissipation structure through a predictive control mechanism to obtain a heat balance control parameter; and analyzing thermal management performance according to the control parameters to obtain performance change trend data, predicting degradation of the heat dissipation structure, and obtaining a degradation prediction result and a compensation control instruction. According to the method, predictive thermal load analysis based on the track position and heat dissipation structure self-adaptive regulation and control are established, so that progressive failure of a thermal management system caused by asymmetric thermal circulation of the elliptical track is avoided.

The embodiment of the invention provides a space target maneuvering detection method and system based on an information energy momentum tensor, and relates to the technical field of spaceflight detection. Comprising the steps of obtaining orbital elements of an observation satellite and a space target; according to the number of the orbits, the number of the wide-area invariant opposite orbits is obtained; obtaining a space target relative position after dimension normalization under an observation satellite LVLH coordinate system according to the wide-area invariant relative orbital element; according to the relative position and the number of the relative orbits, the relative motion acting amount and the relative speed are obtained; carrying out filtering processing on the relative motion state of the observation satellite according to a mean filtering algorithm; acquiring a corresponding energy momentum tensor according to the filtered relative motion action quantity, relative position and relative speed; and obtaining a time-varying curve of the energy momentum tensor, and carrying out maneuvering detection on the space target. According to the method, the detection capability on low thrust and weak signal maneuver can be remarkably enhanced, and the limitation that a traditional method is high in dependence on model precision and insensitive to micro maneuver is overcome.

The present invention discloses a method, device, medium, and product for calculating patterned achievable orbital elements and making decisions about degraded orbits, relating to the field of rocket achievable orbital element calculation. The method comprises obtaining the rocket's initial flight state parameters, target orbital elements, and safety altitude constraints under non-lethal thrust descent, establishing a patterned optimal control problem model based on the rocket's dynamics differential equations and operational constraints, and convexifying it to obtain a convex optimal control subproblem. The convex optimal control subproblem is discretized and slack variables are introduced to obtain a convex optimization subproblem. The method further comprises setting the rocket's mode coefficients in different modes, iteratively solving the convex optimization subproblem using a convex optimization algorithm, obtaining achievable orbital elements for different modes, making decisions on rocket orbit degradation, obtaining the final rocket achievable orbital elements, and determining a reference trajectory for the rocket's operation. The present invention can improve the calculation efficiency of achievable orbital elements and enhance mission adaptability.

The present invention discloses a method and system for optimizing the orbital parameters of a double star, including: Step 1) According to the six orbital elements of the primary star, calculate the time interval to obtain the <time, position, velocity> information matrix of the primary star and the theoretical solution set of the <time, position, velocity> information matrix of the secondary star; Step 2) Set the constraint conditions, establish an optimization model for the orbit of the secondary star, and according to the set solution granularity, use the dynamic mutation genetic algorithm. By setting the mutation operation probability that changes dynamically with the iteration step size, optimize the six orbital elements of the secondary star to obtain the actual solution of the <time, position, velocity> information matrix of the secondary star, and use the number of occurrences of the actual solution in the theoretical solution set as the fitness function; Step 3) Calculate the fitness function to obtain the continuous duration of maintaining communication between the two stars of the double star, and judge whether the continuous duration remains unchanged. If it is judged to be yes, go to Step 4), otherwise, go to Step 2); Step 4) Output the positions of the primary star and the secondary star that meet the conditions.

The invention discloses a fuel-saving control on-satellite autonomous planning method for rapidly transferring an elliptic orbit different-plane target to the same orbit, and particularly relates to a fuel-saving control on-satellite autonomous solving method for rapidly transferring the elliptic orbit different-plane target to the same orbit from a different orbit plane. Comprising the following steps: judging orbital plane normal correction or orbital shape correction according to orbital elements of an initial satellite and a target; solving a fuel-saving orbital plane normal correction control strategy according to orbital plane normal orbital element difference of the satellite and the target; solving a fuel-saving orbit shape correction control strategy according to the in-orbit element difference of the orbit planes of the satellite and the target; and a high-precision orbit extrapolation model is adopted to carry out iterative correction on the fuel-saving control strategy. The method is suitable for orbital transfer of any elliptical orbit target by any near-earth space elliptical orbit satellite, and the control precision can reach ten meters; and the calculation amount is small, and the method is suitable for on-satellite autonomous task planning and solving.

A space information recorder (100) acquires space object information (500), which is orbit forecast information of a plurality of space objects (60), from a management business device (40) used by a management business operator that manages the plurality of space objects (60), and records the space object information (500). The space object information (500) includes a forecast epoch, a forecast orbital element, and a forecast error of each of the plurality of space objects (60). When it is foreseen that a space object A included in the plurality of space objects (60) will intrude into a range at orbital altitudes of 300 km to 1000 km in which a satellite group of LST 10:00 to 11:00 is present, a time period from intrusion to exit and orbit forecast information are recorded.

An ephemeris server executes a method including the following: (a) (i) obtaining broadcast ephemeris information by accessing a broadcast information source on a wide area communication network; (ii) extracting orbit trajectory information of each satellite from the broadcast ephemeris information; and (iii) creating modified orbit trajectory information of each satellite from the extracted orbit trajectory information and storing it in a real-time database; (b) at each of a plurality of specified time points, updating the modified trajectory information in the real-time database based on an evaluation of an update criterion; and (c) providing the modified trajectory information to each of a plurality of devices, each device having a GNSS receiver capable of processing a broadcast signal of a satellite using the modified trajectory information.

The application provides a single-satellite or multi-satellite re-track interference orbit control method based on an effective vertical baseline, and belongs to the technical field of space-borne SAR re-track interference. The single-satellite or multi-satellite re-track interference orbit control method is based on an eccentricity vector and a relative inclination, a reference-observation virtual formation model of re-track interference is established; based on the geometric relationship between the target detection area latitude and the subsatellite point latitude and the orbit root number, the relationship between the target detection area latitude and the effective vertical baseline is obtained, and the virtual formation effective vertical baseline of the reference-observation satellite is determined; and the in-plane / out-of-plane orbit control strategy suitable for single-satellite or multi-satellite re-track interference is obtained by taking the effective vertical baseline as an evaluation index. The application solves the problem that the orbit control strategy research of single-satellite re-track interference is relatively insufficient at home and abroad, can provide a new method and a new direction for establishing a remote sensing system of China, and can be applied to space-based situation awareness to support the improvement of the space-based attack and defense system of China.

The invention provides a differential interference SAR constellation three-dimensional surface deformation field measurement precision analysis method and system, and the method comprises the steps: inputting the orbital elements and load observation capabilities of all satellites in a constellation, and setting a to-be-analyzed observation region; simulating and calculating the observation time of each satellite in the observation area to be analyzed in a strict regression period; calculating observation sight vectors of the satellites to the simulation grid points in the observation area; analyzing and estimating the sight deformation measurement precision of each satellite to the observation area; and establishing an observation equation, and analyzing and estimating the three-dimensional deformation measurement precision of the whole constellation on the observation area. According to the method, the problem of differential interference SAR constellation three-dimensional surface deformation monitoring precision analysis is successfully solved, and important technical support can be provided for construction of a subsequent differential interference SAR surface deformation monitoring system.

The application relates to a satellite orbit and satellite attitude determination method and device after satellite-rocket separation. The method comprises the following steps: calculating a first position vector and a first velocity vector of a satellite in an epoch equatorial coordinate system according to orbit elements of a rocket in a second earth-centered equatorial coordinate system and a satellite-rocket separation time; determining a second position vector and a second velocity vector of the satellite in a J2000 earth-centered inertial coordinate system based on the first position vector and the first velocity vector; taking the second position vector and the second velocity vector as orbit initial values, and determining orbit parameters of the satellite before injection into an orbit by using an orbit recursion algorithm; determining an attitude direction cosine matrix of the satellite in the J2000 earth-centered inertial coordinate system at the satellite-rocket separation time based on a launch time; and taking the attitude direction cosine matrix as attitude initial values, and determining attitude parameters of the satellite before injection into the orbit by using a star gyroscope combination filtering method. The scheme can quickly determine the orbit and the attitude of the satellite, and improves the autonomous detection capability of the satellite.

The invention provides a solar system planet orbit parameter inversion method and system, and belongs to the technical field of astronomical data processing and celestial body orbit calculation, and the method comprises the steps: obtaining the Chebyshev polynomial order and coefficient of a planet; carrying out time normalization processing on a planet position observation moment, and carrying out planet position accurate calculation by adopting Chebyshev polynomial interpolation; estimating the number of initial Kepler orbits; establishing a mathematical model between the planetary position and the Kepler orbital element, constructing an objective function, and performing Jacobian matrix calculation and orbital element solution iterative optimization processing; and performing quadratic polynomial fitting on the multiple groups of Kepler orbit elements to generate Kepler orbit elements in a quadratic polynomial coefficient form. The method is used for meeting the requirement for low calculation amount of medium-precision solar system planetary orbit parameters and planetary positions in the fields of astronomical research, astronomical observation, spaceflight orbit design, space exploration task planning and the like.

The application discloses a low-fuel-consumption low-orbit satellite orbit recovery method, and belongs to the technical field of satellite orbit recovery.The application aims at the problem that the existing orbit maneuver method needs frequent orbit adjustment, which is not conducive to the long-term operation of the satellite in orbit with low fuel consumption.The application comprises the following steps: establishing an orbit dynamics equation considering the earth oblateness perturbation; using a semi-analytical method to solve the initial orbit elements of a low-fuel-consumption transition orbit according to the current orbit and the target orbit of the satellite; simultaneously, combining the orbit dynamics equation, calculating the drift time required for the satellite to drift on the transition orbit by using the earth oblateness perturbation without fuel consumption and the final orbit elements of the transition orbit at the end of the transition; and based on the Pontryagin maximum principle, calculating the optimal control rate sequence of the thruster fuel of the satellite from the current orbit to the transition orbit and from the transition orbit to the target orbit, and controlling the thruster to realize the orbit recovery of the satellite.The application is used for satellite orbit recovery.

The application discloses a scanning method and device of a non-cooperative target, electronic equipment and program product, and relates to the technical field of spaceflight measurement and control, wherein the scanning method comprises the following steps: acquiring initial orbit elements of the non-cooperative target, celestial body parameters of a target celestial body and antenna parameters of an antenna; determining the number of beam pointing positions required for the antenna to scan one orbit period of the non-cooperative target based on the initial orbit elements, the celestial body parameters and the antenna parameters; determining an acceleration multiple based on the orbit period of the non-cooperative target, the number of beam pointing positions and the antenna parameters; and generating guide data based on the initial orbit elements and the acceleration multiple, so as to control the antenna to scan the non-cooperative target through the guide data. The application solves the technical problem that the non-cooperative target cannot be efficiently scanned in the related art, thereby reducing the probability of capturing the non-cooperative target.

The invention provides a method and a device for checking correctness of data injected into a relay terminal of a spacecraft, and relates to the technical field of spacecraft measurement and control. The method comprises the following steps: acquiring two adjacent injection data which are injected into a relay terminal and aim at a target relay satellite; the two adjacent times of injection data comprise two groups of relay satellite orbit information; determining an actual phase difference between the two groups of orbit elements according to the two groups of relay satellite orbit information, and comparing the actual phase difference with a predetermined theoretical phase difference; and determining a relay satellite orbit information inspection result according to a comparison result. The device executes the method. According to the method and the device for checking the correctness of the data injected by the relay terminal of the spacecraft provided by the embodiment of the invention, the correctness of the data injected by the relay terminal can be checked efficiently, so that accurate measurement and control of the spacecraft are realized.