50 results about "Orbital elements" patented technology

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 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 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 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 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 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 discloses a calculation method and device for a satellite VDES load coverage area, a storage medium and calculation equipment, and the method comprises the steps: determining the position and attitude of a satellite according to the orbital element number of the satellite, a designated moment and a satellite working mode; determining the direction of a VDES antenna based on the position and attitude of the satellite and the load installation parameters; based on the VDES antenna pointing, constructing a polar plane of a satellite about the surface of an earth ellipsoid, and determining a parameter equation of a space ellipse where the polar plane intersects with the surface of the earth ellipsoid; according to the parameter equation, traversing angle parameters in the space ellipse to obtain a tangent point track and an intersection point track; and calculating a load coverage area of the VDES antenna based on the tangent point track and the intersection point track. The technical problem that the VDES load coverage area is difficult to calculate in the prior art is solved.

The invention discloses an autonomous control method for a sun synchronous regression orbit, and belongs to the technical field of spacecraft orbit design and autonomous control, and the method comprises the steps: executing an in-plane and out-plane instruction calculation step based on a current instantaneous orbit element vector when determining that a satellite passes through a descending node based on a real-time position vector and a real-time adjacent moment position vector; obtaining in-plane maneuvering final execution time and an out-plane maneuvering instruction; and executing in-plane maneuver based on the in-plane maneuver final execution time, and executing out-plane maneuver when the satellite passes through the ascending node based on the out-plane maneuver instruction. The invention provides an autonomous control method for a sun-synchronous regression orbit, and solves the technical problem that an existing high-precision autonomous orbit control method for the sun-synchronous strict regression orbit is not suitable for the sun-synchronous strict regression orbit at a non-morning-night non-midnight intersection point. And the long-term stability and robustness of the control are enhanced while the high precision of the autonomous control of the sun-synchronous orbit is ensured.

The invention discloses a satellite orbit thrust section inversion method based on pattern search, solves the problem of how to accurately invert parameters of a middle unknown thrust section by using two extremely short radar observation arc sections, and belongs to the field of satellite orbit thrust section inversion. Comprising the following steps: finding epoch moments of two observation arc segments; using a Laplace-LM algorithm to obtain an initial orbital element, and substituting the initial orbital element into POD to obtain a precise orbital element; constructing a satellite orbit dynamic model based on the initial dynamic parameters, and performing orbit propagation on a first precision orbit element of the first observation arc section from the first epoch moment to a second epoch moment of the second observation arc section to obtain a second propagation precision orbit element; constructing a target function according to the position difference between the second propagation precision orbit element and the second precision orbit determination result; introducing a pattern search algorithm to obtain a target dynamic parameter when the target function is minimum; according to the method, the thrust section can be accurately inverted, and the real propulsion strategy corresponding to the target dynamic parameter is summarized.

This invention provides an autonomous orbit-keeping method and system for low-Earth orbit inter-satellite laser communication constellations, comprising: generating upper and lower limits for satellite phase-keeping thresholds and periodically updating the thresholds based on the annual changes in laser link performance; recursively calculating the satellite's mean orbit elements at the current moment based on the autonomous positioning results of the onboard GNSS navigation receiver, and calculating the mean latitude argument deviation relative to the nominal orbit; autonomously calculating and generating an autonomous orbit control strategy onboard using a limit loop phase-keeping method; after generating the orbit-keeping strategy, transmitting it to the ground control center, and reviewing the orbit-keeping scheme autonomously generated onboard according to the integrated space-ground autonomous orbit control decision-making process; if the review is successful, the ground issues a permission command for orbit control implementation, and the satellite executes the orbit control operation according to the strategy; if the review fails, the ground re-enters the orbit-keeping strategy. This invention significantly reduces the complexity of ground system constellation management.

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 and system for interferometric imaging detection of satellite formations flying in the same direction in space. The method includes: setting initial parameters for a central satellite or virtual central satellite and N accompanying satellites in a formation; calculating the relative orbital elements of the N accompanying satellites; calculating the absolute orbital elements based on the relative orbital elements; predicting the position of each accompanying satellite based on the absolute orbital elements, and calculating the detection baseline formed by every two accompanying satellites; calculating the baseline distribution uniformity measurement function based on the detection baseline, using it as the objective function, and exploring optimization using an optimization algorithm to obtain the optimal solution; repeating the above steps based on the optimal solution to obtain a detection baseline with continuity and uniformity in both the radial and circumferential directions; and performing imaging inversion based on the detection baseline and the principle of interferometric imaging. This invention effectively improves the density of baseline coverage and reduces the deployment difficulty compared to deploying multiple satellites in a circle.

The present disclosure relates to a method and device for obtaining rigorous regression orbit parameters, and relates to the technical field of spacecraft, which can solve the problem of long time consumption of existing algorithms. The specific technical solution is: an orbit analytical method is adopted, and the J2 and J3 perturbation terms of the earth flattening are considered to obtain a set of rough regression orbit root numbers; a calculation function of the plane-instantaneous conversion is adopted to obtain the instantaneous orbit root numbers corresponding to the rough regression orbit root numbers; a numerical method is used to recursively correct the semi-major axis in the parameters of the instantaneous orbit root numbers, and the optimized initial value of the regression orbit root numbers is obtained; each orbit parameter and the regression period in the optimized initial value of the regression orbit root numbers are optimized and improved in a small range by using a nonlinear equation set, so that the rigorous regression orbit parameters with high regression accuracy are obtained. The present disclosure is simple and easy to understand in the calculation process, has high solving efficiency, short calculation time consumption, and can save a large amount of time for the design and analysis stage of the rigorous regression orbit.

For maneuvering a spacecraft, a method calculates a virtual point that represents a plurality of spacecraft orbiting in a spacecraft formation. The method iteratively calculates a desired trajectory relative to the virtual point for a given spacecraft of the plurality of spacecraft using a relative orbital elements (ROE) state x of the virtual point and the given spacecraft. Model predictive control maintains the given spacecraft within an ROE constraint for a given time interval subject to a fuel consumption. In response to determining a drift trajectory of the given spacecraft is outside of a trajectory tolerance of the desired trajectory, maneuvering the given spacecraft to within the trajectory tolerance using a maneuver trajectory determined using the model predictive control.

The invention discloses a remote sensing shooting window planning method and device, equipment, a medium and a product, and relates to the field of satellite shooting, and the method comprises the steps: obtaining six orbit elements of a satellite and coordinates of an expected shooting point; calculating an instantaneous three-dimensional position and a satellite speed of the satellite according to the six orbit elements; coordinate system conversion is carried out on the instantaneous three-dimensional position and the satellite speed, and satellite longitude and latitude height coordinates are determined; determining real-time attitude angle parameters of the satellite according to the converted instantaneous three-dimensional position, the converted coordinates of the expected shooting point and the satellite longitude and latitude height coordinates; according to the coordinates of the expected shooting point, the longitude and latitude coordinates of the satellite and the real-time attitude angle parameters, determining the rotation angle of a pan-tilt; the rotation angle of the holder is used for adjusting a remote sensing shooting window. According to the invention, the autonomy of the shooting window can be improved, and the time matching precision and the imaging positioning precision of earth observation are 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.

This invention discloses a high-dynamic tracking method and system for low-Earth orbit satellites. The method is as follows: Step 1: Obtain the two-line orbital element number TLE of the satellite, and calculate the theoretical position and angular velocity of the satellite at the current moment using the orbital propagation model, as a priori guidance for antenna pointing; Step 2: Construct a particle swarm in a two-dimensional search space, where each particle represents a possible antenna pointing angle; Step 3: Collect the received signal strength (RSSI) corresponding to the positions of multiple particles in the particle swarm, and fit the current signal strength gradient vector using the least squares method based on the geometric differences in the spatial distribution of particles; Step 4: Calculate the position of each particle at the next moment according to the generalized velocity update equation, which includes three driving forces: swarm intelligence term, cooperative gradient term, and orbital feedforward term; Step 5: Convert the updated optimal swarm position into a servo control command to drive the antenna rotation.

This application discloses a multi-objective optimization method and related apparatus for ultra-low Earth orbit (UEO) electric propulsion satellite constellations, relating to the field of satellite constellation optimization design. The method includes: constructing a multi-objective function with the satellite constellation's performance index function as the optimization objective; constructing a multi-objective design optimization problem model for the satellite constellation based on the satellite orbital dynamics model as constraints; classifying the performance index functions of the satellite constellation according to priority based on the index function priority classification principle to obtain the priority of each performance index function; solving the multi-objective design optimization problem model for the satellite constellation based on the highest priority performance index function to obtain an initial solution; iteratively optimizing the initial solution based on the priorities of all performance index functions to obtain a non-dominated solution set; determining the optimal orbital element parameter set based on the non-dominated solution set; and obtaining the optimized satellite constellation based on the optimal orbital element parameter set, thereby improving the speed, success rate, and robustness of the optimization solution.

The invention discloses a satellite-to-cluster target intersection orbit planning method for space debris removal. The method comprises the following steps: fitting a low-dimensional parameterized constraint model; generating all possible parameter combinations according to the orbital elements to be optimized in the parameterized constraint model and the parameter space and the step length of the orbital elements; performing non-thrust simulation on the virtual mother satellite of which the initial position is each parameter combination, and generating a passive rendezvous sequence formed by the virtual mother satellite and the space debris in the task period; all the passive rendezvous sequences are divided into a plurality of subsequences, the subsequences of the same time window are sorted in a descending order according to the number of space debris, and k subsequences ranked in the top are reserved; a sub-sequence is selected from each time window to form a candidate task path, all possible candidate task paths are generated, and then the candidate task paths meeting dynamic verification are reserved; and carrying out descending sorting on the reserved candidate task paths, and selecting and outputting the first Q candidate task paths.