47 results about "Relative orbit" patented technology

The invention relates to an autonomous navigation method for lunar satellite formation and a second-order filtering navigation architecture, and the autonomous navigation method comprises the steps: obtaining orbit prediction information based on a kinetic model of the lunar satellite formation, and obtaining collaborative theory observation information based on a collaborative observation model, carrying out fusion processing on orbit prediction information, collaborative theory observation information and multi-source observation data of in-orbit observation by using an extended Kalman filtering algorithm, and estimating an absolute orbit state of the lunar satellite formation; establishing a relative motion model of the slave satellite under the local coordinate system of the master satellite, and establishing a relative measurement model; and taking the absolute orbit state of the master satellite as a reference datum, obtaining a preliminary prediction relative orbit state based on the relative motion model of the slave satellite, and obtaining relative measurement data including distance measurement and angle measurement based on the relative measurement model, and fusing the relative measurement data and the preliminarily predicted relative orbit state through an extended Kalman filtering algorithm to estimate the relative orbit state of the slave satellite. Navigation may be independent of ground support.

The invention relates to the technical field of spacecraft orbit control and space safety, and discloses a space debris orbit prediction and avoidance method, which comprises the following steps that: a ground computing center screens candidate targets and generates enhanced data packets for uploading; the spaceborne computer combines real-time navigation data to carry out geometric feature screening, and an instant high-risk target list is generated; controlling the star sensor to execute optical observation, and resolving an optical measurement relative orbit state by utilizing displacement superposition and Kalman filtering; selecting a high confidence coefficient or a conservative probability threshold according to the optical measurement state acquisition condition, and generating an avoidance instruction when the collision probability exceeds the limit; and the propulsion system responds to the instruction to execute avoidance, verifies the effect by confirming the semi-major axis variation and executes track recovery by selecting an aircraft. According to the method, the ground-air collaborative screening and shift superposition enhancement technology is adopted, the problems that the low-precision ephemeris false alarm rate is high and the detection capability of a satellite-borne sensor on a dark and weak target is insufficient are solved, and the accuracy of evasion decision making and the execution reliability are improved.

The invention relates to a satellite group cooperative area positioning method based on communication ranging, and belongs to the technical field of spacecraft navigation. Three reference satellites are selected in a satellite group space area, and a positioning reference is constructed; determining an initial relative motion state Xim0 of any member satellite m in the LVLH system based on a positioning reference; a triangular vector closed loop is formed by the relative positions of the satellite i, the satellite j and the satellite k; according to the triangular vector closed loop, establishing a relative orbit state consistency constraint model; introducing the consistency constraint model into a filtering algorithm, establishing a distributed Kalman filtering algorithm, and obtaining a measurement updating equation, namely star group distributed state estimation; according to the invention, cluster dynamic ad hoc network relative navigation is realized on the premise of limited satellite-borne measurement and control resources and complex space optical environment, the global positioning navigation problem under limited information interaction is solved, and the method has the capability of fast and accurate area positioning of an in-and-out satellite group system and global state sensing under limited resources.

The invention provides a laser terminal calibration method and system based on inter-satellite orbit mutual transmission, and the method comprises the steps: carrying out the static calibration of a laser terminal through employing a star sensor or ground assistance before the establishment of an inter-satellite communication link, and enabling the pointing error of an optical axis to be smaller than 0.1 degree; after a communication link is established, orbit parameters of two satellites are exchanged in real time, a relative orbit normal vector is calculated through coordinate transformation, and a fusion reference is dynamically generated in combination with a static calibration result. According to the method, the real-time track mutual transmission data is used as dynamic correction input, the initial static calibration result is not only relied on any more, and fusion benchmark optimization of calibration and track mutual transmission is achieved. Compared with a traditional static calibration method depending on a ground station or a single-star inertial reference, the method breaks through the precision bottleneck of static calibration through the fusion of the orbit dynamic data and the static calibration result, effectively eliminates the orbit dynamic error, improves the laser calibration precision to a micro-radian level, and improves the calibration precision. And the establishment efficiency and reliability of the inter-satellite laser link are greatly improved.

The application provides a GNSS / inter-satellite link distributed cooperative real-time orbit determination method and device for a mega constellation, first, each satellite in a cluster performs preliminary orbit determination based on local GNSS observation data and IMU data to obtain preliminary determination results of each satellite in the cluster; then, a cluster head satellite collects the preliminary determination results to determine relative orbit parameters in the cluster; then, a backbone node collects the relative orbit parameters to determine a regional absolute orbit reference; a global anchor satellite determines unified constellation orbit parameters; finally, according to the regional absolute orbit reference and the unified constellation orbit parameters, orbit determination results of each satellite are determined. Through a three-level hierarchical cooperative architecture, the application disperses data processing and determination tasks to different level nodes, avoids data processing bottlenecks of a central node and communication link congestion problems in centralized technology, reduces complex interaction between nodes and inter-satellite link resource occupation in a thousand-star level scale in traditional distributed technology, and shortens iteration convergence time.

The invention specifically discloses a relative orbit dynamics multi-dimensional ground simulation method using a ducted fan, and relates to the technical field of spacecraft ground simulation. According to the method, a ducted fan compensation platform is designed, the ducted fan compensation platform comprises a ducted fan, a main structure frame and a fan driving unit, the ducted fan compensation platform is integrally fixed to the air floatation simulator, and the method comprises the specific steps that a target spacecraft reference coordinate system and a simulator body coordinate system are established, and a space-time scaling model is determined; a simulator pose is obtained in real time through a pose measurement system; starting a jet system to generate relative displacement, and inversely calculating an actual space relative distance and a nonlinear acceleration by combining a C-W equation and a space-time scaling model; the ducted fan is controlled to provide xoy plane nonlinear compensation force, and z-axis direction compensation force is calculated through mathematical simulation coupling. The problem that a traditional air floating platform cannot simulate the dynamic characteristics and multi-dimensional motion of a spacecraft orbit is solved, and the precision of a spacecraft ground simulation test is remarkably improved.

The application discloses a kind of star group between heterogeneous measurement data association and high-precision fusion orbit determination method, belong to space relative navigation technical field.The method includes grouping to the member star in star group, constructs the measurement set of the angle measurement and ranging of all member stars in time t, then calculates the position assumption set of member star, calculates the distance assumption value between member stars under each association assumption, constructs association degree evaluation index by distance assumption value and actual distance measurement, constructs cost matrix according to association degree evaluation index, and then obtains optimal association result, obtains the ranging and angle measurement data corresponding to all member stars at current time relative orbit determination, including using optimal estimation method to carry out state estimation, and including determining system state, constructing system measurement equation and system equation;Finally, the above steps are repeatedly executed for each measurement time, until filter convergence, obtain the high-precision orbit determination result of member star.

The invention relates to a lunar satellite formation relative navigation method based on factor graph optimization, which comprises the following steps: in a sliding window with a set length, constructing a factor graph by using relative orbit states of a series of satellite formations as state nodes and constraint factors connected with the state nodes; and optimizing the factor graph to obtain the optimal estimation of the relative orbit state of the satellite formation at the current moment. According to the relative navigation method, historical data are utilized, the historical orbit state is optimized and corrected again in the process of optimizing the factor graph, error accumulation is effectively restrained, and navigation precision and robustness are remarkably improved.

The application discloses a satellite orbit transfer strategy planning method and device, equipment and a storage medium. The method comprises the following steps: initializing an initial time of a satellite and performing orbit period extrapolation to update a t time to an extrapolated time; initializing an orbit transfer flag as not completed, calculating a first orbit parameter of the satellite at the t time and a second orbit parameter of a target virtual satellite at the t time, and determining a relative orbit parameter of the satellite to the target virtual satellite; determining each orbit parameter correction flag according to the relative orbit parameter and a preset orbit threshold value, and determining a corresponding orbit control sub-mode based on the each orbit parameter correction flag; when the orbit transfer flag is completed, determining a firing parameter of the satellite based on the orbit control sub-mode, and performing orbit transfer maneuver and orbit transfer phase adjustment processing. The scheme can quickly select the corresponding orbit control sub-mode according to the current orbit transfer requirement, quickly generate the calculation and simulation verification of the electric propulsion orbit transfer parameter, and make the satellite accurately enter the predetermined orbit position.

The present application discloses a method, device, equipment and medium for controlling formation flight of ultra-low-orbit electric propulsion satellites, which relates to the field of satellite flight control technology. The method comprises: establishing a relative orbital dynamics model under the action of perturbation force based on the absolute orbital dynamics model of the ultra-low-orbit electric propulsion satellite under the action of perturbation force, and combining the relative motion mathematical model under the action of control force to establish a relative motion mathematical model under the joint action of the two forces; then combining the Hamiltonian function, co-state equation and canonical equation to establish a mathematical model of the orbit optimization control problem of the ultra-low-orbit electric propulsion satellite formation flight; converting the mathematical model of the above-mentioned optimization control problem into a two-point boundary value problem; based on special points, analytically calculating the initial values ​​of the co-state variables and the initial values ​​of the mass co-state; and determining the optimal control scheme according to the initial values ​​of the co-state variables, the initial values ​​of the mass co-state and the mathematical model of the two-point boundary value problem. The present application can improve the solution success rate, convergence speed and control accuracy of the optimization control.

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 application relates to a kind of space spin wobble non-cooperative target safety approach control method.The method includes the following steps: establishing the dynamics and kinematics model of double spacecraft relative attitude error and relative orbit error, and compacting relative attitude error and relative orbit error model into six degrees of freedom form;By defining the constraint function of service spacecraft safety and motion boundary, constraint auxiliary variable, construct spacecraft auxiliary system for rendezvous and docking and consider safety constraints;Design shortest trajectory sliding mode variable for spacecraft auxiliary system, the shortest trajectory sliding mode variable makes service spacecraft attitude and relative orbit from initial state to the desired attitude orbit required by task with the shortest path maneuver;Design nonsingular auxiliary switching function, avoid singular phenomenon when the derivative of shortest trajectory sliding mode variable is solved about time;Build safety approach controller for space spin wobble non-cooperative target considering safety constraints.

The invention relates to a space spinning nutation non-cooperative target safe approaching control method. The method comprises the following steps: establishing dynamics and kinematics models of the relative attitude error and the relative orbit error of the double spacecrafts, and compacting the relative attitude error and relative orbit error model into a six-degree-of-freedom form; the method comprises the following steps: defining constraint functions and constraint auxiliary variables of safety and motion boundaries of a service spacecraft, and constructing a spacecraft auxiliary system which is used for rendezvous and docking and takes safety constraints into consideration; designing a shortest trace direction sliding mode variable for the spacecraft auxiliary system, wherein the shortest trace direction sliding mode variable enables the attitude and the relative orbit of the service spacecraft to maneuver to an expected attitude orbit required by a task from an initial state in a shortest path; a non-singular auxiliary switching function is designed, and the singular phenomenon is avoided when a time-related derivative is solved by a shortest trace direction sliding mode variable; and constructing a security approximation controller considering security constraints for the spatial spin nutation non-cooperative target.

The invention provides an angle-measurement-only multi-pulse self-adaptive closed-loop rendezvous guidance method based on second-order cone programming, and belongs to the relative navigation and rendezvous guidance technology of spacecrafts. According to the method, firstly, an angle-only-measurement relative navigation dynamical model and a measurement model based on a CW equation are constructed under LVLH coordinates, system state observability analysis is carried out, the orbit maneuvering condition enabling the state to be observable is obtained, and the condition is used for constructing an optimal guidance index. Secondly, analyzing and modeling the rendezvous guidance error source, and deducing to obtain a relative orbit guidance error covariance model through error covariance analysis; thirdly, on the basis of a second-order cone programming method, guidance fuel consumption and system state observability conditions are fused, and a multi-target weighted cost function of optimal rendezvous guidance is jointly constructed; meanwhile, based on the established guidance error covariance model, a cost function weight index system with a self-adaptive adjustment function is designed, and real-time self-adaptive adjustment of a guidance strategy along with change of system performance is achieved. The method solves the problems of high fuel consumption, poor flexibility and the like caused by separate design of the relative navigation and the rendezvous guidance model in the traditional only-sight-line measurement rendezvous guidance method, and has the ability of adaptively adjusting and optimizing the rendezvous fuel consumption and the relative navigation system observability.

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.

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 embodiment of the application discloses a two-degree-of-freedom discontinuous rotation control method and system for a solar wing, the discontinuous rotation control method comprises the following steps: obtaining the angle between a sun vector and the normal of a solar wing along the direction of orbit operation and the angle between the sun vector and the normal of the solar wing in the direction perpendicular to the orbit plane, so as to obtain the angle between the sun vector and the normal of the solar wing in a first coordinate projection plane and a second coordinate projection plane, and further obtain the driving target instruction of a first control shaft and a second control shaft; determining the position of the satellite, if the satellite is in a sunlit area, driving the first control shaft to make the sun vector coincide with the normal of the solar wing in the first coordinate projection plane, then driving the second control shaft to adjust the position angle of the solar wing relative to the orbit plane, so that the normal of the solar wing coincides with the sun vector in the second coordinate projection plane; if the satellite is in a shadow area, driving the first control shaft to execute a rotation instruction.

The present invention relates to a micro-nano satellite relative orbit PID sliding film control method and system based on particle swarm optimization. The method includes: establishing a nonlinear relative orbital dynamics model, i.e., a dynamic equation, between a reference satellite and a maneuvering tracking satellite under interference; establishing an ideal orbit for the maneuvering tracking satellite based on various constraints of the micro-nano satellite maneuver; introducing an error between the maneuvering tracking satellite orbit and the ideal orbit, and establishing a sliding film surface using PID control; performing particle swarm optimization on the sliding film surface parameters; and when the orbital state of the disturbed micro-nano satellite is near the sliding film surface, under the action of a control function, causing the orbital state of the disturbed micro-nano satellite to reach the sliding film surface and reach a state to be tracked on the sliding film surface. Based on the fact that the sliding mode is independent of object parameters and interference, the present invention is insensitive to parameter changes and disturbances. Even if the maneuvering satellite deviates from the ideal trajectory and appears near the sliding film surface, under the action of the control function, the system reaches the sliding film surface and finally reaches a state to be tracked on the sliding film surface.

The application discloses a spacecraft formation configuration decoupling control method and belongs to the technical field of satellite formation control. The application aims at solving the problems of serious approximation error accumulation and high fuel consumption of a spacecraft formation control method based on a CW equation. The method comprises the following steps: calculating relative orbit elements of a sub-satellite and a main satellite, obtaining a radial relative motion equation, a trace relative motion equation and a normal relative motion equation of the sub-satellite and the main satellite, and obtaining a modified equation; determining a velocity pulse according to a dynamic equation between the relative orbit elements and the velocity pulse; determining a required change amount of the relative orbit elements, and then determining a required change amount of a relative orbit inclination vector, a required change amount of a relative semi-major axis, a required change amount of a relative eccentricity vector and a required change amount of a relative angular distance; controlling the relative orbit inclination vector by adopting a single-pulse normal control strategy; controlling the relative semi-major axis and the relative eccentricity vector by adopting a double-pulse trace pulse strategy; and controlling the relative angular distance by adopting a multi-period trace control strategy. The application is used for formation control.

This invention discloses an autonomous approach guidance and control method, electronic equipment, and medium for spacecraft. It determines the approach control sequence and dwell time based on the spacecraft's relative orbit to the target, nominal approach position, attitude maneuver allowance, and nominal guidance transfer time. Then, it corrects the guidance transfer time and adjusts the approach control sequence accordingly. Next, it determines the dwell control sequence based on the spacecraft's relative orbit to the target, upper and lower limits of the nominal dwell time, and lateral acceleration. Finally, it determines the lateral control strategy, normal control strategy, and radial control strategy based on the spacecraft's relative orbit to the target and attitude maneuver allowance. This method combines fixed-time CW guidance with independent lateral, normal, and radial three-axis guidance, achieving high-precision approach control within a given time range.

The application relates to the technical field of satellite autonomous control, and discloses a satellite round game pursuit-escape game decision method based on Alpha-Zero, a terminal and a medium. The method sets physical constants and game rules, establishes a relative orbit motion model of satellites of both sides of pursuit and escape based on a C-W equation, and thus constructs a satellite round game pursuit-escape game environment; a strategy value neural network based on an Alpha-Zero framework is constructed, and when impulse orbit change decision needs to be made in each round, Monte Carlo tree search is performed based on the satellite round game pursuit-escape game environment and the strategy value neural network; the parameters of the strategy value neural network are iteratively optimized through self-game iteration, and an optimal strategy value neural network is obtained; after the real-time collected satellite orbit state is input into the optimal network after feature coding, an impulse orbit change strategy is generated and the satellite is controlled to execute. The application realizes efficient autonomous game of the satellite in a complex space environment.

The invention provides a semi-physical full-process ground verification system for satellite in-orbit vision autonomous navigation. The semi-physical full-process ground verification system comprises a dynamics simulation system, an image simulation system and an autonomous navigation system. Wherein the image simulation system generates scene photo data and transmits the scene photo data to the autonomous navigation system; the autonomous navigation system performs image recognition and positioning to give navigation information, and transmits the navigation information to the dynamic simulation system; the dynamics simulation system simulates on-orbit control motion and transmits relative orbit information to the image simulation system; and the image simulator simulates and generates a new scene photo according to the new track information. Therefore, the ground closed-loop verification of the semi-physical whole process is completed. According to the invention, a real satellite service computer and a real photoelectric load can be accessed, so that the whole semi-physical process can be simulated in an on-orbit real working state; through high-precision dynamics, satellite in-orbit motion physical characteristics are simulated, and autonomous navigation is carried out through high-precision image simulation.

The application discloses a kind of space reverse satellite formation interference imaging detection method and system of accompanying flight, the method comprises: setting center satellite or virtual center satellite and the initial parameter of N satellite formation satellite of reverse accompanying flight;The relative orbit element of N satellite formation satellite of accompanying flight is calculated;Absolute orbit element is calculated according to relative orbit element;According to absolute orbit element, the position of each satellite of accompanying flight is forecasted, and the detection baseline formed by every two satellite of accompanying flight is calculated;Baseline distribution uniformity measure function is calculated according to detection baseline, as objective function, optimization algorithm is used to explore optimization, and optimal solution is obtained;Based on optimal solution, the above steps are repeated, and the detection baseline with continuity and uniformity in radial and circumferential direction is obtained;According to the detection baseline obtained, imaging inversion is carried out based on interference imaging principle.The application effectively improves the rapid coverage of long-short baseline, and short baseline is more dense.

The present invention provides a relative motion adaptive control method for the close-range detection mission of small spacecrafts, which is characterized by the following steps: Step 1, regarding the close-range detection object as the target spacecraft and the small spacecraft as the service spacecraft, and constructing a relative orbit dynamics model between the two; Step 2, designing a full-process nominal flight trajectory according to the requirements of the close-range detection mission, including parameters of some key hover points, transfer trajectories between hover points, and transfer times; Step 3, respectively representing the relative orbit dynamics model and the nominal flight trajectory in state space; Step 4, regarding the nominal flight trajectory as the reference model and the relative orbit dynamics model as the controlled object, and designing a relative motion adaptive controller by using the adaptive control theory SAC. The present invention still has good control effects under the influence factors such as the existence of modeling uncertainty terms, unknown disturbances, and limited control capabilities in the controlled object.

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

The invention discloses a real-time control method and device for a limited energy reachable set of a spacecraft, a computer and a storage medium, belongs to the technical field of aerospace orbit control, and solves the problems of low real-time performance and result precision of the limited energy reachable set caused by insufficient high-dimensional and nonlinear characteristics of a spacecraft dynamic model in the prior art. And an analytical solution is difficult to directly obtain. The control method comprises the following steps: establishing a relative orbit dynamics system model; defining a finite energy reachable set; constructing a state uncertainty and control driving range model; establishing a gain set based on the model, and describing an initial shape of a finite energy reachable set; the Minkowski sum is used for carrying out equivalent conversion on the gain set, and a finite energy reachable set shape ellipsoid is obtained; and defining an ellipsoid shape function, and obtaining a finite energy reachable set optimal approximate solution through an ellipsoid trace optimal internal and external approximation method. The method is suitable for the technical field of spacecraft orbit dynamics and autonomous control, and can be applied to spacecraft orbit control research and development.

The application relates to a kind of lunar satellite formation autonomous navigation method and second-order filter navigation architecture, and autonomous navigation method includes: obtaining orbit prediction information based on the dynamics model of lunar satellite formation, obtaining collaborative theory observation information based on collaborative observation model, using extended Kalman filter algorithm to carry out fusion processing to orbit prediction information, collaborative theory observation information and on-orbit observation multi-source observation data, and estimating the absolute orbit state of lunar satellite formation;Establish the relative motion model of slave star under the local coordinate system of main star, and establish relative measurement model;And with the absolute orbit state of main star as reference datum, obtain preliminary prediction relative orbit state based on the relative motion model of slave star, obtain relative measurement data containing ranging and angle measurement based on relative measurement model, estimate the relative orbit state of slave star by fusing relative measurement data and preliminary prediction relative orbit state through extended Kalman filter algorithm.Navigation can not rely on ground support.

The invention discloses a heterogeneous measurement data association and high-precision fusion orbit determination method among star groups, and belongs to the technical field of space relative navigation. The method comprises the following steps: grouping member stars in a star group, constructing a measurement set of angle measurement and distance measurement of all the member stars within time t, then calculating a position hypothesis set of the member stars, calculating a distance hypothesis value between the member stars under each association hypothesis, constructing an association degree evaluation index through the distance hypothesis value and actual distance measurement, constructing a cost matrix according to correlation degree evaluation indexes, further obtaining an optimal correlation result, obtaining distance measurement and angle measurement data corresponding to all member satellites at the current moment, then carrying out relative orbit determination, adopting an optimal estimation method to carry out state estimation, determining a system state, and constructing a system measurement equation and a system equation; and finally, repeatedly executing the steps at each measurement moment until filtering convergence is carried out, thereby obtaining a high-precision orbit determination result of the member satellites.

The invention relates to a multi-low-orbit satellite cooperative communication method based on spectrum sensing assistance, which comprises the following steps: in a multi-low-orbit satellite cooperative system, a satellite with a relatively low relative orbit height senses a satellite with a relatively high relative orbit height by applying a spectrum sensing technology, and satellite-ground communication is carried out based on spectrum sensing assistance. The method specifically comprises the following steps: embedding signal cyclic delay characteristics at a sending end by adopting a cyclic delay diversity technology, and performing corresponding processing at a receiving end. According to the invention, through real-time efficient detection and utilization of the spectrum hole, the satellite-ground communication link quality and the transmission performance are effectively improved.

The embodiment of the invention provides a method and system for determining an initial relative orbit of a space non-cooperative target, and the method comprises the steps: shooting a sequence image of the space non-cooperative target through an optical camera carried by a service spacecraft for the space non-cooperative target; the sequence image is shot at the initial moment when the self-service spacecraft finds the space non-cooperative target; determining a kinetic equation and a measurement equation of a relative motion state of the spatial non-cooperative target under a spherical coordinate system based on the sequence image; and constructing a singular value of the matrix to solve the kinetic equation and the measurement equation of the relative motion state to obtain an initial relative orbit of the spatial non-cooperative target. Through singular value decomposition, an efficient approximate solution of an initial relative orbit determination problem constructed by shooting sequence images of the spatial non-cooperative target through an optical camera carried by a service spacecraft can be obtained, efficient calculation is achieved through solution, and high-precision initial relative orbit determination is achieved.