60 results about "Satellite orbit determination" patented technology

The invention discloses a forecasting method and device for an atmospheric resistance coefficient of a low-orbit satellite in the technical field of satellite navigation. The forecasting method comprises the following steps: acquiring satellite orbit determination data and a space environment index in a preset time period before forecasting; according to the satellite orbit determination data, performing simplified dynamic precise orbit determination on a low-orbit satellite, and fitting an atmospheric resistance coefficient in the simplified dynamic precise orbit determination process; performing data preprocessing on the space environment index and the atmospheric resistance coefficient and generating a sample set; constructing an LSTM prediction model, and training, adjusting and evaluating the LSTM prediction model through the sample set; predicting an atmospheric resistance coefficient in a future preset time period according to the finally obtained LSTM prediction model; and applying the atmospheric resistance coefficient in the future preset time period to the atmospheric resistance perturbation model of the low-orbit satellite for orbit forecasting. According to the method, the atmospheric resistance coefficient in the LEO satellite orbit prediction process is predicted in a high-precision manner through the LSTM prediction model, so that the orbit prediction precision is effectively improved.

The invention provides a low-orbit navigation constellation autonomous orbit determination method and system under a weak inter-satellite topological structure, and relates to the technical field of satellite orbit determination, and the method comprises the steps: obtaining various observation data of an inter-satellite link on a predetermined orbit, and carrying out the preprocessing of the various observation data; establishing a synchronous observation equation of orbit and clock correction parameters by adopting a segmented linear polynomial satellite clock correction modeling method, generating orbit, clock correction and related error parameters combining satellite-borne GNSS and inter-satellite link data, and forecasting; based on an error analysis result, multiple key features are extracted from the generated orbit, clock error and related error parameters to serve as input of a parameter prediction model, and a parameter time sequence of four types of prediction including relative weight, inter-satellite link hardware delay, low-orbit satellite clock error and unmodeled error is obtained through output; and uploading the predicted parameter time sequence and the broadcast ephemeris to a main satellite together, and transmitting to each sub-satellite by the main satellite through an inter-satellite link to realize distributed autonomous orbit determination of the sub-satellites.

The invention discloses a multi-satellite layered cooperative task planning method for multi-target situation awareness, and belongs to the field of aerospace. The implementation method comprises the following steps: defining a multi-satellite collaborative planning problem of a multi-target sensing task as a quintuple, and determining satellite resources and time constraint conditions in an inter-satellite collaborative planning problem of the multi-target sensing task; constructing a multi-target sensing task multi-satellite collaborative planning model, describing elements, related constraints and simplified conditions of a multi-target sensing task multi-satellite collaborative planning problem, analyzing and decoupling multiple coupling complex constraints, and designing a layered bidding and tendering collaborative mechanism; on the basis, a multi-target situation awareness hierarchical collaborative task planning method is constructed, a two-dimensional coding structure related to task types is constructed according to multi-dynamic target orbit determination requirements, a mixed profit value function related to a double-satellite orbit determination observation configuration, attitude maneuver cost, double-satellite collaborative rewards and the like is designed, and multi-satellite collaboration is guided. And a multi-satellite collaborative observation sequence is obtained, and multi-satellite layered collaborative task planning is realized.

The invention provides a GEO communication satellite uplink signal Doppler frequency compensation method and device. The method comprises the following steps: acquiring a satellite orbit determination result corresponding to a GEO communication satellite; determining a Doppler compensation frequency curve according to the satellite orbit measurement result and the local frequency of the frequency conversion of a satellite transponder configured in the GEO communication satellite; and compensating a first uplink signal to be transmitted from the ground station to the GEO communication satellite by using the Doppler compensation frequency curve through the compensation frequency mixer, and transmitting the compensated first uplink signal to the GEO communication satellite. According to the invention, under the condition that the user receiver is non-inductive, the carrier frequency and the spreading code rate of the user receiver are changed in equal proportion, so that the user receiver can adopt a receiving algorithm of the carrier-assisted spreading code rate to improve the sensitivity.

The invention discloses a lunar satellite orbit determination method and device based on radio and laser ranging, and relates to the technical field of lunar satellite orbit calculation, and the method comprises the steps: obtaining observation data of a radio tracking station and a laser station; establishing a high-precision kinetic model of the lunar satellite; establishing a lunar satellite radio and laser ranging mathematical model, and processing observation data according to an observation equation to obtain a kinetic parameter resolving value of the satellite at the initial orbit determination moment; and substituting the kinetic parameter resolving value into the satellite shot motion equation, and resolving to obtain an orbit determination result of the lunar satellite. According to the method, radio ranging and laser ranging are combined for orbit determination, precise orbit determination of the lunar satellite is achieved, compared with a traditional orbit determination mode, the precision and reliability are greatly improved, the requirements for the arrangement number and geometric distribution of ground stations are reduced, and the measurement and control pressure of existing ground stations is relieved.

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 discloses a satellite orbit determination method and device, a storage medium and electronic equipment. The method comprises the following steps: acquiring an initial satellite position of a target communication satellite under the condition that the target communication satellite can receive a navigation signal, and determining a reference orbit of the target communication satellite based on spatio-temporal information of a navigation satellite; obtaining a target satellite position and a target non-conservative force model coefficient through orbit fitting; performing orbit integration to determine an integration orbit; comparing the integral orbit with the reference orbit to determine a position error sequence and process noise covariance information; and under the condition that the target communication satellite cannot receive the navigation signal, performing orbit filtering by using the inter-satellite link data, and determining a target operation orbit. The technical problem that a satellite cannot run according to a correct orbit due to the fact that a satellite constellation integrally deviates and rotates is solved.

The invention discloses an orbit determination method and device for a low-orbit satellite, the low-orbit satellite and a storage medium, and the method comprises the steps: obtaining global navigation satellite system (GNSS) observation data and broadcast ephemeris received by a receiver in the low-orbit satellite, and obtaining orbit state information in a PPP-B2b signal broadcasted by a Beidou No.3 system; preprocessing the GNSS observation data according to an observation model based on the GNSS observation data to obtain preprocessed GNSS observation data, and determining first position information of a GNSS satellite according to the broadcast ephemeris and the orbit state information; and determining orbit position information of the low-orbit satellite according to the preprocessed GNSS observation data and the first position information. Therefore, the low-orbit satellite can complete high-precision orbit positioning under the condition of limited computing power.

The invention provides an autonomous orbit determination method and system for satellite formation based on space-based VLBI, and a readable medium. The autonomous orbit determination method comprises the following steps: acquiring time delays and time delay rates of all observation baselines in a satellite formation; obtaining an extrapolation orbit of the satellite according to the kinetic model of the satellite and the orbit information of the initial epoch of the satellite; establishing a first dynamic orbit model, and parameterizing the first dynamic orbit model according to the extrapolation orbit to obtain a second dynamic orbit model; constructing an observation equation according to the time delay, the time delay rate and the second dynamic orbit models of all satellites; constructing a virtual observation equation according to the orbit information of the initial epoch; the observation equation and the virtual observation equations of all satellites are combined, and the correction number of orbit determination parameters of each satellite is solved; and correcting the orbit determination parameter according to the correction number to obtain a corrected orbit determination parameter. Through the autonomous orbit determination method, the orbit determination precision can be improved, the dependence on external information is reduced, and the autonomy of satellite orbit determination is improved.

A low-orbit satellite orbit determination method considering GPS elastic power comprises the steps that firstly, a carrier-to-noise ratio template is used for detecting the opening time period of the elastic power of a GPS satellite, then a non-geometric combination of satellite-borne observation values in the opening and closing periods of the elastic power is extracted, and a three-time 98% quantile is used for determining a non-geometric combination (GF) epoch difference threshold value in the opening and closing periods of the elastic power; for the signal-to-noise ratio jump, a signal-to-noise ratio jump detection amount DR based on the space between satellites is constructed according to a signal power propagation formula, then the distribution condition of the DR in the elastic power starting period is counted, a signal-to-noise ratio jump detection threshold value is determined by using a triple standard deviation criterion, and finally the ambiguity continuity is detected by combining the two methods. The method is designed for change characteristics of satellite-borne observation data of the low-orbit satellite under elastic power starting; the method is suitable for real-time and afterwards low-orbit satellite orbit determination, and can effectively improve the low-orbit satellite orbit determination precision in the GPS elastic power starting period.

The invention relates to the technical field of satellite orbit determination, in particular to a low-orbit giant constellation autonomous orbit determination method based on a star sensor and inter-satellite angle measurement. The method comprises the following steps: step 1, determining a constellation configuration by using inter-satellite quantity measurement: expressing each satellite state quantity in a constellation by using six elements, constructing a partial derivative matrix of inter-satellite quantity measurement relative to state information, constructing a linear equation set according to the state quantity, the quantity measurement and the partial derivative matrix, and determining a constellation configuration; and solving the satellite state quantity describing the constellation configuration through an improved least square method. Step 2, correcting constellation drift by using a star sensor and a camera quantity measurement and combining the constellation configuration in the step 1: acquiring a true value of an included angle between an adjacent satellite direction vector and a certain fixed star direction vector by using the star sensor and the camera, then calculating an estimated value of the included angle between the adjacent satellite direction vector and the certain fixed star direction vector, and correcting the constellation drift; and the difference between the real value and the estimated value of the included angle is minimized to realize constellation correction. By adopting the scheme, orbit determination can be realized without relying on a ground observation station, and data is easy to manage.

A satellite orbit determination method and device based on space-based observation, which relates to the field of computer simulation. It solves the problems that domestic aerospace simulation software is limited in terms of accuracy and generality, and that when the Laplace method and the Gauss method encounter extremely short arcs of dozens of seconds or even more than ten seconds, large errors often occur between the solution and the true value due to the intrinsic ill-conditioning when solving equations. The method includes: using Spacesim to randomly plan and simulate two satellite orbits; using an observation satellite to photograph a target satellite on the tracking satellite orbit to obtain a series of images of the target satellite; using Opencv to identify and process the series of images of the obtained target satellite to determine the trajectory of the target satellite obtained by the observation satellite; and processing the short-arc orbit determination problem of the trajectory of the target satellite obtained by the observation satellite according to the particle swarm algorithm to determine the target satellite orbit. It is applied to the field of space-based observation.

The present invention proposes a method for modeling the ranging error between navigation satellites, belonging to the field of ranging between navigation satellites. First, precise orbit determination and precise clock error estimation are performed using the inter-satellite link observation data, and the inter-satellite orbit determination residuals and clock error estimation residuals are output. Secondly, the residual sequences are normalized to the same orbital period, and a high-order polynomial with a semi-orbital period term is used for fitting. Subsequently, through the F-test and t-test, the overall effectiveness of the model and the significance of each coefficient of the model are checked. Finally, the ranging error model and the ranging error model of the reverse-normalized one-way link are output. According to this process, the ranging error modeling work of the one-way observation value of any link can be completed. The present invention can effectively improve the inter-satellite link observation model, reduce the observation value residuals of orbit estimation and clock error solution, and meet the requirements of high-precision precise orbit determination and clock error estimation.

The invention relates to the technical field of satellite orbit determination, in particular to a short window satellite orbit determination method based on a ground strategy and weight adjustment. Comprising the following steps: step 1, constructing a partial derivative matrix of quantity measurement relative to state information and a state transition matrix of a state quantity; 2, on the basis of weight adjustment, satellite-ground measurement data optimization is carried out, and high-quality observation arc segments are screened; 3, constructing a relative position information equation set by using inter-satellite measurement data, and establishing a drift equation set by combining the estimation parameter of the constellation drift and satellite-ground angle measurement data; correcting the constellation drift distance based on ground station angle measurement data; and 4, carrying out simultaneous solution on the relative position information equation set and the drift amount equation set to obtain an orbit determination result. According to the method, the problems that continuous long-time window earth measurement is needed during satellite orbit determination and the distribution of observation arc sections of an orbit phase observation station in a short-window orbit determination task is uneven are solved, and the accuracy and the flexibility of the orbit determination task under various windows can be effectively improved.

The present disclosure relates to the technical field of satellite communication, and particularly relates to a satellite orbit determination method based on edge computing, a device and an application, wherein the method comprises calculating the position coordinates and velocity coordinates of a satellite by using the edge computing capability on the satellite, transmitting the position coordinates and velocity coordinates to a ground station, and taking the position coordinates and velocity coordinates as observations by the ground station, and optimizing the observations by a two-step method for precise orbit determination and prediction. Through the embodiments of the present disclosure, the availability of orbit determination and prediction is improved when the satellite-ground measurement and control resources or the operation and control center computing resources are limited.

The present disclosure provides a verification method for satellite orbit determination, including: dividing each on-orbit satellite into different categories according to the orbit altitude of each on-orbit satellite, and determining the orbit determination result of each on-orbit satellite; the orbit determination result contains multiple parameters; automatically verifying the multiple parameters in the orbit determination result with a preset adjustable threshold through machine learning to determine whether the orbit determination result meets the orbit determination accuracy requirements; if the orbit determination result meets the orbit determination accuracy requirements, directly publish the orbit determination result; if the orbit determination result does not meet the orbit determination accuracy requirements, continue with the orbit determination optimization strategy, repeat the satellite orbit determination and automatic verification process until the orbit determination result meets the orbit determination accuracy requirements, and continue to complete the orbit determination result publishing process; or transfer to manual processing after the orbit determination optimization strategy can no longer solve the problem. The present disclosure shortens the time for large-scale satellite orbit determination and product release, improves the operation efficiency of the satellite orbit system; and reduces the workload of operators.

The invention relates to a multi-radar synchronous satellite orbit determination resource scheduling method based on reinforcement learning, and the method comprises the following steps: (10) radar perception situation acquisition: a radar acquires a networking radar state and a satellite motion orbit based on historical information, and completes the parameter collection; (20) measuring and calculating a satellite observation geometric positioning precision factor by networking radars: detecting satellites in sequence based on different radar combinations, and calculating geometric positioning precision factors under different strategies; (30) resource scheduling scheme characterization and problem modeling: modeling a time sequence optimization problem based on detection capability constraints of multiple radars, and realizing high-precision target orbit determination; and (40) generating a multi-radar synchronous satellite orbit determination resource scheduling scheme: solving the time sequence optimization problem to obtain the multi-radar synchronous satellite orbit determination resource scheduling scheme, and issuing the scheme. According to the method, rapid solving can be realized, the GDOP maximum value of all targets in the whole scene is reduced, and the orbit determination precision is further improved.

The application discloses a kind of based on parameter self-adapting time-invariant track information acquisition method, comprising: for the orbit equation of spacecraft motion, construct linear random time-invariant system, determine initial state quantity and external disturbance term;Based on initial state quantity and external disturbance term, obtain constant coefficient matrix and constant disturbance input matrix;Input the parameter self-adaptive Runge-Kutta solver established, obtain the list of to-be-solved orbit parameters;According to the adaptive parameter constraint formula, the dependent relationship of each parameter in parameter list is solved, free parameters therein are optimized according to genetic algorithm, iteratively updated until the optimal parameters of determination standard are obtained, these parameters are used as the parameters of Runge-Kutta solver, and the orbit information is iteratively solved by Runge-Kutta solver to realize the acquisition of orbit information.The application can be widely applied to space science and technology satellite orbit solving aspect, and is not only suitable for single scientific satellite orbit determination, but also suitable for scientific satellite constellation orbit formulation.

The invention provides an observation station system error correction method and system, computer equipment and a storage medium, and belongs to the field of spaceflight measurement and control, and the method comprises the steps: respectively obtaining observation matrixes of a laser ranging station SLR and a deep space network observation station, building observation equations of laser ranging and deep space network ranging and speed measurement, building an orbital motion equation for a satellite, and obtaining an orbital motion equation; superposing the SLR observation equation, the UXB observation equation and the orbital motion equation according to weights to construct a normal equation set; calculating theoretical prediction matrixes of the SLR and a deep space network observation station by combining an orbital motion equation, the SLR and an observation equation of the deep space network, calculating respective residual matrixes according to the difference between the prediction matrixes and the observation matrixes, and solving to-be-estimated parameter correction of an equation set by using a least square method; and comparing the orbit correction with a threshold value, obtaining a final to-be-estimated parameter correction through multiple iterations, and calibrating and compensating the measurement process of the deep space network tracking station by using the to-be-estimated parameter correction, thereby improving the satellite orbit determination precision.

The present invention discloses a GNSS partial ambiguity fixing method and system based on element-by-element ratio test. By constructing an element-by-element ratio test method, quality checks are carried out on the elements in the integer ambiguity alternative group one by one using the ratio test method, so as to judge the reliability of all ambiguity elements at one time, solve the key problem of quickly and effectively selecting a reliable ambiguity subset in high-dimensional ambiguity fixing, require lower computing power compared with the classical partial ambiguity fixing method, and ensure the stability of the method at the same time. The present invention solves the problem that high-dimensional ambiguities are difficult to be quickly and reliably fixed as a whole in the fields of existing GNSS high-precision positioning, satellite orbit determination, etc., has the characteristics of rigorous theory, simple model, easy implementation, etc., and can be applied to high-precision position service fields such as autonomous driving.

The application discloses a LS+CNN combined UT1-UTC prediction method considering effective angular momentum, relates to the field of geophysics and related computer science technology, and comprises the following steps: acquiring a UT1-UTC basic data set and an effective angular momentum data set, constructing an LS model and a CNN model, combining an LS fitting extrapolation result with an output result of the CNN model, and obtaining a high-precision UT1-UTC prediction value. The model has high precision and reliability in long-term UT1-UTC prediction, can overcome the shortcomings of a traditional linear model in UT1-UTC parameter prediction, and improves the prediction precision and stability of UT1-UTC compared with the traditional method. The method realizes high-precision long-term UT1-UTC prediction, and provides a new UT1-UTC parameter prediction method for the fields of navigation satellite orbit determination, astronomical observation and geophysical research.

The application discloses a satellite orbit determination method and device, a storage medium and an electronic device. The method comprises the following steps: acquiring an initial satellite position of a target communication satellite in the case that the target communication satellite can receive a navigation signal, and determining a reference orbit of the target communication satellite based on space-time information of a navigation satellite; obtaining the target satellite position and a target non-conservative force model coefficient through orbit fitting; then performing orbit integration to determine an integrated orbit; comparing the integrated orbit and the reference orbit to determine a position error sequence and process noise covariance information; and performing orbit filtering using intersatellite link data in the case that the target communication satellite cannot receive the navigation signal, so as to determine a target operation orbit. The application solves the technical problem that the satellite cannot operate according to a correct orbit due to overall offset rotation of a satellite constellation.

The invention provides an optical formation satellite orbit determination method and system, and the method comprises the steps: S1, building a kinetic model of a formation satellite system according to the stress condition of a satellite; the formation satellite system is used for formation satellites; s2, acquiring an inter-satellite distance and an inter-satellite angle based on the kinetic model; s3, establishing a state equation and an observation equation according to the inter-satellite distance and the inter-satellite angle; s4, carrying out nonlinear transformation on the random variable to obtain a nonlinear transformation result; carrying out statistical averaging on the nonlinear transformation result, and then filtering to obtain an unbiased result; and S5, constructing an iterative forecasting equation based on the unbiased result, the state equation and the observation equation, and further determining the orbit of the formation satellite. According to the method, distance measurement and angle measurement information is utilized, the orbit determination precision is improved, the principle is simple, programming calculation is convenient, the method is suitable for the field of deep space formation satellite orbit determination, and compared with traditional VLBI measurement, the orbit determination precision is improved.

The application relates to a low-orbit satellite orbit determination method, device and medium based on a multi-thread parallel processing architecture, and belongs to the technical field of satellite orbit determination. The method comprises the following steps: adopting a multi-thread parallel architecture to modularly decouple a satellite orbit determination business process, splitting the satellite orbit determination business process into three types of independent threads, i.e., a positioning thread, an orbit integration thread and an orbit prediction thread, and configuring the three types of threads in turn from high to low in priority: the positioning thread has the highest priority, the orbit prediction thread is in the middle, and the orbit integration thread has the lowest priority; meanwhile, relying on a hierarchical priority scheduling mechanism, GNSS observation data is preferentially guaranteed to be processed in real time in the positioning thread, and the latest effective orbit determination basic observation information is continuously output; the second priority is guaranteed to enable the orbit prediction thread to quickly respond to user requests and data retrieval requirements of an inter-satellite link; and finally, the kinetic integration with a large amount of calculation is processed, so that the real-time performance and business continuity of the orbit determination are effectively improved.

The application discloses a low-orbit satellite navigation enhancement and ground CORS cooperative system and relates to the technical field of satellite navigation positioning, which comprises a space segment, a ground segment and a user segment; the space segment comprises a low-orbit satellite and is used for broadcasting navigation enhancement signals; the ground segment comprises an enhanced CORS network and a star-ground cooperative control center; each ground reference station simultaneously receives GNSS signals and navigation enhancement signals and generates observation data; the star-ground cooperative control center performs precise orbit determination based on low-orbit satellite downlink data and observation data, and uploads star-ground calculation results to the low-orbit satellite to update the generation parameters of the navigation enhancement signals; the user segment comprises a hybrid enhancement receiver, which is used for receiving GNSS signals, navigation enhancement signals and differential data and performing multi-source fusion positioning calculation. The application can improve the low-orbit satellite orbit determination precision, shorten the user initialization time, expand the service coverage, reduce the system construction cost and enhance the integrity guarantee capability.

The present disclosure provides a method and system for autonomous orbit determination of a low-orbit navigation constellation under a weak inter-satellite topology structure, which relates to the field of satellite orbit determination technology, including: obtaining various types of observation data of an inter-satellite link on a predetermined orbit, and pre-processing the various types of observation data; adopting a piecewise first-order polynomial satellite clock error modeling method to establish synchronous observation equations of orbit and clock error parameters, generate orbit, clock error and related error parameters of joint on-board GNSS and inter-satellite link data, and predict them; based on the error analysis results, extract multiple key features from the generated orbit, clock error and related error parameters as inputs of a parameter prediction model, and output four types of predicted parameter time series: relative weight, inter-satellite link hardware delay, low-orbit satellite clock error and unmodeled error; uploading the predicted parameter time series and broadcast ephemeris together to a main satellite, and the main satellite transmits them to each sub-satellite through an inter-satellite link, so as to realize distributed autonomous orbit determination of the sub-satellite operation.

The application relates to a combined orbit determination method based on SLR and intersatellite links, which comprises the following steps: S100, determining a nonlinearized intersatellite observation equation according to intersatellite observation data; S200, determining a nonlinearized SLR observation equation according to SLR observation data; S300, determining a state transition matrix and a linearized satellite orbit determination observation equation according to a satellite motion equation and the nonlinearized satellite orbit determination observation equation; S400, determining a linearized SLR observation equation according to the nonlinearized SLR observation equation, the state transition matrix and the linearized satellite orbit determination observation equation; S500, determining a linearized intersatellite observation equation according to the nonlinearized intersatellite observation equation, the state transition matrix and the linearized satellite orbit determination observation equation; and S600, performing satellite combined orbit determination according to the linearized intersatellite observation equation and the linearized SLR observation equation. The application can simultaneously exert the technical advantages of SLR and intersatellite links, and can obtain a high-precision satellite orbit.

The invention discloses a navigation satellite orbit determination method and device based on satellite-ground combined observation, and the method comprises the steps: obtaining fixed star angular distance measurement data information of a navigation satellite in a preset semi-autonomous operation mode; performing inter-satellite observation by using a Ka inter-satellite link to obtain inter-satellite relative distance information; satellite-ground relative distance information is obtained by using satellite-ground observation of a Ka-band satellite-ground link; according to a preset semi-autonomous operation mode, obtaining satellite orbit parameter information of the semi-autonomous operation mode by using a preset navigation satellite orbit parameter starting value; and performing evaluation processing on the satellite orbit parameter information in the semi-autonomous operation mode to obtain an optimal satellite-ground joint support mode. Therefore, the invention provides a multi-mode optimization technical scheme based on satellite-ground joint support, and the calculation precision of the navigation satellite parameters is improved.

This application relates to a method for multi-dimensional feature extraction and intelligent identification of anomalies in low-Earth orbit (LEO) satellite orbit determination links. The method includes: extracting features from preprocessed signals and observation data from three dimensions—signal layer, data layer, and geometric layer—using a multi-dimensional feature extraction module to construct a seven-dimensional feature vector; building a differentiated fingerprint database representing different anomaly mechanisms based on the seven-dimensional feature vector; classifying and identifying anomaly types using a pre-trained lightweight hybrid intelligent classification model, and outputting the identified anomaly type and confidence level; and outputting the anomaly type and confidence level to the LEO satellite orbit determination system for adaptive fault-tolerant decision-making. This method overcomes the fundamental dependence of traditional precise orbit determination methods on a friendly, stable, and continuous observation environment, as well as the inherent defects of traditional robust filtering techniques, such as detection lag, unclear diagnosis, and limited strategy.

The present invention discloses a method and device for determining the orbit of a lunar-orbiting satellite based on radio and laser ranging, which relates to the technical field of lunar-orbiting satellite orbit calculation, and includes: acquiring the observation data of radio tracking stations and laser stations; establishing a high-precision dynamic model of the lunar-orbiting satellite; establishing a mathematical model for radio and laser ranging of the lunar-orbiting satellite, processing the observation data according to the observation equation, and obtaining the solution values of the dynamic parameters of the satellite at the initial orbit determination moment; substituting the solution values of the dynamic parameters into the satellite perturbed motion equation to calculate and obtain the orbit determination result of the lunar-orbiting satellite. The present invention uses combined orbit determination by radio ranging and laser ranging to achieve precise orbit determination of the lunar-orbiting satellite. Compared with the traditional orbit determination method, the accuracy and reliability are greatly improved, and the requirements for the number and geometric distribution of ground stations are reduced, thus reducing the measurement and control pressure on existing ground stations.