28 results about "Satellite orbit determination" patented technology

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 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 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 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 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 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.

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 application relates to a kind of navigation satellite orbit determination precision promotion method based on pseudo-range deviation correction, belong to satellite navigation field, the method is combined with global distribution ground station dual-frequency pseudo-range observation value and external precise orbit clock difference product constructs ionosphere-free combination observation equation, through precise point positioning calculation receiver coordinate and constructs posteriori pseudo-range residual equation;According to the grouping of receiver model and the application of zero-sum constraint, align receiver clock difference, through multi-epoch joint network least square batch processing each type receiver is solved to each navigation satellite dual-frequency ionosphere-free combination pseudo-range deviation value;According to the receiver model of ground station, select the pseudo-range deviation value of corresponding type, deduct from dual-frequency ionosphere-free combination pseudo-range observation, obtain the pseudo-range observation of correction;Using the pseudo-range observation of correction and carrier phase observation, joint solution navigation satellite precise orbit parameter.The present application has significantly improved the navigation satellite orbit determination precision and spatial reference consistency level.

The embodiment of the invention relates to the field of satellite orbit determination measurement, in particular to a multi-target polling type CEI orbit determination system which comprises a signal receiving system, a signal transmission system, a signal frequency conversion system, an acquisition related processing system and an orbit real-time resolving system. According to one specific embodiment of the system, a signal receiving system receives a satellite signal set; the signal transmission system obtains a satellite transmission signal set; the signal frequency conversion system obtains a satellite standard signal set; acquiring a satellite monitoring signal set obtained by a related processing system; and the orbit real-time resolving system obtains a satellite ephemeris information set. According to the implementation mode, the reliability of the multi-target polling type CEI orbit measurement system is improved.

The application discloses a multi-satellite real-time orbit determination method in a one-launch-multiple-satellite mode, which comprises the following steps: building a one-launch-multiple-satellite real-time orbit determination system; dividing the multi-satellite orbit determination into different calculation work groups according to the satellite-launcher separation sequence, each calculation work group corresponding to one satellite-launcher separation; distributing each satellite to the corresponding calculation work group according to the satellite-launcher separation process; each satellite corresponding to an independent orbit determination program; the single-satellite orbit determination program in the calculation work group only collects orbit determination data without running the orbit determination calculation process when no satellite-launcher separation signal of the satellite is received; and when the satellite-launcher separation signal of the satellite is received, the orbit determination calculation process is started to obtain the initial orbit of the satellite.

The invention provides a satellite inter-satellite link distance measurement deviation calibration method and system. The method comprises the following steps: collecting and preprocessing Ka wave band inter-satellite link observation data of a target time period; epoch normalization is carried out, and inter-satellite link observation values at different moments are reduced to the same moment; combining observation values, and respectively constructing a clock-error-free observation quantity and a geometric-error-free observation quantity; an inter-satellite link observation equation is constructed, satellite orbits, clock errors and hardware time delay parameters are estimated, and the time delay parameters are divided into two modeling modes according to satellites and according to links; carrying out convergence judgment and iterative calculation according to a post-test residual statistical index; and finally, the distance measurement deviation of each one-way link is extracted through the difference between the satellite-related time delay parameter and the link-related time delay parameter. According to the method, the internal correlation between the time delay characteristics of different links and the distance measurement deviation is fully considered, the actual situation of a Beidou system is better met, the inter-satellite link orbit determination residual error can be remarkably reduced, the satellite orbit determination precision is improved, and the method has high stability and reliability and good actual application value.

The invention provides a satellite precise orbit determination and weight determination method based on observation value post-test residual errors, and relates to the technical field of satellite orbit determination. According to the method, the observed value post-test residual error is obtained based on equal-weight weighting, the observed value precision sky map is constructed through the observed value post-test residual error, observation value weighting is carried out according to the observed value precision sky map, parameter estimation and orbit integration are iteratively carried out to solve the satellite orbit, and external coincidence precision evaluation is carried out. According to the method, the noise characteristics of the satellite observation value can be effectively described, the random model is obviously optimized, and the satellite orbit with higher precision is finally determined.

The invention provides a low earth orbit satellite orbit determination method, orbit determination equipment, a readable storage medium and a product, and relates to the technical field of communication, the method comprises the following steps: collecting multi-source data, the multi-source data comprising low earth orbit satellite self-loading data, real-time observation data collected by a ground continuous operation reference station (CORS) and environmental data collected by an auxiliary sensor; carrying out multi-source data fusion on the low-orbit satellite self-loading data, the real-time observation data and the environment data to obtain comprehensive positioning data; inputting the comprehensive positioning data into a low-orbit satellite state space model containing position, speed and acceleration, and performing state estimation on the comprehensive positioning data through the low-orbit satellite state space model to obtain orbit state parameters of the low-orbit satellite, the orbit state parameters comprise at least one of the position, the speed and the acceleration of the low-orbit satellite; according to the method, the orbit determination precision of the low-orbit satellite can be improved, errors are reduced, and the robustness and real-time performance of the system are improved.

The invention provides a double-satellite formation satellite orbit determination method and device based on inter-satellite relative optical imaging and microwave measurement, and the method comprises the steps: building a satellite motion equation based on orbit dynamics, and obtaining a state equation; establishing an observation equation based on inter-satellite relative optical imaging and microwave measurement; measuring the accurate orbit of the reference satellite A; the reference star A performs optical imaging on the target star B to obtain azimuth information of the star B relative to the star A; carrying out microwave distance measurement on the reference satellite A and the target satellite B to obtain relative distance information; calculating orbit parameters of the satellite B, and obtaining position coordinates and a satellite orbit of the satellite B according to the observation coordinate value of the satellite B relative to the satellite A and the state equation; and obtaining an orbit determination result of the target satellite at any moment. The economic benefit of satellite in-orbit operation can be improved.

The invention discloses a satellite orbit determination method and a ground operation control center, and the method is applied to the ground operation control center, and comprises the steps: obtaining a first state parameter of a first satellite at a first moment; observation data received by the first satellite from the target satellite at the target moment are obtained, first observation data are obtained, the operation orbit of the target satellite is higher than that of the first satellite, and the first moment is the moment before the target moment; determining a second state parameter of the first satellite at the target moment based on the first state parameter; and determining a target state parameter of the target satellite at the target moment based on the second state parameter and the first observation data. Through the method and the device, the problem that the maneuvering satellite cannot accurately determine the orbit in the related technology is solved, and the effect of improving the satellite orbit determination accuracy is further achieved.

The invention discloses a TLE forecast precision evaluation method for star chain constellation satellites, and relates to the technical field of satellite orbit determination, and the method comprises the steps: obtaining ephemeris data and TLE data of star chain satellites; determining a position variance according to the TLE data, and dividing the ephemeris data based on the position variance to obtain a critical epoch moment and multiple segments of ephemeris data; performing data forecasting processing of a preset epoch moment sequence on the TLE data to obtain a TLE forecasting result; dividing the TLE forecast result based on the critical epoch moment to obtain multiple segments of TLE forecast data; determining the position deviation of the two segments according to the multiple segments of ephemeris data and the multiple segments of TLE forecast data; and determining a TLE forecast precision evaluation result of the star chain satellite based on the two sections of position deviation. According to the method, the evaluation time coverage range and the result fineness can be remarkably improved, and more accurate orbit forecast error quantification support is provided for ground tracking and monitoring.

The present application relates to a kind of low-orbit satellite orbit determination method and system based on BDS-3 navigation message SISRE precision promotion.Method includes: according to the broadcast of BDS-3 navigation message broadcast parameter, the broadcast clock difference of current time in 1 hour is calculated;Based on historical 24 hours broadcast clock difference data, using fitting method to correct current time clock speed;Based on the corrected current time clock speed, the initial clock difference of next time is calculated;Clock difference jump is detected, when clock difference jump occurs, 24 hours of BDS-3 navigation message broadcast parameter is accumulated again, repeat the above steps;And using the corrected broadcast clock difference parameter, combined with low-orbit satellite on-board observation data completes autonomous orbit determination.The method proposed in the present application, calculation step is simple, can greatly reduce SISRE error, so that on-board autonomous orbit determination has high precision and timeliness, meet the requirements of occultation detection and satellite measurement and other important scientific tasks.

This invention discloses a weighting method for adjustment observations in a multi-system navigation satellite joint orbit determination network, belonging to the field of satellite navigation and precise orbit determination data processing technology. The method includes: S1 observation data preprocessing and quality control; S2 orbit numerical integration and fitting; S3 prior weighted parameter estimation and residual editing; S4 grouping observations by station and navigation system dimensions, and initializing variance factors with equal weights; S5 estimating the variance scaling factor for each group using a simplified Helmert formula omitting the redundancy trace correction term and updating the weight matrix, combined with robust estimation to reduce the weight of outlier residuals, iterating until convergence; S6 re-estimating ambiguity with fixed constraints, outputting precise orbit and clock error products. This invention uses the above method, replacing fixed prior weights with data-driven simplified variance component estimation, to characterize the non-homogeneous characteristics of observations with almost no increase in computational burden, thus improving the accuracy and reliability of multi-system joint orbit determination.

The present application relates to satellite orbit determination technical field, specifically to low-orbit giant constellation autonomous orbit determination method based on star sensor and inter-satellite angle measurement. The method comprises: step 1, determining constellation configuration by using inter-satellite measurement: using six number expression each satellite state quantity in constellation, constructing partial derivative matrix of inter-satellite measurement relative to state information, constructing linear equation set according to state quantity, measurement and partial derivative matrix, solving satellite state quantity describing constellation configuration by improved least square method. Step 2, using star sensor and camera measurement, combining constellation configuration in step 1, correcting constellation drift: using star sensor and camera, obtaining the real value of the angle between the direction vector of adjacent satellite and the direction vector of a certain star, then calculating the estimated value of the angle between the direction vector of adjacent satellite and the direction vector of a certain star, then minimizing the difference between the real value and the estimated value of the angle, realizing constellation correction. The above scheme can realize orbit determination without relying on ground station and is easy to manage data.

The application discloses a Beidou-3 satellite orbit determination and time synchronization method and system. The method comprises the following steps: setting clock difference sampling points at certain intervals in an orbit determination arc segment; in the orbit determination arc segment, orbit determination and time synchronization calculation are performed by using observation values to obtain calculation parameters; theoretical observation values of the observation values are calculated according to the calculation parameters; and residual errors are calculated according to the observation values and the theoretical observation values, and observation values with residual errors exceeding a threshold value are removed. The application directly uses original non-synchronized observation single-pass inter-satellite link data to realize orbit determination and time synchronization, and simultaneously calculates the clock speed parameter of each satellite and the satellite orbit, clock difference, clock speed and other parameters, thereby avoiding the dependence on external satellite clock speed data (such as clock difference rate data in a navigation message), being more rigorous in theory, improving the orbit determination and time synchronization precision, and only adding one clock speed parameter for each satellite, thereby having little influence on the orbit determination and time synchronization calculation efficiency.

The application discloses a low-orbit satellite orbit determination precision improvement method and device under geomagnetic storm conditions and belongs to the technical field of satellite navigation. The method proposes an adaptive orbit determination framework of dynamic model and stochastic model optimization. At the dynamic level, the segmented estimation length of empirical force and atmospheric resistance coefficient is dynamically adjusted according to the geomagnetic activity level, and the resolution capability for rapid changes of non-conservative force is improved. At the stochastic model level, an external physical driving weight function based on the SYM-H index and an internal data diagnosis weight function based on observation residuals are constructed, and real-time adaptive adjustment of observation value variance is realized. Through the "geomagnetic index-observation residual" double driving mechanism, the low-orbit satellite orbit determination error during the strong magnetic storm is reduced from the decimeter level to the centimeter level, and the precision and robustness of the low-orbit satellite orbit determination under the space weather event are significantly improved.

The invention provides a satellite orbit measurement data simulation method and system. The method comprises the following steps: selecting observation station input required by data simulation according to the distribution of a satellite measurement and control network; target input required by data simulation is selected according to space target data, a first space position reference system of an observation station and a target is converted to a J2000 coordinate system for simulation operation, actual measurement scene simulation is carried out according to visibility influence factors in actual observation, and sparse observation is formed; a dynamic model of space target operation is constructed, a time position sequence is formed, various types of orbit measurement data are solved, and random noise is added; on the basis of orbit measurement data with random noise, simulation data closer to actually measured data attributes are formed. According to the method, various visibility criteria are provided, and certain guidance is provided for spaceflight measurement and control tasks; meanwhile, a large amount of observation data can be generated by the mass satellite orbit measurement data simulation method, and support can be provided for capability demonstration of a cataloguing system.

The invention relates to a navigation satellite orbit determination precision improving method based on pseudo-range deviation correction, which belongs to the field of satellite navigation, and comprises the following steps: constructing an ionosphere-free combined observation equation by combining a global distributed ground observation station double-frequency pseudo-range observation value and an external precision orbit clock difference product; calculating the coordinates of the receiver through precise single-point positioning, and constructing a post-test pseudo-range residual equation; grouping is carried out according to the models of the receivers, zero sum constraint is applied, clock errors of the receivers are aligned, and the dual-frequency ionospheric-free combination pseudo-range deviation values of the receivers of all the models to all the navigation satellites are calculated through multi-epoch combined whole network least square batch processing; according to the model of a ground station receiver, selecting a pseudo-range deviation value of a corresponding model, and deducting from the dual-frequency ionosphere-free combination pseudo-range observed quantity to obtain a corrected pseudo-range observed quantity; and utilizing the corrected pseudo-range observed quantity and the carrier phase observed quantity to jointly solve the navigation satellite precision orbit parameters. According to the method, the consistency level of the navigation satellite orbit determination precision and the space reference is remarkably improved.

The invention discloses a satellite atmospheric resistance modeling method based on a geometric effective area and a relative flow direction. The method belongs to the technical field of aerospace dynamics and satellite orbit determination. The defects that only resistance coefficient modeling is carried out in an existing satellite atmospheric resistance model, and the reference area is fixed to be a constant are overcome. Comprising the following steps: S1, defining a relative speed and a coordinate system; s2, effective windward area calculation based on a geometric model: constructing a satellite geometric model, and calculating a geometric effective windward area of a satellite; s3, constructing a resistance coefficient angle model: constructing the resistance coefficient angle model through the azimuth angle and the pitch angle of the relative flow direction under the body-fixed coordinate system; s4, defining an effective windward area-coefficient combination quantity: combining the resistance coefficient angle model with the geometric effective windward area of the satellite, and defining the effective windward area-coefficient combination quantity; and S5, constructing an improved atmospheric resistance acceleration model: constructing the improved atmospheric resistance acceleration model through the effective windward area-coefficient combination quantity.

The application relates to a high-precision ranging integrated system and a ranging method, equipment and medium, and belongs to the technical field of satellite orbit determination. The system comprises a transmitting module, which is used for transmitting a downlink spread spectrum signal after frequency up-conversion, radio frequency filtering and power amplification, transmitting one way and sending the other way to a local oscillator and self-checking module; a receiving module, which is used for filtering, amplifying and twice down-converting an uplink radio frequency signal, and outputting an uplink intermediate frequency signal to the local oscillator and self-checking module; a clock module, which is used for keeping the system clock synchronous and consistent; a local oscillator and self-checking module, which is used for providing a local oscillator frequency, converting the downlink spread spectrum signal processed by the transmitting module into a calibration intermediate frequency signal, and independently outputting the uplink intermediate frequency signal and the calibration intermediate frequency signal to a digital baseband module; and the digital baseband module, which is used for baseband signal processing, calibration of an uplink signal path time delay and real-time correction of a carrier phase measurement value. The system can meet the high-precision ranging requirement of high-orbit satellite orbit determination.