A Method, Device, Equipment and Medium for Determining the Tracking and Control Orbit of Low-Earth Orbit Satellites

Abstract

The present invention discloses a method, device, equipment and medium for determining the tracking and control orbit of a low-earth orbit satellite. The method includes: obtaining on-board GNSS pseudorange observations and GNSS broadcast ephemeris, and calculating the self-positioning result of the low-earth orbit satellite through the pseudorange single-point positioning algorithm; when the number of GNSS satellites used in the pseudorange single-point positioning is lower than a first threshold, or the satellite spatial distribution eigenvalue of the positioning epoch is greater than a second threshold, eliminating the corresponding self-positioning result; using a sliding window to perform polynomial fitting on the filtered self-positioning results, eliminating the self-positioning results with fitting residuals greater than a third threshold, and then performing interpolation processing to obtain an optimized self-positioning result; combining the dynamic fitting method for orbit determination to obtain the final tracking and control orbit of the low-earth orbit satellite. The present invention has a low calculation cost and good real-time performance, gets rid of the dependence on ground stations, is less affected by the orbit maneuver of low-earth orbit satellites, and can reflect the relatively accurate satellite position.

Description

Technical Field

[0001] The present invention relates to the technical field of satellite orbit measurement and control, and particularly to a method, device, equipment and medium for determining the measurement and control orbit of a low-orbit satellite. Background Art

[0002] The orbit measurement and control of low-orbit satellites is crucial for ensuring the normal operation and mission execution of satellites. Accurate orbit determination can provide precise position and velocity information of the satellite, thus supporting effective attitude control, orbit correction, collision avoidance and other operations. In addition, orbit measurement and control can also ensure smooth communication between the satellite and the ground station, and help the smooth progress of tasks such as scientific research, remote sensing observation and navigation positioning.

[0003] The orbit of a low-orbit satellite is affected by various disturbing factors such as atmospheric drag and solar radiation pressure. Traditional orbit measurement and control methods such as ground station observation or on-board GNSS (Global Navigation Satellite System) observation usually have difficulty providing high-precision orbit states in a short time. On-board self-positioning data can provide instant orbit information, which does not depend on real-time observation data of ground measurement and control stations, and reduces the influence of data transmission delay. Through on-board reception and autonomous navigation and positioning algorithms, the satellite can complete orbit determination without ground control, which can effectively alleviate the deficiency of relying on ground stations. It helps to reduce operation and maintenance costs and improve the efficiency of orbit determination. Especially for high-frequency low-orbit satellite monitoring tasks, it can achieve a lower-cost orbit measurement and control solution. Up to now, there has been little research on the measurement and control orbit determination of low-orbit satellites based on on-board self-positioning data. In the prior art, the following several technical methods are relatively relevant to this method:

[0004] The patent application with the publication number of CN117008155A discloses a method for improving the orbit accuracy of on-orbit single-point positioning of a low-orbit satellite. According to the position-velocity time series of the low-orbit satellite transmitted back during a period, six Keplerian elements of the low-orbit satellite at the initial moment are calculated, and two self-defined coefficients and pseudo-random pulse parameters within N time periods are fitted. On this basis, the acceleration of the low-orbit satellite is integrated using the Collocation multi-step method, and the dynamically smoothed orbit of the low-orbit satellite at any moment within the integration interval is obtained using the integration data, which can achieve precise orbit determination of low-orbit satellites at low cost. However, the implementation process of this scheme is relatively complex and requires segmentation of the entire orbit determination arc. Moreover, two self-defined coefficients need to be obtained by least squares fitting first, and the calculation process is complex and time-consuming. The two self-defined coefficients respectively represent the tangential and normal accelerations of the solar radiation pressure, while the influence of the solar radiation pressure perturbation on the low-orbit satellite is not significant. Atmospheric drag is the largest error source of the perturbation force received by the low-orbit satellite, and this scheme does not consider the improvement of atmospheric drag, so there may be large orbit errors.

[0005] The patent application with the publication number CN113008243A discloses a method and system for autonomous orbit determination of a low-Earth orbit satellite constellation, including obtaining inter-satellite link ranging data and on-board GNSS receiver observation data. First, the PPP solution is used to obtain the approximate position of the low-Earth orbit satellite based on the on-board observation data. Secondly, the orbit of the low-Earth orbit satellite is fitted based on dynamics. Finally, the joint adjustment of the low-Earth orbit constellation network is carried out through the inter-satellite link data to achieve high-precision safe autonomous orbit determination of the low-Earth orbit satellite on the satellite. However, this method includes multiple key steps, and the required data types and data volumes are relatively large. It is necessary to use both the precise point positioning method and the dynamic orbit determination method. The implementation process is cumbersome and the computational amount is large.

[0006] The patent application with the publication number CN110058287A discloses a method, device and system for orbit determination of a low-Earth orbit satellite. The orbit and clock offset of the navigation satellite are calculated based on the broadcast ephemeris, and the correction information of the positioning information is introduced from the ground system. The low-Earth orbit satellite is orbited using the accurate corrected positioning information, thereby improving the orbit determination accuracy of the low-Earth orbit satellite. However, it is not easy to obtain the correction information of the navigation satellite through the ground system, and the accuracy of the navigation satellite orbit calculated by combining the broadcast ephemeris and the correction information cannot be guaranteed, resulting in limited orbit determination accuracy of the low-Earth orbit satellite. This method depends on both the navigation satellite and the ground station at the same time, and the autonomy is relatively poor.

[0007] Therefore, the existing technologies mainly have the following technical problems:

[0008] (1) Traditional orbit measurement and control methods such as ground station observation or on-board GNSS observation usually have difficulty providing high-precision orbit states in a short time; generally, a relatively precise dynamic model is required to realize the orbit determination of a low-Earth orbit satellite through ground station observation means and on-board GNSS observation means. In addition, the dynamic orbit determination method requires ensuring real-time observation data transmission and strict data processing algorithms of the ground station. The implementation process is computationally intensive and relatively complex;

[0009] (2) During the orbit control process of a low-Earth orbit satellite, the attitude and orbit of the satellite will change frequently, which will lead to a decrease in the timeliness and accuracy of on-board observation or ground observation data; the dynamic changes of the satellite may cause deviations in the observation data, thereby affecting the orbit determination accuracy; using the orbit control acceleration modeling will inevitably increase the computational complexity and calculation time consumption. How to realize the orbit determination of a low-Earth orbit satellite during the orbit control period has become a difficult problem;

[0010] (3) Since the satellite side cannot implement overly complex algorithms, the on-board self-positioning results are generally obtained by solving a simple single-point positioning algorithm. Due to the high-dynamic motion of low-earth orbit satellites and the quality problems of on-board observation data, there may be outliers in the on-board self-positioning data, and the outliers deviate significantly from the actual orbit of low-earth orbit satellites. In addition, the on-board self-positioning data of low-earth orbit satellites are discrete points with unequal intervals, which do not conform to the smooth characteristics of low-earth orbit satellite orbits, and directly using them for orbit determination may lead to divergence of orbit fitting. Summary of the Invention

[0011] Aiming at the deficiencies in the prior art, the purpose of the present invention is to provide a method, device, equipment and medium for determining the tracking and control orbit of a low-earth orbit satellite to solve at least one of the above technical problems.

[0012] In the first aspect, a method for determining the tracking and control orbit of a low-earth orbit satellite is provided, including the following steps:

[0013] Obtain self-positioning results: Obtain on-board GNSS pseudorange observations and GNSS broadcast ephemeris, and solve the self-positioning results of the low-earth orbit satellite through the pseudorange single-point positioning algorithm;

[0014] Filter self-positioning results: When the number of GNSS satellites used in the pseudorange single-point positioning is lower than the first threshold, the corresponding self-positioning results are excluded; calculate the satellite spatial distribution eigenvalue of the positioning epoch, and when the satellite spatial distribution eigenvalue is greater than the second threshold, the corresponding self-positioning results are excluded;

[0015] Optimize self-positioning results: Use the sliding window method to perform polynomial fitting on the filtered self-positioning results, exclude the self-positioning results with fitting residuals greater than the third threshold, and perform interpolation processing according to the finally fitted polynomial to obtain the optimized self-positioning results;

[0016] Satellite orbit determination: Based on the optimized self-positioning results, use the dynamic fitting method to determine the orbit of the low-earth orbit satellite to obtain the final tracking and control orbit of the low-earth orbit satellite.

[0017] Further, the process of solving the self-positioning results of the low-earth orbit satellite through the pseudorange single-point positioning algorithm specifically includes:

[0018] On-board GNSS pseudorange observations The observation equation is expressed as follows:

[0019] ;

[0020] In the formula, the subscripts , respectively represent the frequency point and the on-board receiver, and the superscript represents the GNSS satellite, represents the geometric distance from the GNSS satellite to the center of the on-board receiver; and respectively represent the receiver clock error and satellite clock error included in the pseudorange; and respectively represent the hardware delay biases at the receiver end and satellite end of the frequency point , where the hardware delay at the receiver end is received by the receiver clock error during positioning; represents the tropospheric delay error of the pseudorange measurement, represents the ionospheric delay of the frequency point , is the pseudorange observation noise, is the speed of light in vacuum;

[0021] The error equation after linearization of the observation equation of the spaceborne GNSS pseudorange observation value is expressed as follows:

[0022] ;

[0023] In the formula, is the observation value correction vector, is the design matrix, is the pseudorange observation value residual vector, is the unknown parameter correction vector, , where , , are the spaceborne receiver coordinate corrections;

[0024] The unknown parameter correction vector can be obtained by the least square principle:

[0025] ;

[0026] In the formula, is the coefficient matrix of the normal equation, is the observation value weight matrix;

[0027] Iterative calculation is carried out until the unknown parameter correction vector satisfies the iterative convergence condition, and then the spaceborne receiver coordinates at the current epoch are obtained as the self-positioning result of the low-earth orbit satellite.

[0028] Furthermore, the satellite spatial distribution eigenvalue is calculated by the following method:

[0029] The geometric distance from the GNSS satellite to the center of the spaceborne receiver is expanded by Taylor series at the approximate position of the low-earth orbit satellite, and the first-order term is retained. The linearized equation is:

[0030] ;

[0031] In the formula, , , , is the geometric distance from the GNSS satellite to the center of the onboard receiver calculated from the approximate position, Indicates GNSS satellite coordinates;

[0032] make ,have , subscript n is the number of GNSS satellites observed by the onboard receiver in the current epoch;

[0033] Then the satellite spatial distribution characteristic value is .

[0034] Furthermore, the self-positioning result optimization specifically includes:

[0035] Set the sliding window length and sliding step size. The sliding window length is equal to the sliding step size.

[0036] Divide the filtered self-positioning results into multiple segments of self-positioning results through a sliding window;

[0037] For each segment of the self-positioning results after division, a polynomial fitting is performed, and the self-positioning results whose fitting residuals are greater than the third threshold are eliminated. The polynomial fitting and elimination process is iterated until the fitting residuals of the retained self-positioning results are all less than the third threshold, thereby obtaining the final fitting polynomial;

[0038] Interpolation processing is performed according to the final fitted polynomial to obtain the optimized self-positioning result.

[0039] Furthermore, the satellite orbit determination specifically includes:

[0040] Assumptions The self-positioning position on the satellite at time ;

[0041] The position and velocity at the initial orbit time can be obtained by fitting the final fitting polynomial, and the position and velocity at the initial orbit time can be obtained by integration. The low-orbit satellite position at the time , and then establish the N-dimensional observation equation as follows:

[0042] = ;

[0043] In the formula, for The satellite motion state transfer matrix at time , , , are the corrections of the position, velocity and dynamic parameters of the low-orbit satellite at the time of initial orbit;

[0044] The least squares optimal solution is obtained for the N-dimensional observation equation, and iterative calculations are performed until , , the iterative convergence condition is satisfied, and then the final orbit of the low-Earth orbit satellite is obtained through orbit integration.

[0045] Furthermore, the first threshold is 8; the second threshold is 2; the third threshold is 10 m.

[0046] In a second aspect, a device for determining the measurement and control orbit of a low-Earth orbit satellite is provided, including:

[0047] An initial self-positioning module, configured to obtain on-board GNSS pseudorange observations and GNSS broadcast ephemeris, and calculate the self-positioning result of the low-Earth orbit satellite through the pseudorange single-point positioning algorithm;

[0048] A filtering module, configured to discard the corresponding self-positioning result when the number of GNSS satellites used for pseudorange single-point positioning is lower than the first threshold; calculate the satellite spatial distribution eigenvalue of the positioning epoch, and discard the corresponding self-positioning result when the satellite spatial distribution eigenvalue is greater than the second threshold;

[0049] An optimization module, configured to perform polynomial fitting on the filtered self-positioning results in a sliding window manner, discard the self-positioning results with fitting residuals greater than the third threshold, and perform interpolation processing according to the finally fitted polynomial to obtain the optimized self-positioning result;

[0050] An orbit determination module, configured to perform orbit determination on the low-Earth orbit satellite based on the optimized self-positioning result using the dynamic fitting method to obtain the final measurement and control orbit of the low-Earth orbit satellite.

[0051] In a third aspect, a device for determining the measurement and control orbit of a low-Earth orbit satellite, configured on the low-Earth orbit satellite, includes:

[0052] An on-board receiver, configured to receive on-board GNSS pseudorange observations and GNSS broadcast ephemeris, and transmit them to the low-Earth orbit satellite orbit determination module;

[0053] A satellite-side calculation module, configured to implement the method for determining the measurement and control orbit of a low-Earth orbit satellite as described above.

[0054] In a fourth aspect, an electronic device is provided, including:

[0055] A memory, on which a computer program or instruction is stored;

[0056] A processor, configured to load and execute the computer program or instruction to implement the method for determining the measurement and control orbit of a low-Earth orbit satellite as described above.

[0057] Fifth aspect, there is provided a computer-readable storage medium storing a computer program or instructions, which when executed by a processor implement the low-Earth orbit satellite TT&C orbit determination method as described above.

[0058] The present invention provides a low-Earth orbit satellite TT&C orbit determination method, apparatus, device and medium, having the following beneficial effects:

[0059] (1) For the problem of large errors in ground-based observation data during orbit control, it is proposed to use GNSS observation data and GNSS broadcast ephemeris received by an on-board receiver for pseudorange single-point positioning to real-time calculate the preliminary self-positioning of the low-Earth orbit satellite, which can achieve relatively accurate orbit determination during orbit control;

[0060] (2) For the problem that there are some outliers in the on-board self-positioning results, it is proposed to use the satellite spatial distribution eigenvalue and the number of satellites used in the self-positioning calculation to eliminate the outliers in the self-positioning results, improving the quality of the original self-positioning data from the positioning principle;

[0061] (3) The short-term low-Earth orbit satellite orbit conforms to polynomial characteristics. In order to better use the discrete on-board self-positioning results for dynamic orbit determination, it is proposed to use the polynomial fitting method with a sliding window to fit and smooth the self-positioning data, and further eliminate the outliers in the self-positioning results through the fitting residuals. Using the smoothed self-positioning data can effectively improve the orbit determination accuracy;

[0062] (4) The implementation process of the present invention is relatively simple. Using the pseudorange single-point positioning algorithm reduces the computational complexity, and only using pseudorange measurement data also reduces the data processing complexity and computational amount. Therefore, the overall computational cost is low and it has good real-time performance, and it can achieve the low-Earth orbit satellite TT&C orbit determination without relying on a ground station; since the self-positioning results are calculated in real time, they are less affected by the orbit maneuver of the low-Earth orbit satellite and can reflect the relatively accurate satellite position; for the large-scale low-Earth orbit constellation, the massive low-Earth orbit satellite TT&C tasks are heavy and complex, and the low-Earth orbit satellites have frequent orbit maneuvers. The present invention provides an idea for the TT&C orbit determination problem of low-Earth orbit satellites in this situation. Description of the Drawings

[0063] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for the description of the embodiments or the prior art. Obviously, the following drawings are only some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0064] Figure 1 It is the flowchart of the low-Earth orbit satellite TT&C orbit determination method provided by the embodiment of the present invention. Detailed Embodiments

[0065] To make the objectives, technical solutions, and advantages of the present invention clearer, the technical solutions of the present invention will be described in detail below. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all of the embodiments. All other implementation manners obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts fall within the scope protected by the present invention.

[0066] In the description of the present invention, it should be understood that terms such as "first" and "second" are only used for descriptive purposes and cannot be construed as indicating or implying relative importance or order. In addition, in the description of the present invention, unless otherwise specified, the meaning of "a plurality of" refers to at least two.

[0067] As Figure 1 shown, an embodiment of the present invention provides a method for determining the tracking and control orbit of a low-earth orbit satellite, including the following steps:

[0068] S1: Obtain the self-positioning result: Obtain the on-board GNSS pseudorange observation value and the GNSS broadcast ephemeris, and calculate the self-positioning result of the low-earth orbit satellite through the pseudorange single-point positioning algorithm.

[0069] Specifically, the pseudorange single-point positioning algorithm includes the following process:

[0070] On-board GNSS pseudorange observation value The observation equation is expressed as follows:

[0071] ;

[0072] In the formula, the subscripts , respectively represent the frequency point and the on-board receiver, and the superscript represents the GNSS satellite, represents the geometric distance from the GNSS satellite to the center of the on-board receiver; and respectively represent the receiver clock error and the satellite clock error included in the pseudorange; and respectively represent the hardware delay deviation at the receiver end and the satellite end of the frequency point , where the hardware delay at the receiver end is received by the receiver clock error during positioning; represents the tropospheric delay error of the pseudorange measurement, represents the ionospheric delay of the frequency point , is the pseudorange observation noise, is the speed of light in vacuum;

[0073] Geometric distance from GNSS satellite to center of on-board receiver , where is the GNSS satellite coordinate, and is the center coordinate of the on-board receiver; the center coordinate of the on-board receiver is the quantity to be solved. During the initial calculation, the center coordinate of the on-board receiver is approximated by the approximate position and then, by continuously optimizing the coordinate correction of the on-board receiver, the final center coordinate position of the on-board receiver is obtained as the self-positioning result of the low-earth orbit satellite; the process of optimizing the coordinate correction of the on-board receiver is as follows:

[0074] The error equation after linearizing the observation equation of the on-board GNSS pseudorange observation is expressed as follows:

[0075] ;

[0076] In the formula, is the vector of observation value corrections; is the design matrix; is the vector of pseudorange observation residuals, which is the residual vector after subtracting the geometric distance from the pseudorange observation and various errors; is the vector of unknown parameter corrections, , where , , are the coordinate corrections of the on-board receiver;

[0077] The vector of unknown parameter corrections can be obtained by the least squares principle:

[0078] ;

[0079] In the formula, is the coefficient matrix of the normal equation, is the weight matrix of the observations;

[0080] Iterative calculations are carried out until the vector of unknown parameter corrections satisfies the iterative convergence condition, and then the coordinate of the on-board receiver at the current epoch is obtained as the self-positioning result of the low-earth orbit satellite.

[0081] Regarding the problem of large errors in the ground-based observation data during orbit control, it is proposed to use the GNSS observation data received by the on-board receiver and the GNSS broadcast ephemeris to perform pseudorange single-point positioning and real-time solve the preliminary self-positioning of the low-earth orbit satellite, which can achieve relatively accurate orbit determination during orbit control.

[0082] S2: Self-positioning result filtering: When the number of GNSS satellites used for pseudorange single-point positioning is lower than the first threshold, the corresponding self-positioning result is excluded; calculate the eigenvalue of the satellite spatial distribution at the positioning epoch. When the eigenvalue of the satellite spatial distribution is greater than the second threshold, the corresponding self-positioning result is excluded.

[0083] Through step S1, the central coordinates of the on-board receiver at multiple epoch moments (i.e., the self-positioning result of the low-earth orbit satellite) can be obtained. However, due to the high-dynamic operation of the low-earth orbit satellite, the number of GNSS satellites observed by the on-board receiver is too small or the spatial distribution configuration is poor in some time periods, which will lead to a large deviation in the central coordinates of the on-board receiver obtained in step S1. To address this problem, in this embodiment, two strategies are adopted to filter the self-positioning results of the low-earth orbit satellite over a period of time: The first strategy is to eliminate the self-positioning result when the number of GNSS satellites used in pseudorange single-point positioning is lower than the first threshold; the second strategy is to calculate the satellite spatial distribution eigenvalue of the positioning epoch, and eliminate the self-positioning result when the satellite spatial distribution eigenvalue is greater than the second threshold. The above first threshold and second threshold are set according to empirical values. In this embodiment, the first threshold is 8; the second threshold is 2 to ensure the accuracy of the self-positioning results of the high-dynamic low-earth orbit satellite.

[0084] The smaller the satellite spatial distribution eigenvalue, the higher the on-board self-positioning accuracy of the low-earth orbit satellite, and it is calculated by the following method:

[0085] The geometric distance from the GNSS satellite to the center of the on-board receiver at the approximate position of the low-earth orbit satellite is expanded by Taylor series, and the first-order term is retained. The linearized equation is:

[0086] ;

[0087] In the formula, , , , is the geometric distance from the GNSS satellite to the center of the on-board receiver calculated from the approximate position, represents the th GNSS satellite coordinates;

[0088] Let , there is , and the subscript n is the number of GNSS satellites observed by the on-board receiver at the current epoch;

[0089] Then there is the satellite spatial distribution eigenvalue .

[0090] Regarding the problem that there are some outliers in the on-board self-positioning results, it is proposed to eliminate the outliers of the self-positioning results by using the satellite spatial distribution eigenvalue and the number of GNSS satellites used in the self-positioning calculation, and improve the quality of the original self-positioning data from the positioning principle.

[0091] S3: Self - positioning result optimization: The filtered self - positioning results are subjected to polynomial fitting using a sliding window method. Self - positioning results with fitting residuals greater than the third threshold are removed. Interpolation is performed based on the finally fitted polynomial to obtain optimized self - positioning results.

[0092] Specifically, the self - positioning result optimization includes:

[0093] Set the sliding window length and the sliding step size. The sliding window length is equal to the sliding step size. In this embodiment, both of these parameters are set to 600 s;

[0094] The filtered self - positioning results are divided into multiple segments of self - positioning results through the sliding window;

[0095] For each segment of the divided self - positioning results, polynomial fitting is performed. Preferably, 5 - th order polynomial fitting is performed, and self - positioning results with fitting residuals greater than the third threshold (such as 10 m, which can be set according to specific accuracy requirements) are removed. The polynomial fitting and removal process is iterated until the fitting residuals of the remaining self - positioning results are not greater than the third threshold, obtaining the finally fitted polynomial;

[0096] Interpolation is performed based on the finally fitted polynomial. In this embodiment, uniform interpolation is performed to a series of self - positioning results with a time interval of 30 s to obtain optimized self - positioning results.

[0097] The short - term low - Earth orbit satellite orbit conforms to polynomial characteristics. In order to better use the discrete on - satellite self - positioning results for dynamic orbit determination, a polynomial fitting method using a sliding window is proposed to fit and smooth the self - positioning data, and outliers in the self - positioning results are further removed through fitting residuals. Using the smoothed self - positioning data can effectively improve the orbit determination accuracy.

[0098] S4: Satellite orbit determination: Based on the optimized self - positioning results, the low - Earth orbit satellite is orbited using a dynamic fitting method to obtain the final low - Earth orbit satellite TT&C orbit.

[0099] Specifically, the satellite orbit determination process includes:

[0100] Assume the on - satellite self - positioning position at time ;

[0101] The position and velocity at the initial orbit time can be fitted through the finally fitted polynomial. Using the position and velocity at the initial orbit time, the position of the low - Earth orbit satellite at time is obtained through integration , and then the following N - dimensional observation equation is established:

[0102] = ;

[0103] In the formula, is the satellite motion state transition matrix at time , , are respectively the corrections of the position, velocity and dynamic parameters of the low-earth orbit satellite at the initial orbit time;

[0104] The least-squares optimal solution is obtained for the N-dimensional observation equation and iterative calculations are carried out until , , meet the iterative convergence conditions, and then the final orbit of the low-earth orbit satellite is obtained through orbit integration.

[0105] A method for determining the tracking and control orbit of a low-earth orbit satellite provided by the above embodiment calculates the real-time satellite position based on on-board GNSS pseudorange data and broadcast ephemeris; aiming at the gross error problem existing in high-dynamic positioning, it is proposed to combine the eigenvalue of satellite spatial distribution and the threshold of the number of satellites used for positioning to eliminate the gross error in the self-positioning result; for the problem that the self-positioning result is unevenly spaced discrete points, a sliding window polynomial is used to fit the self-positioning result, and the fitting residual is used to further eliminate the outliers in the discrete points; a series of smoothed self-positioning results are obtained based on the polynomial interpolation after fitting, and the final tracking and control orbit of the low-earth orbit satellite is determined through dynamic orbit fitting. The data used in the present invention only includes single-frequency on-board GNSS pseudorange data and broadcast ephemeris, and the types and amounts of data used are small, so the cost is low; only simple pseudorange single-point positioning algorithms and dynamic orbit determination methods are used, and the implementation process is relatively simple and the calculation amount is small; therefore, the technical solution of the present invention has a relatively simple implementation process, low calculation cost and good real-time performance, and can realize the determination of the tracking and control orbit of a low-earth orbit satellite without relying on a ground station. Since the self-positioning result is calculated in real time and is processed by multiple outlier eliminations, it gets rid of the dependence on the ground station, is less affected by the orbit maneuver of the low-earth orbit satellite, can reflect the relatively accurate satellite position, and the orbit determination accuracy is better than the result of ground station data tracking and control orbit determination. For the large-scale low-earth orbit constellation, the measurement and control tasks of a large number of low-earth orbit satellites are heavy and complex, and the low-earth orbit satellites have frequent orbit maneuvers. The present technical solution provides an idea for the measurement and control orbit determination problem of low-earth orbit satellites in this situation.

[0106] It should also be supplemented and explained that using a low-orbit satellite on-board receiver for single-point positioning can itself obtain a rough orbital position. However, this result is obtained by solving geometric methods and does not have orbital characteristics. The integration of the dynamic fitting orbit determination method can endow it with dynamic orbital characteristics, making it more in line with the real satellite orbit. The combination of two key technologies is required to realize the measurement and control orbit determination without relying on the ground station system. The difficulty lies in the reasonable preprocessing of the single-point positioning results. Otherwise, if the data preprocessing is inappropriate, it may lead to the divergence of orbit determination errors. The present invention uses an effective method to eliminate the outliers of single-point positioning and perform smoothing processing that conforms to orbital characteristics. The preprocessed data is used as the input for orbit determination. Through this processing, the gross errors in the single-point positioning results can be eliminated, making the orbit determination results for low-orbit satellite measurement and control have higher accuracy and reliability.

[0107] The embodiment of the present invention also provides a device for determining the measurement and control orbit of a low-orbit satellite, including:

[0108] An initial self-positioning module, configured to obtain on-board GNSS pseudorange observations and GNSS broadcast ephemeris, and calculate the self-positioning result of the low-orbit satellite through the pseudorange single-point positioning algorithm;

[0109] A filtering module, configured to, when the number of GNSS satellites used for pseudorange single-point positioning is lower than a first threshold, eliminate the corresponding self-positioning result; calculate the satellite spatial distribution characteristic value of the positioning epoch, and when the satellite spatial distribution characteristic value is greater than a second threshold, eliminate the corresponding self-positioning result;

[0110] An optimization module, configured to perform polynomial fitting on the filtered self-positioning results in a sliding window manner, eliminate the self-positioning results with fitting residuals greater than a third threshold, and perform interpolation processing according to the finally fitted polynomial to obtain the optimized self-positioning result;

[0111] An orbit determination module, configured to perform orbit determination on the low-orbit satellite using the dynamic fitting method based on the optimized self-positioning result to obtain the final measurement and control orbit of the low-orbit satellite.

[0112] It should be understood that the functional unit modules in various embodiments of the present invention can be concentrated in one processing unit, or each unit module can exist physically alone, or two or more unit modules can be integrated in one unit module, and can be implemented in the form of hardware or software.

[0113] The embodiment of the present invention also provides a device for determining the measurement and control orbit of a low-orbit satellite, configured on the low-orbit satellite, including:

[0114] An on-board receiver, configured to receive on-board GNSS pseudorange observations and GNSS broadcast ephemeris, and transmit them to the low-orbit satellite orbit determination module;

[0115] The satellite-side computing module is configured to implement the method for determining the tracking and control orbit of a low-Earth orbit satellite as described above.

[0116] An embodiment of the present invention further provides an electronic device, including:

[0117] A memory, on which a computer program or instruction is stored;

[0118] A processor, configured to load and execute the computer program or instruction to implement the method for determining the tracking and control orbit of a low-Earth orbit satellite as described above.

[0119] An embodiment of the present invention further provides a computer-readable storage medium, on which a computer program or instruction is stored, and when the computer program or instruction is executed by a processor, the method for determining the tracking and control orbit of a low-Earth orbit satellite as described above is implemented.

[0120] It can be understood that the same or similar parts in the above embodiments can be referred to each other, and the content not detailed in some embodiments can be seen in the same or similar content in other embodiments.

[0121] It should be understood that the functional unit modules in each embodiment of the present invention can be concentrated in one processing unit, or each unit module exists physically alone, or two or more unit modules are integrated in one unit module, and can be implemented in the form of hardware or software.

[0122] Those skilled in the art should understand that the embodiments of the present application can be provided as a method, a system, or a computer program product. Therefore, the present application can adopt the form of a complete hardware embodiment, a complete software embodiment, or an embodiment combining software and hardware aspects. Moreover, the present application can adopt the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code.

[0123] The present application is described with reference to the flowcharts and / or block diagrams of methods, devices (systems), and computer program products according to the embodiments of the present application. It should be understood that each process and / or block in the flowchart and / or block diagram, and the combination of processes and / or blocks in the flowchart and / or block diagram can be implemented by computer program instructions. These computer program instructions can be provided to the processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing devices to generate a machine, so that the instructions executed by the processor of the computer or other programmable data processing devices generate means for implementing the functions specified in Figure 1 one process or multiple processes and / or blocks Figure 1 one block or multiple blocks.

[0124] These computer program instructions can also be stored in a computer-readable memory that can direct a computer or other programmable data processing apparatus to function in a particular manner, such that the instructions stored in the computer-readable memory produce an article of manufacture including instruction means that implement the function specified in one or more of the flow Figure 1 acts or multiple acts and / or boxes Figure 1 specified in one or more of the boxes or multiple boxes.

[0125] These computer program instructions can also be loaded onto a computer or other programmable data processing apparatus to cause a series of operational steps to be performed on the computer or other programmable apparatus to produce a computer-implemented process, whereby the instructions executed on the computer or other programmable apparatus provide steps for implementing the function specified in one or more of the flow Figure 1 acts or multiple acts and / or boxes Figure 1 specified in one or more of the boxes or multiple boxes.

[0126] Although the embodiments of the present invention have been shown and described above, it is to be understood that the above embodiments are exemplary and should not be construed as limiting the present invention, and that those of ordinary skill in the art can make changes, modifications, substitutions, and variations to the above embodiments within the scope of the present invention.

Claims

1. A method for determining a low-orbit satellite tracking and control orbit, characterized in that: The steps include: Obtain self-positioning results: Obtain onboard GNSS pseudo-range observations and GNSS broadcast ephemeris, and obtain the self-positioning results of low-orbit satellites through pseudo-range single-point positioning algorithm; Self-positioning result filtering: When the number of GNSS satellites used for pseudo-range single-point positioning is lower than the first threshold, the corresponding self-positioning result is eliminated; Calculate the satellite spatial distribution characteristic value of the positioning epoch, and when the satellite spatial distribution characteristic value is greater than a second threshold, remove the corresponding self-positioning result; Self-positioning result optimization: The filtered self-positioning results are fitted with a polynomial using a sliding window method, and the self-positioning results with fitting residuals greater than the third threshold are eliminated. Interpolation processing is performed based on the final fitted polynomial to obtain the optimized self-positioning results. Satellite orbit determination: Based on the optimized self-positioning results, the low-orbit satellite is orbit determined using the dynamic fitting method to obtain the final low-orbit satellite tracking and control orbit; The satellite spatial distribution characteristic value is calculated by the following method: The geometric distance from the GNSS satellite to the center of the onboard receiver Approximate position of satellite in low orbit Perform Taylor series expansion at , retain the first-order term, and the linearized equation is: ； In the formula, , , , is the geometric distance from the GNSS satellite to the center of the onboard receiver calculated from the approximate position, Indicates GNSS satellite coordinates; make ,have , subscript n is the number of GNSS satellites observed by the onboard receiver in the current epoch; Then the satellite spatial distribution characteristic value is .

2. The method for determining the low-orbit satellite tracking and control orbit according to claim 1, characterized in that: The self-positioning result of the low-orbit satellite is obtained by solving the pseudo-range single-point positioning algorithm, specifically including: Spaceborne GNSS pseudorange observations The observation equation is expressed as follows: ； In the formula, the subscript , Represent the frequency and satellite receiver respectively, the superscript represents a GNSS satellite, Indicates the geometric distance from the GNSS satellite to the center of the onboard receiver; and They represent the receiver clock error and satellite clock error contained in the pseudorange respectively; and Respectively represent the frequency The hardware delay deviation between the receiver and the satellite, where the hardware delay at the receiver is received by the receiver clock error during positioning; represents the tropospheric delay error of pseudorange measurement, Indicates frequency The ionospheric delay, is the pseudorange observation noise, is the speed of light in a vacuum; The error equation after linearization of the observation equation of the satellite-borne GNSS pseudorange observation value is expressed as follows: ； In the formula, is the observation correction vector, is the design matrix, is the residual vector of pseudorange observations, is the unknown parameter correction vector, ,in , , is the coordinate correction number of the satellite receiver; The unknown parameter correction vector can be obtained by the least squares principle: ； In the formula, is the coefficient matrix of the normal equation, is the observation value weight matrix; Iterative calculation is performed until the unknown parameter correction vector meets the iterative convergence condition, and then the coordinates of the onboard receiver at the current epoch are obtained as the self-positioning result of the low-orbit satellite.

3. The method for determining the low-orbit satellite tracking and control orbit according to claim 1, characterized in that: The self-positioning result optimization specifically includes: Set the sliding window length and sliding step size. The sliding window length is equal to the sliding step size. Divide the filtered self-positioning results into multiple segments of self-positioning results through a sliding window; For each segment of the self-positioning results after division, a polynomial fitting is performed, and the self-positioning results whose fitting residuals are greater than the third threshold are eliminated. The polynomial fitting and elimination process is iterated until the fitting residuals of the retained self-positioning results are all less than the third threshold, thereby obtaining the final fitting polynomial; Interpolation processing is performed according to the final fitted polynomial to obtain the optimized self-positioning result.

4. The method for determining the low-orbit satellite tracking and control orbit according to claim 1, characterized in that: The satellite orbit determination specifically includes: Assumptions The self-positioning position on the satellite at time ; The position and velocity at the initial orbit time can be obtained by fitting the final fitting polynomial, and the position and velocity at the initial orbit time can be obtained by integration. The low-orbit satellite position at the time , and then establish the N-dimensional observation equation as follows: = ； In the formula, for The satellite motion state transfer matrix at time , , , are the corrections of the position, velocity and dynamic parameters of the low-orbit satellite at the time of initial orbit; Find the least squares optimal solution for the N-dimensional observation equation and perform iterative calculation until , , The iterative convergence conditions are met, and then the final orbit of the low-orbit satellite is obtained through orbit integration.

5. The method for determining the low-orbit satellite tracking and control orbit according to claim 1, characterized in that: The first threshold is 8; the second threshold is 2; and the third threshold is 10m.

6. A low-orbit satellite tracking and control orbit determination device, characterized in that: include: The initial self-positioning module is used to obtain the onboard GNSS pseudo-range observation values ​​and GNSS broadcast ephemeris, and obtain the self-positioning result of the low-orbit satellite through the pseudo-range single-point positioning algorithm; A filtering module, used for discarding the corresponding self-positioning result when the number of GNSS satellites used for pseudo-range single point positioning is lower than a first threshold; Calculate the satellite spatial distribution characteristic value of the positioning epoch, and when the satellite spatial distribution characteristic value is greater than a second threshold, remove the corresponding self-positioning result; An optimization module is used to perform polynomial fitting on the filtered self-positioning results by using a sliding window method, eliminate the self-positioning results whose fitting residuals are greater than a third threshold, and perform interpolation processing according to the final fitted polynomial to obtain an optimized self-positioning result; The orbit determination module is used to determine the orbit of the low-orbit satellite using the dynamic fitting method based on the optimized self-positioning results to obtain the final low-orbit satellite tracking and control orbit; The satellite spatial distribution characteristic value is calculated by the following method: The geometric distance from the GNSS satellite to the center of the onboard receiver Approximate position of satellite in low orbit Perform Taylor series expansion at , retain the first-order term, and the linearized equation is: ； In the formula, , , , is the geometric distance from the GNSS satellite to the center of the onboard receiver calculated from the approximate position, Indicates GNSS satellite coordinates; make ,have , subscript n is the number of GNSS satellites observed by the onboard receiver in the current epoch; Then the satellite spatial distribution characteristic value is .

7. A low-orbit satellite tracking and control orbit determination device, characterized in that: Configured on low-orbit satellites, including: The onboard receiver is used to receive onboard GNSS pseudorange observations and GNSS broadcast ephemeris and transmit them to the low-orbit satellite orbit determination module; The satellite-side computing module is configured to implement the low-orbit satellite tracking and control orbit determination method as described in any one of claims 1 to 5.

8. An electronic device, characterized in that: include: Memory on which computer programs or instructions are stored; A processor is used to load and execute the computer program or instruction to implement the low-orbit satellite tracking and control orbit determination method as described in any one of claims 1 to 5.

9. A computer-readable storage medium having a computer program or instruction stored thereon, characterized in that: When the computer program or instruction is executed by the processor, the method for determining the low-orbit satellite tracking and control orbit as described in any one of claims 1 to 5 is implemented.

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