Satellite orbit determination method and device, storage medium and electronic equipment

Abstract

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.

Description

Technical Field

[0001] The present application relates to the field of computers, and in particular, to a method and device for determining a satellite operating orbit, a storage medium, and an electronic device. Background Art

[0002] In the field of satellite navigation technology, precise orbit control of a satellite constellation is the key to achieving high-precision positioning services. However, in related technologies, relying solely on inter-satellite link observation data for autonomous orbit determination, it is difficult to avoid the overall offset and rotation errors of the satellite constellation, resulting in each satellite deviating from its predetermined orbit. Therefore, in related technologies, there is a technical problem that the satellite cannot operate along the correct orbit due to the overall offset and rotation of the satellite constellation.

[0003] In view of the above problems, no effective solution has been proposed yet. Summary of the Invention

[0004] Embodiments of the present application provide a method and device for determining a satellite operating orbit, a storage medium, and an electronic device, so as to at least solve the technical problem that the satellite cannot operate along the correct orbit due to the overall offset and rotation of the satellite constellation.

[0005] According to an aspect of an embodiment of the present application, a method for determining a satellite operating orbit is provided, including: when a target communication satellite can receive a navigation signal, obtaining an initial satellite position of the target communication satellite, and determining a reference orbit of the target communication satellite based on the spatio-temporal information of a navigation satellite, where the spatio-temporal information includes the position and clock error of the navigation satellite, and the reference orbit is an operating orbit determined by the target communication satellite during the autonomous orbit determination process; performing orbit fitting on the initial satellite position, the reference orbit, and initial non-conservative force model parameters to obtain a target satellite position and target non-conservative force model coefficients; performing orbit integration on the target communication satellite according to pre-determined conservative force model coefficients, the target satellite position, and the target non-conservative force model coefficients to determine an integrated orbit of the target communication satellite; comparing the integrated orbit with the reference orbit to determine a position error sequence of the target communication satellite, and determining process noise covariance information at different integration durations in the integrated orbit based on the position error sequence, where the position error sequence is used to indicate the change of the position deviation of the target communication satellite relative to the target operating orbit over time during the autonomous orbit determination process; when the target communication satellite cannot receive the navigation signal, using inter-satellite link data for orbit filtering to determine the target operating orbit.

[0006] According to another aspect of the embodiments of the present application, there is also provided a device for determining a satellite operating orbit, including: an acquisition module, configured to acquire an initial satellite position of the target communication satellite when the target communication satellite can receive a navigation signal, and determine a reference orbit of the target communication satellite based on the spatio-temporal information of the navigation satellite, where the spatio-temporal information includes the position and clock offset of the navigation satellite, and the reference orbit is an operating orbit determined by the target communication satellite during the autonomous orbit determination process; a fitting module, configured to perform orbit fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters to obtain a target satellite position and target non-conservative force model coefficients; an integration module, configured to perform orbit integration on the target communication satellite according to the pre-determined conservative force model coefficients, the target satellite position, and the target non-conservative force model coefficients to determine an integrated orbit of the target communication satellite; a comparison module, configured to compare the integrated orbit and the reference orbit to determine a position error sequence of the target communication satellite, and determine process noise covariance information at different integration durations in the integrated orbit based on the position error sequence, where the position error sequence is used to indicate the change of the position deviation of the target communication satellite relative to the target operating orbit over time during the autonomous orbit determination process; a determination module, configured to perform orbit filtering using inter-satellite link data to determine the target operating orbit when the target communication satellite cannot receive the navigation signal.

[0007] Optionally, the device is configured to compare the integrated orbit and the reference orbit in the following manner to determine the position error sequence of the target communication satellite, and determine the process noise covariance information at different integration durations in the integrated orbit based on the position error sequence: acquire an integrated orbit point corresponding to the integrated orbit at any moment, and acquire a reference orbit point corresponding to the reference orbit at the any moment; determine error data corresponding to the target communication satellite at the any moment according to the integrated orbit point and the reference orbit point, where the error data includes at least one of position error data and clock offset error data; generate the position error sequence according to the error data.

[0008] Optionally, the device is configured to perform orbit filtering using inter-satellite link data to determine the target operating orbit when the target communication satellite cannot receive the navigation signal in the following manner: when the target communication satellite cannot receive the navigation signal, determine the initial operating state of the predicted orbit according to the initial orbit state quantity and the initial process noise covariance information; use the inter-satellite link data to correct the predicted orbit according to the Kalman filtering algorithm to obtain the target operating orbit; wherein, the inter-satellite link data is used to determine the observation residual in the process of the orbit filtering, and the observation residual represents the difference between the inter-satellite ranging observation value indicated by the inter-satellite link data and the expected inter-satellite ranging value of the predicted orbit.

[0009] Optionally, the device is further configured to: determine that the target communication satellite cannot receive the navigation signal when the navigation signal is denied; determine that the target communication satellite cannot receive the navigation signal when the target communication satellite is in a state without ground support.

[0010] Optionally, the device is configured to perform orbit integration on the target communication satellite according to a pre-determined conservative force model coefficient, the target satellite position, and the target non-conservative force model coefficient in the following manner to obtain an integrated orbit: use the least squares method to fit the target non-conservative force model coefficient to determine the target non-conservative force fitting coefficient; use the conservative force model coefficient and the target non-conservative force fitting coefficient to determine the dynamic model of the target communication satellite; based on the dynamic model and the target satellite position, solve the state information of the target communication satellite at any moment to obtain the integrated orbit, wherein the state information includes at least one of the position data, clock data, and velocity data of the target communication satellite.

[0011] Optionally, the device is configured to perform orbit fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters in the following manner to obtain the target satellite position and the target non-conservative force model coefficient: determine the initial non-conservative force model parameters according to the type of the target communication satellite; use the least squares method to fit the initial satellite position, the reference orbit, and the initial non-conservative force model parameters to obtain the target satellite position and the target non-conservative force model coefficient.

[0012] Optionally, the device is configured to determine the initial non-conservative force model parameters according to the type of the target communication satellite in the following manner: when the type of the target communication satellite is a low-Earth orbit type, the initial non-conservative force model parameters include at least one of an atmospheric drag model parameter, a solar radiation pressure model parameter, and a magnetic force influence parameter; when the type of the target communication satellite is a medium-Earth orbit type, the initial non-conservative force model parameters include at least one of the solar radiation pressure model parameter, a lunar gravitational perturbation parameter, and a third-body gravitational model parameter; when the type of the target communication satellite is a geosynchronous orbit type, the initial non-conservative force model parameters include at least one of an Earth gravity field model parameter, the lunar gravitational perturbation parameter, and a solar gravitational perturbation parameter.

[0013] Optionally, the device is further configured to: the inter-satellite link data includes at least one of a first geometric distance, a first distance correction term, a first error correction term, and a first observation residual term between the target communication satellite and any one of the other communication satellites when the target communication satellite transmits a signal to the any one of the other communication satellites; and at least one of a second geometric distance, a second distance correction term, a second error correction term, and a second observation residual term between the any one of the other communication satellites and the target communication satellite when the any one of the other communication satellites transmits a signal to the target communication satellite; wherein, the first distance correction term is determined by the product of the speed of light and the clock deviation between the transmission time of the signal transmitted by the target communication satellite and the reception time of the signal received by the any one of the other communication satellites; the second distance correction term is determined by the product of the speed of light and the clock deviation between the transmission time of the signal transmitted by the any one of the other communication satellites and the reception time of the signal received by the target communication satellite; the first error correction term and the second error correction term are both determined by the satellite antenna phase center deviation, the relativistic effect, and the atmospheric delay; the first observation residual term and the second observation residual term are both determined by the observation noise generated during the acquisition of the inter-satellite link data.

[0014] Optionally, the device is further configured to: determine the first geometric distance based on the position of the target communication satellite at the signal transmission time and the position of any communication satellite at the signal reception time; determine the second geometric distance based on the position of any communication satellite at the signal transmission time and the position of the target communication satellite at the signal reception time; adjust the first geometric distance by using at least one of the first distance correction term, the first error correction term, and the first observation residual term, and perform normalization processing on the first geometric distance by using the clock error of the predicted orbit generated in the orbit filtering; adjust the second geometric distance by using at least one of the second distance correction term, the second error correction term, and the second observation residual term, and perform normalization processing on the second geometric distance by using the clock error of the predicted orbit; perform first-order Taylor expansion on the first geometric distance and the second geometric distance respectively, retain the first-order terms, and obtain the first linearized geometric distance and the second linearized geometric distance; normalize the first linearized geometric distance and the second linearized geometric distance to the first target geometric distance and the second target geometric distance at the reference time respectively; and determine the target operating orbit based on the first target geometric distance, the second target geometric distance, and the position error sequence.

[0015] Optionally, the device is configured to perform orbit fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters in the following manner to obtain the target satellite position and the target non-conservative force model coefficients: the initial non-conservative force model parameters are determined by the type of the target communication satellite, where the type of the target communication satellite is any one of a low-earth orbit type, a medium-earth orbit type, and a high-earth orbit type.

[0016] Optionally, the device is configured to perform orbit filtering by using the inter-satellite link data to determine the target operating orbit in the case where the target communication satellite cannot receive the navigation signal in the following manner: in the case where the target communication satellite cannot receive the navigation signal, obtain the relative distance and relative clock difference between the target communication satellite and other communication satellites indicated by the inter-satellite link data, where the target communication satellite and the other communication satellites are in the same satellite constellation; perform the orbit filtering based on the relative distance and the relative clock difference to determine the target operating orbit, where in the process of the orbit filtering, the initial orbit state quantity is determined by the integrated orbit, and the initial process noise covariance is determined by the process noise covariance information.

[0017] According to another aspect of the embodiments of the present application, there is also provided a computer-readable storage medium, in which a computer program is stored, where the computer program is configured to execute the method for determining the satellite operating orbit as described above when running.

[0018] According to another aspect of the embodiments of the present application, there is provided a computer program product or a computer program, which includes computer instructions stored in a computer-readable storage medium. The processor of the computer device reads the computer instructions from the computer-readable storage medium, and the processor executes the computer instructions, so that the computer device executes the method for determining the satellite orbit as described above.

[0019] According to another aspect of the embodiments of the present application, there is also provided an electronic device, including a memory and a processor. A computer program is stored in the memory, and the processor is configured to execute the method for determining the satellite orbit as described above through the computer program.

[0020] In the embodiments of the present application, when the target communication satellite can receive navigation signals, the initial satellite position of the target communication satellite is obtained, and the reference orbit of the target communication satellite is determined based on the space-time information of the navigation satellite; orbit fitting is performed on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters to obtain the target satellite position and the target non-conservative force model coefficients; orbit integration is performed on the target communication satellite according to the pre-determined conservative force model coefficients, the target satellite position, and the target non-conservative force model coefficients to determine the integrated orbit of the target communication satellite; the integrated orbit and the reference orbit are compared to determine the position error sequence of the target communication satellite, and the process noise covariance information at different integration times in the integrated orbit is determined based on the position error sequence; when the target communication satellite cannot receive navigation signals, orbit filtering is performed using inter-satellite link data to determine the target operating orbit. By comprehensively using non-conservative force model parameters, position error sequences, and inter-satellite link data, the accumulation of the overall rotation error of the low-earth orbit constellation is effectively suppressed, the autonomous operation ability and orbit determination accuracy of the satellite network are significantly improved, the requirements for real-time autonomous determination and robust service of the entire network orbit of the satellite constellation in an adversarial scenario are met, and further the technical problem that the satellite cannot operate in the correct orbit due to the overall offset rotation of the satellite constellation is solved. Description of the Drawings

[0021] The drawings described herein are used to provide a further understanding of the present application, and constitute a part of the present application. The schematic embodiments of the present application and their descriptions are used to explain the present application, and do not constitute an improper limitation to the present application. In the drawings:

[0022] Figure 1 is a schematic diagram of an application environment of an optional method for determining a satellite orbit according to an embodiment of the present application;

[0023] Figure 2 is a schematic flowchart of an optional method for determining a satellite orbit according to an embodiment of the present application;

[0024] Figure 3 is a schematic flowchart of an optional method for determining a satellite operating orbit according to an embodiment of the present application;

[0025] Figure 4 is a schematic diagram of two-way inter-satellite link observation of an optional method for determining a satellite operating orbit according to an embodiment of the present application;

[0026] Figure 5 is a schematic structural diagram of an optional device for determining a satellite operating orbit according to an embodiment of the present application;

[0027] Figure 6 is a schematic structural diagram of an optional product for determining a satellite operating orbit according to an embodiment of the present application;

[0028] Figure 7 is a schematic structural diagram of an optional electronic device according to an embodiment of the present application. Detailed implementation manners

[0029] In order to enable those skilled in the art to better understand the solutions of the present application, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present application.

[0030] It should be noted that the terms "first", "second", etc. in the specification and claims of the present application and the above drawings are used to distinguish similar objects, and are not necessarily used to describe a specific order or sequence. It should be understood that such data can be interchanged under appropriate circumstances so that the embodiments of the present application described herein can be implemented in an order other than those illustrated or described herein. In addition, the terms "comprising" and "having" and any variations thereof are intended to cover non-exclusive inclusion. For example, a process, method, system, product or device comprising a series of steps or units is not necessarily limited to those steps or units clearly listed, but may include other steps or units not clearly listed or inherent to these processes, methods, products or devices.

[0031] The present application will be described below with reference to the embodiments:

[0032] According to one aspect of the embodiments of the present application, a method for determining a satellite operating orbit is provided. Optionally, in this embodiment, the above method for determining a satellite operating orbit can be applied to, for example, Figure 1 the hardware environment composed of a server 101 and a terminal device 103 as shown. AsFigure 1 As shown in the figure, the server 101 is connected to the terminal device 103 through a network, and can be used to provide services for the terminal device or the application installed on the terminal device. The application 107 can be a video application, an instant messaging application, a browser application, an educational application, a game application, etc. The database 105 can be set on the server or independently of the server, and is used to provide data storage services for the server 101. For example, a game data storage server. The above network can include but is not limited to: a wired network, a wireless network. Among them, the wired network includes: a local area network, a metropolitan area network, and a wide area network. The wireless network includes: Bluetooth, WIFI, and other networks that implement wireless communication. The terminal device 103 can be a terminal configured with an application, and can include but is not limited to at least one of the following: a mobile phone (such as an Android mobile phone, an iOS mobile phone, etc.), a laptop computer, a tablet computer, a personal digital assistant, a MID (Mobile Internet Devices), a PAD, a desktop computer, a smart TV, a smart voice interaction device, a smart home appliance, a vehicle-mounted terminal, an aircraft, a virtual reality (VR) terminal, an augmented reality (AR) terminal, a mixed reality (MR) terminal, and other computer devices. The above server can be a single server, or a server cluster composed of multiple servers, or a cloud server.

[0033] Combined with Figure 1 As shown in the figure, the method for determining the above satellite orbit can be executed by an electronic device, and the electronic device can be a terminal device or a server. The method for determining the above satellite orbit can be implemented separately by the terminal device or the server, or jointly implemented by the terminal device and the server.

[0034] The above is only an example, and this embodiment does not make specific limitations. Optionally, as an alternative implementation manner, as Figure 2 shown in the figure, the method for determining the above satellite orbit includes:

[0035] S202, when the target communication satellite can receive the navigation signal, obtain the initial satellite position of the target communication satellite, and determine the reference orbit of the target communication satellite based on the space-time information of the navigation satellite. Among them, the space-time information includes the position and clock error of the navigation satellite, and the reference orbit is the operating orbit determined by the target communication satellite during the autonomous orbit determination process;

[0036] Optionally, in the embodiments of the present application, obtaining the initial satellite position of the target communication satellite refers to measuring or estimating information such as the three-dimensional space coordinates of the target communication satellite at a specific moment. The target communication satellite here can be a low-earth orbit satellite, a medium-earth orbit satellite, a geosynchronous orbit satellite, etc., and the present application does not make any limitations in this regard; the above spatio-temporal information refers to the precise position and clock deviation information of navigation satellites (such as GPS or Beidou satellites) at a specific moment; the above reference orbit is the expected operating orbit calculated and determined by the target communication satellite itself during the autonomous orbit determination process based on the spatio-temporal information provided by these navigation satellites, combined with its own dynamic model and observation data, including but not limited to the expected position and velocity vector of the target satellite during the autonomous orbit determination period.

[0037] S204, perform orbit fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters to obtain the target satellite position and the target non-conservative force model coefficients;

[0038] Optionally, in the embodiments of the present application, performing orbit fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters means using an optimization algorithm to adjust the satellite position and the dynamic model parameters to minimize the difference between the observed value and the model prediction value; the above target satellite position represents the current optimal estimated position information of the target communication satellite; the above target non-conservative force model coefficients represent the non-conservative force parameters that affect the operating orbit of the target communication satellite, such as solar radiation pressure, atmospheric drag, etc.

[0039] S206, perform orbit integration on the target communication satellite according to the pre-determined conservative force model coefficients, the target satellite position, and the target non-conservative force model coefficients to determine the integrated orbit of the target communication satellite;

[0040] Optionally, in the embodiments of the present application, the above orbit integration refers to the prediction process of the operating orbit of the target communication satellite based on the dynamic model, using the conservative force model coefficients (such as gravity parameters), the target satellite position, and the target non-conservative force model coefficients, and the above integrated orbit is the result of this prediction. The integrated orbit reflects the predicted trajectory of the position and velocity of the satellite over time considering the influence of all acting forces.

[0041] S208, compare the integrated orbit with the reference orbit to determine the position error sequence of the target communication satellite, and determine the process noise covariance information at different integration durations in the integrated orbit based on the position error sequence, where the position error sequence is used to indicate the change of the position deviation of the target communication satellite relative to the target operating orbit over time during the autonomous orbit determination process;

[0042] Optionally, in the embodiments of the present application, the above comparison refers to comparing the integral orbit with the reference orbit to determine the difference between the two orbits, that is, the above position error sequence; and the position error sequence contains the position deviation information of the target communication satellite relative to the target operating orbit at different time points, which is used to identify and correct the overall rotation and drift errors of the constellation. The above process noise covariance information refers to a matrix that describes the statistical characteristics of the prediction errors of satellite states (such as position, velocity, clock deviation, etc.) in orbit filtering algorithms such as Kalman filtering or similar state estimation. This matrix quantifies the state estimation uncertainty caused by factors such as incomplete dynamic models, external random disturbances (such as non-conservative forces like solar radiation pressure and atmospheric drag), and measurement noise. During the autonomous orbit determination process, it is used to adjust the weights of the filter for the prediction model and observation data to achieve optimal state estimation. The process noise covariance information at different integral durations reflects the uncertainty level of satellite state prediction over time and is a key parameter for optimizing the orbit solution accuracy and robustness.

[0043] S210. When the target communication satellite cannot receive navigation signals, use the inter-satellite link data for orbit filtering to determine the target operating orbit.

[0044] Optionally, in the embodiments of the present application, the above orbit filtering refers to the process of real-time updating and correcting the satellite orbit through a filtering algorithm based on inter-satellite link data and the position error sequence when the target communication satellite loses external GNSS signals. The above inter-satellite link data provides observations of the relative distance and relative velocity between satellites, while the position error sequence describes the change of the satellite's deviation from its target orbit over time; the above target operating orbit refers to the real-time operating trajectory determined by the target communication satellite after being corrected by the orbit filtering algorithm when it loses external GNSS signals, which synthesizes the position error sequence and inter-satellite link data and reflects the latest position and velocity information of the target communication satellite within the constellation, ensuring that the target communication satellite can maintain its operation on the correct orbit even in a complex environment.

[0045] It should be noted that the above satellite constellation can be understood as a group of satellite clusters operating on a predetermined orbit.

[0046] Exemplarily, first accurately obtain the initial position of the target communication satellite, and then use the high-precision spatio-temporal information (including position and clock deviation) of the navigation satellite to determine the reference operating orbit of the target communication satellite; next, perform orbit fitting on the initial satellite position, the preset reference orbit, and the initial non-conservative force model parameters based on the satellite type to obtain the optimized target satellite position and the model coefficients reflecting the actual non-conservative force influence. These non-conservative forces here can include, but are not limited to, solar radiation pressure, atmospheric drag, etc.

[0047] Next, using the determined conservative force model coefficients (such as the Earth's gravitational parameter) and the target satellite position and target non-conservative force model coefficients obtained by fitting, orbit integration is performed on the target communication satellite to predict its future orbit, obtaining an integrated orbit. By comparing the integrated orbit with the reference orbit, the position deviation of the target communication satellite during autonomous orbit determination can be determined, that is, the position error sequence. Based on the position error sequence, the uncertainty level of satellite state prediction is further analyzed, and the most suitable process noise covariance information at different integration durations is determined through statistical analysis and adaptive algorithms.

[0048] When the target communication satellite cannot receive navigation signals, orbit filtering is performed using inter-satellite link data to determine the target operating orbit. During the orbit filtering process, the initial orbit state variables are determined by the previous integrated orbit, and the initial process noise covariance information is determined based on the previously determined process noise covariance information. The inter-satellite link data provides the relative distance and relative clock difference information between the target communication satellite and other satellites within the constellation. By combining the inter-satellite link data and prior constraints (i.e., the initial orbit state variables and process noise covariance information), the filtering algorithm can effectively suppress the overall rotation error of the constellation, maintain the stability and operating accuracy of the satellite network, and ensure reliable services can still be provided in an adversarial or GNSS-denied environment.

[0049] In an exemplary embodiment, Figure 3 is a schematic flowchart of an optional method for determining a satellite operating orbit according to an embodiment of the present application, as Figure 3 shown, including but not limited to:

[0050] S1, using the target communication satellite to collect the spatio-temporal information of GNSS satellites (the above-mentioned navigation satellites) in the current autonomous processing cycle (including but not limited to the position, clock difference, etc. of GNSS satellites), and calculating the initial satellite position and reference orbit of the communication satellite at the current initial moment.

[0051] S2, perform orbit fitting on the initial satellite position, reference orbit, and related non-conservative forces (atmospheric drag, Earth albedo radiation pressure, solar radiation pressure, etc., which are related to the satellite type, such as low Earth orbit, high Earth orbit, etc.) (including but not limited to algorithms such as the least squares method, etc.), to obtain a more accurate satellite position, denoted as the target satellite position and non-conservative force model coefficients.

[0052] S3, combining the conservative force model coefficients and the target non-conservative force fitting coefficients obtained by fitting the non-conservative force model coefficients, and using a numerical integration method to perform orbit integration on the target communication satellite. It can be understood that the current target satellite position is used to predict the satellite position at the next moment in the future, and the satellite positions at multiple moments are recursively obtained to obtain an integrated orbit.

[0053] S4. Compare the integral orbit with the reference orbit to obtain the position error of the target communication satellite at different future moments, that is, the position error sequence. Further, conduct a detailed statistical analysis on the position error sequence, including calculating the mean, variance of the error, and the autocorrelation function of the time series, etc., to identify the statistical characteristics of the error. Then, adopt an adaptive algorithm to adjust the process noise covariance information to achieve dynamic update of the process noise covariance information and better match the uncertainty of satellite state prediction.

[0054] S5. When in the GNSS denial or no ground support state, use the process noise covariance information and inter-satellite link data for orbit filtering (including but not limited to Kalman filtering, etc.) to determine the target operating orbit of the communication satellite.

[0055] Among them, the above inter-satellite link data is directly obtained by mutual observation among the communication satellites in the satellite constellation. Taking communication satellite A and communication satellite B as examples, it can be expressed as , and the formula represents various information included in the inter-satellite link data ( represents the geometric distance between satellites, represents the satellite clock error of communication satellite B at time t1, represents when (i.e., the moment when satellite A transmits the signal, considering the signal transmission delay), the satellite clock error of communication satellite A, represents the error corrections such as the satellite antenna phase centers of communication satellite A and communication satellite B, relativistic effects, and atmospheric delays, represents the random noise or unmodeled residuals in the observation), and c is the speed of light.

[0056] Through the embodiments of the present application, when the target communication satellite can receive navigation signals, the initial satellite position of the target communication satellite is obtained, and the reference orbit of the target communication satellite is determined based on the spatio-temporal information of the navigation satellite, where the spatio-temporal information includes the position and clock offset of the navigation satellite, and the reference orbit is the operating orbit determined by the target communication satellite during the autonomous orbit determination process; orbit fitting is performed on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters to obtain the target satellite position and the target non-conservative force model coefficients, where the initial non-conservative force model parameters are determined by the type of the target communication satellite; the target communication satellite is orbit integrated according to the pre-determined conservative force model coefficients, the target satellite position, and the target non-conservative force model coefficients to determine the integrated orbit of the target communication satellite; the integrated orbit and the reference orbit are compared to determine the position error sequence of the target communication satellite, and the process noise covariance information at different integration durations in the integrated orbit is determined based on the position error sequence, where the position error sequence is used to indicate the change of the position deviation of the target communication satellite relative to the target operating orbit over time during the autonomous orbit determination process; when the target communication satellite cannot receive navigation signals, inter-satellite link data is used for orbit filtering to determine the target operating orbit, where the initial orbit state quantity in the process of orbit filtering is determined by the integrated orbit, and the initial process noise covariance information in the process of orbit filtering is determined by the process noise covariance information, and the inter-satellite link data is used to indicate the relative distance and relative clock offset between the target communication satellite and other communication satellites, and the target communication satellite and other communication satellites are in the same satellite constellation. By comprehensively using the non-conservative force model parameters, the position error sequence, and the inter-satellite link data, the accumulation of the overall rotation error of the low-earth orbit constellation is effectively suppressed, the autonomous operation ability and orbit determination accuracy of the satellite network are significantly improved, the requirements for real-time autonomous orbit determination and robust service of the entire satellite constellation in an adversarial scenario are met, and thus the technical problem that the satellite cannot operate along the correct orbit due to the overall offset and rotation of the satellite constellation is solved.

[0057] As an alternative solution, the comparison of the integrated orbit and the reference orbit to determine the position error sequence of the target communication satellite and the determination of the process noise covariance information at different integration durations in the integrated orbit based on the position error sequence include: obtaining the integrated orbit point corresponding to the integrated orbit at any moment, and obtaining the reference orbit point corresponding to the reference orbit at the any moment; determining the error data corresponding to the target communication satellite at the any moment according to the integrated orbit point and the reference orbit point, where the error data includes at least one of position error data and clock offset error data; and generating the position error sequence according to the error data.

[0058] Optionally, in the embodiments of the present application, the above-mentioned integral orbit points refer to the predicted position and velocity information of the target communication satellite calculated at preset time intervals during the integration process; the above-mentioned reference orbit points are the expected positions and velocities of the satellite predicted based on the initial conditions and the dynamic model at the same time intervals; by calculating the difference between the integral orbit points and the reference orbit points, the error data of the above-mentioned target communication satellite at each time point can be obtained, including but not limited to position error, clock error, and velocity error.

[0059] In an exemplary embodiment, during the autonomous orbit determination process of the target communication satellite, an integral orbit point is obtained through orbit integration every 300 seconds. At the same time, according to the reference orbit determined by the GNSS ephemeris data, a set of reference orbit points is also provided at the same time point. By comparing the coordinates of the two orbit points at the 300-second moment, the position error data is calculated. By comparing the deviation of the clock error, the clock error data is determined, and by comparing the velocity vectors, the velocity error data is obtained. Arranging these error data in chronological order constitutes the above-mentioned position error sequence that describes the change of the satellite orbit deviation over time.

[0060] Through the embodiments of the present application, by comparing the integral orbit and the reference orbit, the technical effect of accurately monitoring and evaluating the orbit error of the target communication satellite is achieved, and the purpose of accurately identifying its orbit deviation when the satellite loses the external navigation signal and providing data support for subsequent orbit filtering and correction is achieved.

[0061] As an optional solution, in the case that the above-mentioned target communication satellite cannot receive the above-mentioned navigation signal, using the inter-satellite link data for orbit filtering to determine the above-mentioned target operating orbit includes: in the case that the above-mentioned target communication satellite cannot receive the above-mentioned navigation signal, determining the initial operating state of the predicted orbit according to the above-mentioned initial orbit state quantity and the above-mentioned initial process noise covariance information; using the above-mentioned inter-satellite link data to correct the above-mentioned predicted orbit according to the Kalman filtering algorithm to obtain the above-mentioned target operating orbit; wherein, the above-mentioned inter-satellite link data is used to determine the observation residual in the process of the above-mentioned orbit filtering, and the above-mentioned observation residual represents the difference between the inter-satellite ranging observation value indicated by the above-mentioned inter-satellite link data and the expected inter-satellite ranging value of the above-mentioned predicted orbit.

[0062] Optionally, in the embodiments of the present application, the above prediction orbit can be understood as the predicted orbit position and state parameters of the satellite at the next moment or in a future period based on the currently known satellite state information (such as position, velocity, and clock deviation, etc.) and the dynamic model. The above inter-satellite ranging observation value refers to the actual distance data directly measured through the inter-satellite link between the target communication satellite and other communication satellites in the constellation. The expected inter-satellite ranging value refers to the expected distance between the target communication satellite and other communication satellites in the constellation calculated according to the prediction orbit.

[0063] It should be noted that the Kalman filter algorithm continuously adjusts the orbit parameters of the satellite through two stages of state prediction and observation update to minimize the error. The above position error sequence is used to dynamically adjust the weight of the prediction orbit during the orbit filtering process to ensure that the orbit filtering pays more attention to those data points with smaller errors and more accurate positions. The inter-satellite link data is used to calculate the observation residual, that is, the difference between the actual inter-satellite ranging observation value and the expected inter-satellite ranging calculated according to the prediction orbit.

[0064] In an exemplary embodiment, when the target communication satellite encounters a GNSS denial environment during the autonomous orbit determination process and is unable to receive external navigation signals, first, based on the obtained position error sequence, the weight of the prediction orbit in the orbit filtering process is adjusted to make the prediction closer to the true state of the satellite. Subsequently, using the inter-satellite link data, the observation residual of the inter-satellite ranging is calculated, that is, the gap between the actual ranging and the predicted value. Combining the two for Kalman filtering to update the state vector of the satellite, including parameters such as position, velocity, and clock deviation. Through continuous prediction and update, the high-precision determination of the satellite orbit is maintained, effectively suppressing the accumulation of the overall rotation error of the constellation.

[0065] Through the embodiments of the present application, the strategy of using Kalman filtering to fuse the position error sequence and the inter-satellite link data realizes the technical effect of accurately determining the operating orbit of the target communication satellite in a complex environment where the target communication satellite cannot receive navigation signals, achieving the purpose of effectively controlling the overall rotation error of the constellation and maintaining the autonomous operation ability of the satellite network.

[0066] As an optional solution, the above method further includes: when the above navigation signal is denied, determining that the above target communication satellite cannot receive the above navigation signal; when the above target communication satellite is in a state without ground support, determining that the above target communication satellite cannot receive the above navigation signal.

[0067] Optionally, in the embodiments of the present application, the above-mentioned navigation signal rejection means that the target communication satellite is in an environment where the signal is shielded due to artificial or natural factors on the Earth's surface or in the atmosphere, resulting in its inability to receive signals from the global navigation satellite system (such as GPS, Beidou, Galileo, etc.); the above-mentioned ground support-free state means that the communication link between the satellite and the ground station is interrupted, and it is impossible to obtain data or instructions transmitted on the ground.

[0068] Exemplarily, when the above-mentioned navigation signal receives signals from the global navigation satellite system (such as GPS, Beidou, Galileo, etc.), and the above-mentioned target communication satellite is in a ground support-free state, the above-mentioned target communication satellite cannot receive the above-mentioned navigation signal;

[0069] In another example, when the above-mentioned navigation signal is rejected and the above-mentioned target communication satellite is in a normal ground support state, the above-mentioned target communication satellite also cannot receive the above-mentioned navigation signal;

[0070] In yet another example, when the above-mentioned navigation signal is in a non-rejected state and the above-mentioned target communication satellite is in a ground support-free state, the above-mentioned target communication satellite still cannot receive the above-mentioned navigation signal, that is, when there is a reception blind spot or signal acquisition obstacle for the target communication satellite, the above-mentioned target communication satellite is still determined to be unable to receive the above-mentioned navigation signal.

[0071] Generally speaking, when at least one of the above-mentioned navigation signal rejection and the target communication satellite being in a ground support-free state occurs, the above-mentioned target communication satellite will be determined to be unable to receive the above-mentioned navigation signal.

[0072] It should be noted that the navigation signal rejection and the ground support-free state can be caused by signal interference, natural phenomena (such as solar storms), equipment failures, satellite deployment strategies (such as optimizing the relative positions between satellites), or specific operation requirements. The solutions proposed in the present application are applicable to all these situations and are not limited thereto.

[0073] In an exemplary embodiment, when the target communication satellite was performing a task and encountered strong signal interference, resulting in its inability to receive navigation signals, the autonomous operation system of the target communication satellite immediately activated the navigation signal rejection detection mechanism, confirmed the rejection state, and then, in the absence of ground support, implemented the method for determining the satellite operating orbit proposed in the present application.

[0074] Through the embodiments of the present application, by adopting the signal rejection detection and the orbit filtering strategy in the ground support-free state, the technical effect of being able to autonomously determine the satellite operating orbit even under extreme conditions where the satellite loses external navigation signals and ground support is achieved, and the purpose of ensuring the stable operation of the satellite constellation and providing high-precision navigation and positioning services in a complex environment is achieved.

[0075] As an alternative, the above-mentioned orbital integration of the target communication satellite based on the pre-determined conservative force model coefficients, the above-mentioned target satellite position, and the above-mentioned target non-conservative force model coefficients to obtain an integrated orbit includes: fitting the above-mentioned target non-conservative force model coefficients by the least squares method to determine the target non-conservative force fitting coefficients; using the above-mentioned conservative force model coefficients and the above-mentioned target non-conservative force fitting coefficients to determine the dynamic model of the above-mentioned target communication satellite; based on the above-mentioned dynamic model and the above-mentioned target satellite position, solving the state information of the above-mentioned target communication satellite at any moment to obtain the above-mentioned integrated orbit, where the above-mentioned state information includes at least one of the position data, clock data, and velocity data of the above-mentioned target communication satellite.

[0076] Optionally, in the embodiments of the present application, the above-mentioned conservative force model coefficients are used to describe the basic dynamic characteristics of satellite motion during orbital integration, while the above-mentioned target non-conservative force model coefficients cover parameters such as solar radiation pressure, atmospheric drag, and non-spherical earth gravity field that affect the satellite orbit, and can be determined by least squares fitting; and the dynamic model can be understood as a mathematical model that describes the motion state of the satellite based on the conservative force and non-conservative force model coefficients.

[0077] It should be noted that the coefficients of the above-mentioned conservative force and non-conservative force models may vary according to the specific type, operating environment, and mission requirements of the target communication satellite. The fitting algorithm can be other algorithms than the least squares method. The specific form of the target non-conservative force model, the complexity of the dynamic model, and the update frequency and accuracy of the state information can all vary according to the specific implementation situation, and the present application does not make specific limitations on this.

[0078] Through the embodiments of the present application, an orbital integration strategy combining dynamic model prediction based on conservative force and non-conservative force model coefficients with least squares fitting is adopted, achieving the technical effect of accurately determining the orbit of the target communication satellite in a complex environment, and achieving the purpose of effectively controlling the overall rotation error and maintaining high-precision orbit determination ability during the autonomous orbit determination of the satellite constellation.

[0079] As an alternative, the above-mentioned orbital fitting of the above-mentioned initial satellite position, the above-mentioned reference orbit, and the initial non-conservative force model parameters to obtain the target satellite position and the target non-conservative force model coefficients includes: determining the above-mentioned initial non-conservative force model parameters according to the type of the above-mentioned target communication satellite; using the least squares method to fit the above-mentioned initial satellite position, the above-mentioned reference orbit, and the above-mentioned initial non-conservative force model parameters to obtain the above-mentioned target satellite position and the above-mentioned target non-conservative force model coefficients.

[0080] Optionally, in the embodiments of the present application, the above-mentioned initial satellite position refers to the starting position information of the target communication satellite at a certain time point, which can be estimated based on ground station observation data or the measurement results of GNSS receivers on the satellite, and the reference orbit is a set of stable orbital paths pre-calculated according to factors such as satellite type, mission requirements, and the geodynamic environment.

[0081] It should be noted that the types of the above-mentioned target communication satellites may include low-earth orbit satellites, medium-earth orbit satellites, high-earth orbit satellites, etc. For each type of satellite, the selection and weight of the non-conservative force model parameters will be different. Therefore, the determination methods of the initial satellite position and the reference orbit can also vary according to the specific application scenario. For example, GNSS precise ephemeris data, ground radar measurements, or data provided by other satellite constellations can be used. Furthermore, the application of the above-mentioned least squares method, such as parameters like the number of iterations, convergence conditions, and data window size, can also be adjusted according to the implementation details. The present application does not make any limitations in this regard.

[0082] In an exemplary embodiment, it is assumed that the target communication satellite is a low-earth orbit satellite, and the non-conservative force model parameters need to consider the effects of atmospheric drag and solar radiation pressure. First, according to the satellite type and mission requirements of the target communication satellite, a set of initial non-conservative force model parameters are selected, including preliminary estimates of the atmospheric drag coefficient and the solar radiation pressure coefficient. Then, based on these initial non-conservative force model parameters, using the initial satellite position data obtained by the GNSS receiver on the satellite and the reference orbit calculated by the geodynamic model, the least squares method is used to fit them. By continuously adjusting the model parameters, the difference between the fitting result and the observation data is minimized, and the above-mentioned target satellite position and target non-conservative force model coefficients are obtained.

[0083] Through the embodiments of the present application, by adopting the orbital fitting method that combines non-conservative force model parameters customized for satellite types with the least squares method fitting, the technical effect of accurately estimating the satellite orbit and non-conservative force model parameters during the autonomous orbit determination process of the satellite constellation is achieved. The purpose of eliminating rank deficiency, ensuring the stability and accuracy of the autonomous orbit determination algorithm, and effectively suppressing the overall rotation error of the constellation is achieved.

[0084] As an alternative solution, determining the above-mentioned initial non-conservative force model parameters according to the type of the above-mentioned target communication satellite includes: when the type of the above-mentioned target communication satellite is a low-earth orbit type, the above-mentioned initial non-conservative force model parameters include at least one of an atmospheric drag model parameter, a solar radiation pressure model parameter, and a magnetic force influence parameter; when the type of the above-mentioned target communication satellite is a medium-earth orbit type, the above-mentioned initial non-conservative force model parameters include at least one of the above-mentioned solar radiation pressure model parameter, a lunar gravitational perturbation parameter, and a third-body gravitational model parameter; when the type of the above-mentioned target communication satellite is a high-earth orbit type, the above-mentioned initial non-conservative force model parameters include at least one of a terrestrial gravity field model parameter, the above-mentioned lunar gravitational perturbation parameter, and a solar gravitational perturbation parameter.

[0085] Optionally, in the embodiments of the present application, the type of the target communication satellite may be classified according to its operating orbit altitude, and may be specifically divided into a low-earth orbit type (such as LEO, Low Earth Orbit), a medium-earth orbit type (such as MEO, Medium Earth Orbit), and a high-earth orbit type (such as GEO, Geostationary Earth Orbit and Molniya orbit, etc.), and the initial non-conservative force model parameters refer to a series of parameters related to non-conservative forces preset according to the satellite type, which are used to describe the external force influence on the satellite orbit.

[0086] Exemplarily, when the type of the above-mentioned target communication satellite is a low-earth orbit type, rather than a medium-earth orbit or a high-earth orbit type, the initial non-conservative force model parameters may include one or more of an atmospheric drag model parameter, a solar radiation pressure model parameter, and a magnetic force influence parameter.

[0087] In another example, when the type of the above-mentioned target communication satellite is a medium-earth orbit type, rather than a low-earth orbit or a high-earth orbit type, the initial non-conservative force model parameters may include one or more of a solar radiation pressure model parameter, a lunar gravitational perturbation parameter, and a third-body gravitational model parameter.

[0088] In yet another example, when the type of the above-mentioned target communication satellite is a high-earth orbit type, rather than a medium-earth orbit or a low-earth orbit type, the initial non-conservative force model parameters may include one or more of a terrestrial gravity field model parameter, a lunar gravitational perturbation parameter, and a solar gravitational perturbation parameter.

[0089] Generally speaking, in the embodiments of the present application, the initial non-conservative force model parameters are closely related to the type of the target communication satellite, and the initial non-conservative force model parameters required for different types of target communication satellites are not exactly the same.

[0090] As an alternative, the above method further includes: the inter-satellite link data includes at least one of the first geometric distance, the first distance correction term, the first error correction term, and the first observation residual term between the target communication satellite and any one of the other communication satellites when the target communication satellite transmits a signal to any one of the other communication satellites; and at least one of the second geometric distance, the second distance correction term, the second error correction term, and the second observation residual term between any one of the other communication satellites and the target communication satellite when any one of the other communication satellites transmits a signal to the target communication satellite; wherein, the first distance correction term is determined by the product of the speed of light and the clock deviation between the transmission time of the signal transmitted by the target communication satellite and the reception time of the signal received by any one of the other communication satellites; the second distance correction term is determined by the product of the speed of light and the clock deviation between the transmission time of the signal transmitted by any one of the other communication satellites and the reception time of the signal received by the target communication satellite; the first error correction term and the second error correction term are both determined by the satellite antenna phase center deviation, the relativistic effect, and the atmospheric delay; the first observation residual term and the second observation residual term are both determined by the observation noise generated during the acquisition of the inter-satellite link data.

[0091] As an alternative, the above method further includes: determining the first geometric distance according to the position at the transmission time of the signal transmitted by the target communication satellite and the position at the reception time of any one of the other communication satellites; determining the second geometric distance according to the position at the transmission time of the signal transmitted by any one of the other communication satellites and the position at the reception time of the target communication satellite; adjusting the first geometric distance using at least one of the first distance correction term, the first error correction term, and the first observation residual term, and performing a normalization process on the first geometric distance using the clock error of the predicted orbit generated in the above orbit filtering; adjusting the second geometric distance using at least one of the second distance correction term, the second error correction term, and the second observation residual term, and performing a normalization process on the second geometric distance using the clock error of the predicted orbit; performing a first-order Taylor expansion on the first geometric distance and the second geometric distance respectively, retaining the first-order term, to obtain the first linearized geometric distance and the second linearized geometric distance; normalizing the first linearized geometric distance and the second linearized geometric distance to the first target geometric distance and the second target geometric distance at the reference time respectively; and determining the target operating orbit based on the first target geometric distance, the second target geometric distance, and the position error sequence.

[0092] Specifically, Figure 4 is a schematic diagram of two-way observation of the inter-satellite link for an alternative method for determining the satellite operating orbit according to an embodiment of the present application. As Figure 4 shown, the original observation equation of the inter-satellite link can be expressed as:

[0093] 。

[0094] 。

[0095] Among them, and represent the actual signal reception time, and represent the actual signal transmission time, and respectively represent the inter-satellite link observation values of communication satellite A transmitting and communication satellite B receiving, and communication satellite B transmitting and communication satellite A receiving in the same satellite constellation. Assuming that communication satellite A is the above-mentioned target communication satellite, corresponds to the above-mentioned first geometric distance, corresponds to the second geometric distance, corresponds to the above-mentioned first distance correction term, corresponds to the above-mentioned second distance correction term, represents the satellite clock error of communication satellite B at time t1, represents at time (i.e., the time when satellite A transmits the signal, considering the signal transmission delay), the satellite clock error of communication satellite A, represents the satellite clock error of communication satellite A at time t2, represents at time (i.e., the time when satellite B transmits the signal, considering the signal transmission delay), the satellite clock error of communication satellite A, represents the error corrections such as the satellite antenna phase center, relativistic effect, and atmospheric delay when communication satellite A transmits and communication satellite B receives, corresponding to the above-mentioned first error correction term, △ BA represents the error corrections such as the satellite antenna phase center, relativistic effect, and atmospheric delay when communication satellite A receives and communication satellite B transmits, corresponding to the above-mentioned second error correction term, represents the random noise or unmodeled residuals in the observation), c is the speed of light, corresponds to the above-mentioned first observation residual term, corresponds to the above-mentioned second observation residual term.

[0096] For example, r A represents the satellite position of communication satellite A, r B represents the satellite position of communication satellite B, and the geometric distance can be expressed as:

[0097] 。

[0098] 。

[0099] After correcting various errors, the observed values are normalized using the predicted orbit clock error. The sampling interval of the inter-satellite link is 300 seconds, the round-trip link interval is 1.5 seconds, the data processing interval is set to 300 seconds, and the two-way observation data is normalized to the reference time (0 seconds, 300 seconds, etc.). At this time, the difference between the reference time and the actual signal transmission and reception times is no more than 2 seconds.

[0100] Next, the geometric distance is Taylor-expanded according to the reference satellite orbit of the actual signal transmission and reception times, and only the first-order term is retained to obtain the linearized geometric distance as follows:

[0101] 。

[0102] 。

[0103] Among them, represents the geometric distance calculated according to the reference satellite orbit, represents the partial derivative of the geometric distance with respect to the satellite orbit, and respectively represent the reference satellite orbit errors at the actual signal transmission and reception times. Here, can be communication satellite A or communication satellite B.

[0104] Among them, the specific form of the partial derivative of the geometric distance with respect to the satellite orbit is as follows:

[0105] 。

[0106] 。

[0107] 。

[0108] 。

[0109] The satellite orbit in the above linearized geometric distance corresponds to the actual signal transmission and reception times. Here, refers to the prior estimate or theoretical value. Therefore, it can be further normalized to the reference time and normalized as follows:

[0110] 。

[0111] 。

[0112] 。

[0113] 。

[0114] Among them, , where T denotes time, x is the position vector of the satellite in the inertial frame, v is the velocity vector of the satellite in the inertial frame, and p are the dynamic parameters, typically parameters of the solar radiation pressure model, atmospheric drag parameters, empirical force model parameters, etc. and denotes the state transition matrix, which can be obtained by integrating the variational equations. Here, denotes an arbitrary time.

[0115] As an optional solution, the above-mentioned initial satellite position, the above-mentioned reference orbit, and the initial non-conservative force model parameters are used for orbit fitting to obtain the target satellite position and the target non-conservative force model coefficients, including: the above-mentioned initial non-conservative force model parameters are determined by the type of the above-mentioned target communication satellite, where the type of the above-mentioned target communication satellite is any one of the low-earth orbit type, medium-earth orbit type, and high-earth orbit type.

[0116] Optionally, in the embodiments of the present application, the above-mentioned orbit fitting can be performed by means of numerical analysis methods, dynamic models, etc. The above-mentioned non-conservative force model parameters may include but are not limited to solar radiation pressure parameters, atmospheric drag parameters, etc.

[0117] It should be noted that the above-mentioned technical solution of the present application is applicable to communication satellites of different types of orbits. Whether it is a low-earth orbit, medium-earth orbit, or high-earth orbit, according to the orbital environment of the satellite and the influence degree of non-conservative forces, the above-mentioned target non-conservative force model coefficients can be set specifically to improve the orbit fitting accuracy. The present application does not make any limitations in this regard.

[0118] In an exemplary embodiment, taking the application scenario of low-earth orbit satellites as an example, during the autonomous operation of a satellite constellation, when using inter-satellite link data for orbit determination, the initial non-conservative force model parameters, including the solar radiation pressure model, atmospheric drag model, etc., can be set in advance according to the characteristics of low-earth orbit satellites. Furthermore, the parameters are continuously optimized through an orbit fitting algorithm to improve the accuracy and stability of orbit determination.

[0119] Through the embodiments of the present application, orbit fitting is performed using the initial non-conservative force model parameters set according to the satellite type, realizing high-precision orbit determination of communication satellites in the case of navigation signal rejection or no ground support, achieving the purpose of autonomous orbit determination and effectively controlling the overall rotation error of the constellation.

[0120] As an alternative solution, when the above-mentioned target communication satellite is unable to receive the above-mentioned navigation signal, using the inter-satellite link data for orbit filtering to determine the above-mentioned target operating orbit includes: when the above-mentioned target communication satellite is unable to receive the above-mentioned navigation signal, obtaining the relative distance and relative clock difference between the above-mentioned target communication satellite and other communication satellites indicated by the above-mentioned inter-satellite link data, where the above-mentioned target communication satellite and the above-mentioned other communication satellites are in the same satellite constellation; performing the above-mentioned orbit filtering based on the above-mentioned relative distance and the above-mentioned relative clock difference to determine the above-mentioned target operating orbit, where during the process of the above-mentioned orbit filtering, the initial orbit state quantity is determined by the above-mentioned integrated orbit, and the initial process noise covariance is determined by the above-mentioned process noise covariance information.

[0121] Optionally, in the embodiments of the present application, the above-mentioned inter-satellite link data refers to the relative distance and relative clock difference data obtained through two-way link communication between the target communication satellite and other communication satellites within the same constellation, including but not limited to the mutual distance measurement values and time synchronization differences exchanged in real time by on-board devices, and can be used for autonomous orbit filtering when the target satellite loses the navigation signal. The on-board devices herein refer to communication antennas, ranging devices, time synchronization modules, etc. on the target communication satellite.

[0122] It should be noted that the embodiments of the present application cover the implementation of various orbit filtering algorithms, such as extended Kalman filtering, unscented Kalman filtering, or particle filtering, etc., which can all be used to process the inter-satellite link data to determine the target operating orbit, and the present application does not make any limitations in this regard.

[0123] In an exemplary embodiment, taking the application scenario of medium-earth orbit communication satellites as an example, when the target communication satellite is temporarily unable to receive the navigation signal due to entering a certain area, the relative distance and relative clock difference between the target satellite and other satellites in the constellation can be obtained through the inter-satellite link. Using the relative distance and relative clock difference, combined with the initial orbit state quantity and process noise covariance obtained from the previous orbit integration, real-time orbit update is performed through the extended Kalman filtering algorithm, ensuring the accuracy and stability of the target satellite's operating orbit during the interruption of the navigation signal.

[0124] Through the embodiments of the present application, using the inter-satellite link data for orbit filtering realizes the autonomous determination of the operating orbit of the target communication satellite when the navigation signal cannot be received, achieving the purpose of maintaining the overall operating performance of the satellite constellation and providing continuous navigation services.

[0125] In an exemplary embodiment, the embodiments of the present application can be applied to the field of satellite navigation technology to implement a method for controlling the overall rotation error of autonomous orbit determination of a satellite constellation. Among them, the satellite orbit error is a key element of the satellite spatial reference, and its accuracy directly determines the service performance of the satellite. Moreover, high-precision navigation enhancement services place extremely high centimeter-level accuracy requirements on the orbit information of low-earth orbit satellites.

[0126] In the prior art, mainly through satellite-to-satellite cooperative control and mutual acquisition of status information, it is possible to achieve autonomous determination of the orbits of low-earth orbit constellations based on inter-satellite links. Generally, the autonomous orbit determination of the constellation is carried out using inter-satellite distance observations. However, inter-satellite ranging can only ensure the accuracy of the relative positions of the satellites in the constellation and cannot overcome the overall rotation and drift of the entire constellation. The overall rotation error of the constellation will be directly reflected in the accuracy of constellation orbit determination. The longer the time, the greater the orbit determination error, posing a serious risk of degradation to high-precision navigation enhancement services.

[0127] Therefore, the embodiments of the present application can implement a method for controlling the overall rotation error of autonomous orbit determination of a satellite constellation suitable for on-board processing, effectively suppressing the error of the overall rotation of the constellation in the case of GNSS denial or lack of ground support for satellites, significantly improving the autonomous solution accuracy of the orbits of low-earth orbit constellations, meeting the requirements for real-time autonomous determination and robust service of the entire network of low-earth orbit constellations in future confrontation scenarios, so as to enhance the autonomous operation ability and improve the overall system operation and maintenance level.

[0128] Exemplarily, the embodiments of the present application can introduce prior orbit parameters to constrain the spatial rotation of the satellite constellation. In other words, in the orbit reference, the non-conservative force coefficients are constrained, and the overall constraint compensation algorithm is adopted. After estimating the constellation rotation error of the filtered solution orbit relative to the reference orbit, the orbit parameters obtained by autonomous orbit determination are directly corrected, thereby realizing the suppression of the overall rotation error of the constellation and enhancing the autonomous operation ability of the satellite network, including but not limited to:

[0129] S1. Use the target communication satellite to collect the spatio-temporal information (including but not limited to the position, clock error, etc. of GNSS satellites) of GNSS satellites (the above-mentioned navigation satellites) in the current autonomous processing cycle, and calculate the initial satellite position and reference orbit of the communication satellite at the current initial moment. For example, use the target communication satellite to collect the GNSS precise ephemeris data file in the current autonomous processing cycle, and calculate the initial satellite position and reference orbit.

[0130] S2. Perform orbit fitting (including but not limited to algorithms such as the least squares method, etc.) on the initial satellite position, reference orbit, and related non-conservative forces (atmospheric drag, earth albedo radiation pressure, solar radiation pressure, etc., which are related to the satellite type, such as low-earth orbit, high-earth orbit, etc.) to obtain a more accurate satellite position, denoted as the target satellite position and the non-conservative force model coefficient.

[0131] S3. Combine the conservative force model coefficients and the target non-conservative force fitting coefficients obtained by fitting the non-conservative force model coefficients, and use numerical integration methods to perform orbit integration on the target communication satellite. It can be understood that the current position of the target satellite is used to predict the satellite position at the next moment in the future, and the satellite positions at multiple moments are recursively obtained to obtain the integrated orbit.

[0132] S4. Compare the integrated orbit with the reference orbit to obtain the position errors of the target communication satellite at different future moments, that is, the position error sequence.

[0133] S5. When in the GNSS denial or no ground support state, use the position error sequence and inter-satellite link data for orbit filtering (including but not limited to Kalman filtering, etc.) to determine the target operating orbit of the communication satellite. That is, use the dynamic model parameters fitted from the previous period's orbit as the prior constraint information for the autonomous orbit determination algorithm, adjust the weight of the predicted orbit in the orbit filtering, and suppress the accumulation of the overall rotation error of the low-earth orbit constellation.

[0134] Through the embodiments of the present application, a technical solution for suppressing the space rotation of a low-earth orbit constellation by introducing prior orbit parameter constraints is introduced. The dynamic model parameters fitted from the previous period's orbit are used as the prior constraint information for the autonomous orbit determination algorithm to eliminate the rank deficiency of the autonomous orbit determination solution, ensure that the autonomous orbit determination algorithm can be correctly and stably solved, thereby effectively suppressing the accumulation of the overall rotation error of the low-earth orbit constellation and significantly improving the autonomous solution accuracy of the orbit clock error of the low-earth orbit constellation.

[0135] It can be understood that in the specific implementation manners of the present application, data related to user information, etc. are involved. When the above embodiments of the present application are applied to specific products or technologies, user permission or consent needs to be obtained, and the collection, use, and processing of relevant data need to comply with the relevant laws, regulations, and standards of relevant countries and regions.

[0136] It should be noted that for the foregoing method embodiments, for the sake of simple description, they are all expressed as a series of action combinations. However, those skilled in the art should know that the present application is not limited by the described action sequence, because according to the present application, certain steps can be performed in other sequences or simultaneously. Secondly, those skilled in the art should also know that the embodiments described in the specification are all preferred embodiments, and the actions and modules involved are not necessarily essential to the present application.

[0137] According to another aspect of the embodiments of the present application, there is also provided a device for determining a satellite operating orbit for implementing the method for determining a satellite operating orbit described above. As Figure 5 shown, the device includes:

[0138] An acquisition module 502, configured to acquire an initial satellite position of a target communication satellite when the target communication satellite can receive a navigation signal, and determine a reference orbit of the target communication satellite based on the spatio-temporal information of a navigation satellite, where the spatio-temporal information includes the position and clock offset of the navigation satellite, and the reference orbit is an operating orbit determined by the target communication satellite during the autonomous orbit determination process;

[0139] A fitting module 504, configured to perform orbit fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters to obtain a target satellite position and target non-conservative force model coefficients;

[0140] An integration module 506, configured to perform orbit integration on the target communication satellite according to pre-determined conservative force model coefficients, the target satellite position, and the target non-conservative force model coefficients to determine an integrated orbit of the target communication satellite;

[0141] A comparison module 508, configured to compare the integrated orbit and the reference orbit to determine a position error sequence of the target communication satellite, and determine process noise covariance information at different integration durations in the integrated orbit based on the position error sequence, where the position error sequence is used to indicate the change of the position deviation of the target communication satellite relative to the target operating orbit over time during the autonomous orbit determination process;

[0142] A determination module 510, configured to perform orbit filtering using inter-satellite link data to determine the target operating orbit when the target communication satellite cannot receive the navigation signal.

[0143] As an optional solution, the above device is used to compare the integrated orbit and the reference orbit in the following manner to determine a position error sequence of the target communication satellite, and determine process noise covariance information at different integration durations in the integrated orbit based on the position error sequence: acquire an integrated orbit point corresponding to the integrated orbit at any moment, and acquire a reference orbit point corresponding to the reference orbit at any moment; determine error data corresponding to the target communication satellite at any moment according to the integrated orbit point and the reference orbit point, where the error data includes at least one of position error data and clock offset error data; generate a position error sequence according to the error data.

[0144] As an alternative solution, the above-mentioned device is used to perform orbit filtering using inter-satellite link data and determine the target operating orbit in the case where the target communication satellite cannot receive navigation signals in the following manner: when the target communication satellite cannot receive navigation signals, determine the initial operating state of the predicted orbit according to the initial orbit state variables and the initial process noise covariance information; use the inter-satellite link data to correct the predicted orbit according to the Kalman filtering algorithm to obtain the target operating orbit; wherein, the inter-satellite link data is used to determine the observation residual in the process of orbit filtering, and the observation residual represents the difference between the inter-satellite ranging observation value indicated by the inter-satellite link data and the expected inter-satellite ranging value of the predicted orbit.

[0145] As an alternative solution, the above-mentioned device is further used to: determine that the target communication satellite cannot receive navigation signals when the navigation signal is denied; determine that the target communication satellite cannot receive navigation signals when the target communication satellite is in a state without ground support.

[0146] As an alternative solution, the above-mentioned device is used to perform orbit integration on the target communication satellite according to the pre-determined conservative force model coefficients, the target satellite position, and the target non-conservative force model coefficients in the following manner to obtain the integrated orbit: use the least squares method to fit the target non-conservative force model coefficients to determine the target non-conservative force fitting coefficients; use the conservative force model coefficients and the target non-conservative force fitting coefficients to determine the dynamic model of the target communication satellite; based on the dynamic model and the target satellite position, solve the state information of the target communication satellite at any moment to obtain the integrated orbit, where the state information includes at least one of the position data, clock data, and velocity data of the target communication satellite.

[0147] As an alternative solution, the above-mentioned device is used to perform orbit fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters in the following manner to obtain the target satellite position and the target non-conservative force model coefficients: determine the initial non-conservative force model parameters according to the type of the target communication satellite; use the least squares method to fit the initial satellite position, the reference orbit, and the initial non-conservative force model parameters to obtain the target satellite position and the target non-conservative force model coefficients.

[0148] As an alternative solution, the above-mentioned device is used to determine the initial non-conservative force model parameters according to the type of the target communication satellite in the following manner: when the type of the target communication satellite is a low-earth orbit type, the initial non-conservative force model parameters include at least one of the atmospheric drag model parameters, the solar radiation pressure model parameters, and the magnetic force influence parameters; when the type of the target communication satellite is a medium-earth orbit type, the initial non-conservative force model parameters include at least one of the solar radiation pressure model parameters, the lunar gravitational perturbation parameters, and the third-body gravitational model parameters; when the type of the target communication satellite is a high-earth orbit type, the initial non-conservative force model parameters include at least one of the earth gravity field model parameters, the lunar gravitational perturbation parameters, and the solar gravitational perturbation parameters.

[0149] As an alternative solution, the above-mentioned device is also used for: when the inter-satellite link data includes at least one of the first geometric distance, the first distance correction term, the first error correction term, and the first observation residual term between the target communication satellite and any one of the other communication satellites when the target communication satellite transmits a signal to any one of the other communication satellites; and at least one of the second geometric distance, the second distance correction term, the second error correction term, and the second observation residual term between any one of the other communication satellites and the target communication satellite when any one of the other communication satellites transmits a signal to the target communication satellite; wherein, the first distance correction term is determined by the product of the speed of light and the clock deviation between the transmission time of the signal transmitted by the target communication satellite and the reception time of the signal received by any one of the other communication satellites; the second distance correction term is determined by the product of the speed of light and the clock deviation between the transmission time of the signal transmitted by any one of the other communication satellites and the reception time of the signal received by the target communication satellite; the first error correction term and the second error correction term are both determined by the satellite antenna phase center deviation, the relativistic effect, and the atmospheric delay; the first observation residual term and the second observation residual term are both determined by the observation noise generated during the acquisition of the inter-satellite link data.

[0150] As an alternative, the above device is further configured to: determine a first geometric distance based on the position of the target communication satellite at the signal transmission time and the position of any communication satellite at the signal reception time; determine a second geometric distance based on the position of any communication satellite at the signal transmission time and the position of the target communication satellite at the signal reception time; adjust the first geometric distance using at least one of a first distance correction term, a first error correction term, and a first observation residual term, and normalize the first geometric distance using the clock error of the predicted orbit generated in the orbit filtering; adjust the second geometric distance using at least one of a second distance correction term, a second error correction term, and a second observation residual term, and normalize the second geometric distance using the clock error of the predicted orbit; perform a first-order Taylor expansion on the first geometric distance and the second geometric distance respectively, retain the first-order terms, to obtain a first linearized geometric distance and a second linearized geometric distance; normalize the first linearized geometric distance and the second linearized geometric distance to a first target geometric distance and a second target geometric distance at a reference time respectively; determine the target operating orbit based on the first target geometric distance, the second target geometric distance, and the position error sequence.

[0151] As an alternative, the device is configured to perform orbit fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters in the following manner to obtain the target satellite position and the target non-conservative force model coefficients: the initial non-conservative force model parameters are determined by the type of the target communication satellite, where the type of the target communication satellite is any one of a low-earth orbit type, a medium-earth orbit type, and a high-earth orbit type.

[0152] As an alternative, the device is configured to perform orbit filtering using inter-satellite link data to determine the target operating orbit in the case where the target communication satellite cannot receive navigation signals in the following manner: in the case where the target communication satellite cannot receive navigation signals, obtain the relative distance and relative clock difference between the target communication satellite and other communication satellites indicated by the inter-satellite link data, where the target communication satellite and other communication satellites are in the same satellite constellation; perform orbit filtering based on the relative distance and relative clock difference to determine the target operating orbit, where in the process of orbit filtering, the initial orbit state quantity is determined by the integrated orbit, and the initial process noise covariance is determined by the process noise covariance information.

[0153] In the embodiments of the present application, the term "module" or "unit" refers to a computer program with a predetermined function or a part of a computer program, which works together with other related parts to achieve a predetermined goal, and can be implemented in whole or in part by using software, hardware (such as a processing circuit or a memory), or a combination thereof. Similarly, one processor (or multiple processors or memories) can be used to implement one or more modules or units. In addition, each module or unit can be a part of the overall module or unit that includes the function of the module or unit.

[0154] Regarding the device in the above embodiments, the specific manners in which each module performs operations have been described in detail in the embodiments related to the method, and will not be elaborated herein.

[0155] According to one aspect of the present application, there is provided a computer program product, which includes a computer program.

[0156] The serial numbers of the embodiments of the present application above are only for description and do not represent the superiority or inferiority of the embodiments.

[0157] Figure 6 A block diagram of a computer system of an electronic device for implementing the embodiments of the present application is schematically shown.

[0158] It should be noted that Figure 6 The computer system 600 of the electronic device shown is only an example and should not impose any limitations on the functions and usage scope of the embodiments of the present application.

[0159] As Figure 6 shown, the computer system 600 includes a central processing unit 601 (CPU), which can perform various appropriate actions and processes according to a program stored in a read-only memory 602 (ROM) or a program loaded from a storage section 608 into a random access memory 603 (RAM). In the random access memory 603, various programs and data required for system operations are also stored. The central processing unit 601, the read-only memory 602, and the random access memory 603 are connected to each other via a bus 604. An input / output interface 605 (Input / Output interface, i.e., I / O interface) is also connected to the bus 604.

[0160] The following components are connected to the input / output interface 605: an input section 606 including a keyboard, a mouse, etc.; an output section 607 including, for example, a cathode ray tube (CRT), a liquid crystal display (LCD), etc., and a speaker, etc.; a storage section 608 including a hard disk, etc.; and a communication section 609 including a network interface card such as a local area network card, a modem, etc. The communication section 609 performs communication processing via a network such as the Internet. A drive 610 is also connected to the input / output interface 605 as needed. A removable medium 611, such as a magnetic disk, an optical disk, a magneto-optical disk, a semiconductor memory, etc., is installed on the drive 610 as needed so that a computer program read from it can be installed into the storage section 608 as needed.

[0161] In particular, according to the embodiments of the present application, the processes described in each method flow chart can be implemented as computer software programs. For example, the embodiments of the present application include a computer program product, which includes a computer program carried on a computer-readable medium, and the computer program contains program codes for executing the methods shown in the flow charts. In such an embodiment, the computer program can be downloaded and installed from the network through the communication part 609, and / or installed from the removable medium 611. When the computer program is executed by the central processing unit 601, various functions defined in the system of the present application are executed.

[0162] In such an embodiment, the computer program can be downloaded and installed from the network through the communication part 609, and / or installed from the removable medium 611. When the computer program is executed by the central processing unit 601, various functions provided by the embodiments of the present application are executed.

[0163] According to another aspect of the embodiments of the present application, an electronic device for implementing the above-mentioned method for determining the satellite operation orbit is further provided. The electronic device may be Figure 1 the terminal device or server shown. This embodiment takes the electronic device as the terminal device as an example for illustration. As Figure 7 shown, the electronic device includes a memory 702 and a processor 704. A computer program is stored in the memory 702, and the processor 704 is configured to execute the steps in any one of the above method embodiments through the computer program.

[0164] Optionally, in this embodiment, the above-mentioned electronic device may be at least one network device among multiple network devices in a computer network.

[0165] Optionally, in this embodiment, the above-mentioned processor may be configured to execute the methods in the embodiments of the present application through the computer program.

[0166] Optionally, those of ordinary skill in the art can understand that Figure 7 the structure shown is only schematic, Figure 7 and it does not limit the structure of the above-mentioned electronic device. For example, the electronic device may further include more or fewer components (such as a network interface, etc.) than those shown in Figure 7 , or have a different configuration from that shown in Figure 7 .

[0167] Among them, the memory 702 can be used to store software programs and modules, such as the program instructions / modules corresponding to the method and device for determining the satellite operation orbit in the embodiments of the present application. The processor 704 executes various functional applications and data processing by running the software programs and modules stored in the memory 702, that is, implements the above-mentioned method for determining the satellite operation orbit. The memory 702 can include a high-speed random access memory, and can also include a non-volatile memory, such as one or more magnetic storage devices, flash memory, or other non-volatile solid-state memories. In some instances, the memory 702 can further include a memory remotely disposed relative to the processor 704, and these remote memories can be connected to the terminal through a network. Examples of the above network include but are not limited to the Internet, enterprise intranet, local area network, mobile communication network, and their combinations. Among them, the memory 702 can specifically but not limitedly be used to store information carried by navigation signals, communication signals, and other information. As an example, as Figure 7 shown, the above memory 702 can but not limitedly include the acquisition module 502, fitting module 504, integration module 506, comparison module 508, and determination module 510 in the above device for determining the satellite operation orbit. In addition, it can also include but not limited to other module units in the above device for determining the satellite operation orbit, which will not be elaborated in this example.

[0168] Optionally, the above transmission device 706 is used to receive or send data via a network. Specific examples of the above network can include a wired network and a wireless network. In one instance, the transmission device 706 includes a network adapter (Network Interface Controller, NIC), which can be connected to other network devices and routers through a network cable, thereby enabling communication with the Internet or local area network. In one instance, the transmission device 706 is a radio frequency (Radio Frequency, RF) module, which is used to communicate with the Internet wirelessly.

[0169] In addition, the above electronic device further includes: a display 708 for displaying the above positioning information; and a connection bus 710 for connecting each module component in the above electronic device.

[0170] In other embodiments, the above terminal device or server can be a node in a distributed system. Among them, the distributed system can be a blockchain system, and the blockchain system can be a distributed system formed by connecting the multiple nodes through network communication. Among them, the nodes can form a peer-to-peer network, and any form of computing device, such as a server, terminal, and other electronic devices, can become a node in the blockchain system by joining the peer-to-peer network.

[0171] According to one aspect of the present application, a computer-readable storage medium is provided. A processor of an electronic device reads the computer instructions from the computer-readable storage medium, and the processor executes the computer instructions, so that the electronic device executes the method for determining the satellite orbit provided in various alternative implementations of the above-mentioned aspect of determining the satellite orbit.

[0172] Optionally, in this embodiment, the above computer-readable storage medium may be configured to store instructions for executing the methods in the various embodiments of the present application.

[0173] Optionally, in this embodiment, those of ordinary skill in the art can understand that all or part of the steps in the various methods of the above embodiments can be completed by a program instructing the relevant hardware of the terminal device. The program can be stored in a computer-readable storage medium, and the storage medium may include: a flash drive, a read-only memory (ROM), a random access memory (RAM), a magnetic disk or an optical disc, etc.

[0174] The serial numbers of the above embodiments of the present application are only for description and do not represent the advantages or disadvantages of the embodiments.

[0175] If the integrated unit in the above embodiment is implemented in the form of a software functional unit and sold or used as an independent product, it can be stored in the above computer-readable storage medium. Based on such an understanding, the technical solution of the present application, in essence, or the part that contributes to the prior art, or all or part of the technical solution, can be embodied in the form of a software product. The computer software product is stored in a storage medium and includes several instructions for causing one or more electronic devices to execute all or part of the steps of the methods described in the various embodiments of the present application.

[0176] In the above embodiments of the present application, the descriptions of the various embodiments have their own emphases. For the parts not detailed in a certain embodiment, reference may be made to the relevant descriptions of other embodiments.

[0177] In the several embodiments provided by the present application, it should be understood that the disclosed application program can be implemented in other ways. Among them, the device embodiments described above are only illustrative. For example, the division of the units is only a logical function division. In actual implementation, there may be other division methods. For example, multiple units or components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the displayed or discussed mutual coupling or direct coupling or communication connection can be through some interfaces, and the indirect coupling or communication connection of the units or modules can be in an electrical or other form.

[0178] The unit described as a separation component may or may not be physically separated. The component shown as a unit may or may not be a physical unit, that is, it may be located in one place or may be distributed over multiple network units. Some or all of the units can be selected according to actual needs to achieve the purpose of the solution of this embodiment.

[0179] In addition, each functional unit in various embodiments of the present application may be integrated into a processing unit, may exist independently as individual physical units, or two or more units may be integrated into one unit. The above-mentioned integrated unit may be implemented in the form of hardware or in the form of a software functional unit.

[0180] The above are only the preferred embodiments of the present application. It should be noted that for those of ordinary skill in the art, without departing from the principle of the present application, several improvements and refinements can be made, and these improvements and refinements should also be regarded as the protection scope of the present application.

Claims

1. A method for determining a satellite operating orbit, characterized in that, Including: When the target communication satellite can receive navigation signals, obtain the initial satellite position of the target communication satellite, and determine the reference orbit of the target communication satellite based on the spatio-temporal information of the navigation satellite, where the spatio-temporal information includes the position and clock offset of the navigation satellite, and the reference orbit is the operating orbit determined by the target communication satellite during the autonomous orbit determination process; Perform orbit fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters to obtain the target satellite position and the target non-conservative force model coefficients; Perform orbit integration on the target communication satellite according to the pre-determined conservative force model coefficients, the target satellite position, and the target non-conservative force model coefficients to determine the integrated orbit of the target communication satellite; Compare the integrated orbit with the reference orbit to determine the position error sequence of the target communication satellite, and determine the process noise covariance information at different integration durations in the integrated orbit based on the position error sequence, where the position error sequence is used to indicate the change of the position deviation of the target communication satellite relative to the target operating orbit over time during the autonomous orbit determination process; When the target communication satellite cannot receive the navigation signal, use the inter-satellite link data to perform orbit filtering to determine the target operating orbit.

2. The method according to claim 1, wherein The comparing the integrated orbit with the reference orbit to determine the position error sequence of the target communication satellite, and determining the process noise covariance information at different integration durations in the integrated orbit based on the position error sequence includes: Obtain the integrated orbit point corresponding to any moment of the integrated orbit, and obtain the reference orbit point corresponding to the any moment of the reference orbit; Determine the error data corresponding to the target communication satellite at the any moment according to the integrated orbit point and the reference orbit point, where the error data includes at least one of position error data and clock offset error data; Generate the position error sequence according to the error data.

3. The method according to claim 1, wherein The when the target communication satellite cannot receive the navigation signal, using the inter-satellite link data to perform orbit filtering to determine the target operating orbit includes: When the target communication satellite cannot receive the navigation signal, determine the initial operating state of the predicted orbit according to the initial orbit state quantity and the initial process noise covariance information; Use the inter-satellite link data to correct the predicted orbit according to the Kalman filtering algorithm to obtain the target operating orbit; Wherein, the inter-satellite link data is used to determine the observation residual in the process of the orbit filtering, and the observation residual represents the difference between the inter-satellite ranging observation value indicated by the inter-satellite link data and the inter-satellite ranging expected value of the predicted orbit.

4. The method according to claim 1, wherein The method further includes: When the navigation signal is denied, determine that the target communication satellite cannot receive the navigation signal; When the target communication satellite is in a state without ground support, determine that the target communication satellite cannot receive the navigation signal.

5. The method according to claim 1, wherein Performing orbital integration on the target communication satellite according to the pre-determined conservative force model coefficients, the target satellite position, and the target non-conservative force model coefficients to obtain an integrated orbit, including: Fitting the target non-conservative force model coefficients using the least squares method to determine the target non-conservative force fitting coefficients; Determining the dynamic model of the target communication satellite using the conservative force model coefficients and the target non-conservative force fitting coefficients; Solving the state information of the target communication satellite at any moment based on the dynamic model and the target satellite position to obtain the integrated orbit, where the state information includes at least one of the position data, clock data, and velocity data of the target communication satellite.

6. The method according to claim 1, characterized in that, Performing orbital fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters to obtain the target satellite position and the target non-conservative force model coefficients, including: Determining the initial non-conservative force model parameters according to the type of the target communication satellite; Performing fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters using the least squares method to obtain the target satellite position and the target non-conservative force model coefficients.

7. The method according to claim 6, wherein The determining the initial non-conservative force model parameters according to the type of the target communication satellite includes: When the type of the target communication satellite is a low-earth orbit type, the initial non-conservative force model parameters include at least one of the atmospheric drag model parameters, the solar radiation pressure model parameters, and the magnetic force influence parameters; When the type of the target communication satellite is a medium-earth orbit type, the initial non-conservative force model parameters include at least one of the solar radiation pressure model parameters, the lunar gravitational perturbation parameters, and the third-body gravitational model parameters; When the type of the target communication satellite is a geosynchronous orbit type, the initial non-conservative force model parameters include at least one of the earth gravity field model parameters, the lunar gravitational perturbation parameters, and the solar gravitational perturbation parameters.

8. The method according to claim 1, characterized in that The method further includes: The inter-satellite link data includes at least one of the first geometric distance, the first distance correction term, the first error correction term, and the first observation residual term between the target communication satellite and any one of the other communication satellites when the target communication satellite transmits a signal to the other communication satellite, where the target communication satellite and the other communication satellites are in the same satellite constellation; and At least one of the second geometric distance, the second distance correction term, the second error correction term, and the second observation residual term between any one of the other communication satellites and the target communication satellite when the other communication satellite transmits a signal to the target communication satellite; Wherein, the first distance correction term is determined by the product of the speed of light and the clock deviation between the transmission time of the signal transmitted by the target communication satellite and the reception time of the signal received by the other communication satellite; The second distance correction term is determined by the product of the speed of light and the clock deviation between the transmission time of the signal transmitted by the other communication satellite and the reception time of the signal received by the target communication satellite. The first error correction term and the second error correction term are both determined by the satellite antenna phase center deviation, the relativistic effect, and the atmospheric delay; The first observation residual term and the second observation residual term are both determined by the observation noise generated during the acquisition of the inter-satellite link data.

9. The method according to claim 8, wherein The method further includes: Determining the first geometric distance according to the position of the target communication satellite at the transmission time of the transmitted signal and the position of any communication satellite at the reception time; determining the second geometric distance according to the position of any communication satellite at the transmission time of the transmitted signal and the position of the target communication satellite at the reception time; Adjusting the first geometric distance using at least one of the first distance correction term, the first error correction term, and the first observation residual term, and normalizing the first geometric distance using the clock error of the predicted orbit generated in the orbit filtering; adjusting the second geometric distance using at least one of the second distance correction term, the second error correction term, and the second observation residual term, and normalizing the second geometric distance using the clock error of the predicted orbit; Performing first-order Taylor expansion on the first geometric distance and the second geometric distance respectively, retaining the first-order terms, to obtain a first linearized geometric distance and a second linearized geometric distance; Normalizing the first linearized geometric distance and the second linearized geometric distance to a first target geometric distance and a second target geometric distance at a reference time respectively; Determining the target operating orbit based on the first target geometric distance, the second target geometric distance, and the position error sequence.

10. The method according to claim 1, characterized in that, The performing orbit fitting on the initial satellite position, the reference orbit, and the initial non-conservative force model parameters to obtain the target satellite position and the target non-conservative force model coefficients includes: The initial non-conservative force model parameters are determined by the type of the target communication satellite, where the type of the target communication satellite is any one of a low-earth orbit type, a medium-earth orbit type, and a high-earth orbit type.

11. The method according to claim 1, wherein The determining the target operating orbit using the inter-satellite link data when the target communication satellite cannot receive the navigation signal includes: When the target communication satellite cannot receive the navigation signal, acquiring the relative distance and relative clock difference between the target communication satellite and other communication satellites indicated by the inter-satellite link data, where the target communication satellite and the other communication satellites are in the same satellite constellation; Performing the orbit filtering based on the relative distance and the relative clock difference to determine the target operating orbit, where during the orbit filtering, the initial orbit state quantity is determined by the integrated orbit, and the initial process noise covariance is determined by the process noise covariance information.

12. A device for determining a satellite's operating orbit, characterized in that, including: An acquisition module, configured to acquire an initial satellite position of the target communication satellite when the target communication satellite can receive a navigation signal, and determine a reference orbit of the target communication satellite based on spatio-temporal information of navigation satellites, where the spatio-temporal information includes positions and clock biases of the navigation satellites, and the reference orbit is an operating orbit determined by the target communication satellite during the autonomous orbit determination process; A fitting module, configured to perform orbit fitting on the initial satellite position, the reference orbit, and initial non-conservative force model parameters to obtain a target satellite position and target non-conservative force model coefficients; An integration module, configured to perform orbit integration on the target communication satellite according to pre-determined conservative force model coefficients, the target satellite position, and the target non-conservative force model coefficients to determine an integrated orbit of the target communication satellite; A comparison module, configured to compare the integrated orbit and the reference orbit to determine a position error sequence of the target communication satellite, and determine process noise covariance information at different integration durations in the integrated orbit based on the position error sequence, where the position error sequence is used to indicate the change of the position deviation of the target communication satellite relative to the target operating orbit over time during the autonomous orbit determination process; A determination module, configured to perform orbit filtering using inter-satellite link data to determine the target operating orbit when the target communication satellite cannot receive the navigation signal.

13. A computer-readable storage medium, characterized in that, The computer-readable storage medium includes a stored computer program, where the computer program, when being run by an electronic device, executes the method according to any one of claims 1 to 11.

14. A computer program product, comprising a computer program, characterized in that, When being executed by a processor, the computer program implements the steps of the method according to any one of claims 1 to 11.

15. An electronic device, comprising a memory and a processor, characterized in that, A computer program is stored in the memory, and the processor is configured to execute the method according to any one of claims 1 to 11 through the computer program.

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