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

Abstract

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

Description

TECHNICAL FIELD

[0001] The present application relates to the field of computers, in particular to a satellite running orbit determination method and device, a storage medium and an electronic device. BACKGROUND

[0002] In the field of satellite navigation technology, accurate orbit control of a satellite constellation is the key to achieving high-precision positioning services. However, in related technologies, autonomous orbit determination is only dependent on inter-satellite link observation data, which cannot avoid the overall shift and rotation error of the satellite constellation, causing each satellite to deviate from its predetermined orbit. Therefore, the related technologies have the technical problem that the satellites cannot run according to the correct orbit due to the overall shift and rotation of the satellite constellation.

[0003] At present, there is no effective solution to the above problems. SUMMARY

[0004] The satellite running orbit determination method and device, the storage medium and the electronic device provided by the embodiments of the present application can at least solve the technical problem that the satellites cannot run according to the correct orbit due to the overall shift and rotation of the satellite constellation.

[0005] According to an aspect of an embodiment of the present application, a satellite running orbit determination method is provided, including: in a case where 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 space-time information of a navigation satellite, wherein the space-time information includes a position and a clock error of the navigation satellite, and the reference orbit is a running orbit determined by the target communication satellite in an 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 and the reference orbit to determine a position error sequence of the target communication satellite, and determining process noise covariance information under different integration time lengths in the integrated orbit based on the position error sequence, wherein the position error sequence is used to indicate the change of the position deviation of the target communication satellite relative to a target running orbit over time in the autonomous orbit determination process; in a case where the target communication satellite cannot receive the navigation signal, performing orbit filtering using inter-satellite link data to determine the target running orbit.

[0006] According to another aspect of the embodiments of the present application, a satellite orbit determination apparatus is also provided, comprising: an acquisition module configured to acquire an initial satellite position of a target communication satellite when the target communication satellite is capable of receiving a navigation signal, and determine a reference orbit of the target communication satellite based on space-time information of a navigation satellite, wherein the space-time information comprises a position and a clock error of the navigation satellite, and the reference orbit is an orbit determined by the target communication satellite in an 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 predetermined 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, determine a position error sequence of the target communication satellite, and determine process noise covariance information under different integration time lengths in the integrated orbit based on the position error sequence, wherein the position error sequence is used to indicate a change of a position deviation of the target communication satellite relative to a target orbit over time in the autonomous orbit determination process; and a determination module configured to perform orbit filtering using inter-satellite link data to determine the target orbit when the target communication satellite is incapable of receiving the navigation signal.

[0007] Optionally, the apparatus is configured to compare the integrated orbit and the reference orbit, determine the position error sequence of the target communication satellite, and determine the process noise covariance information under different integration time lengths in the integrated orbit based on the position error sequence by: acquiring an integrated orbit point corresponding to any time of the integrated orbit, and acquiring a reference orbit point corresponding to the any time of the reference orbit; determining error data corresponding to the any time of the target communication satellite according to the integrated orbit point and the reference orbit point, wherein the error data comprises at least one of position error data and clock error data; and generating the position error sequence according to the error data.

[0008] Optionally, the apparatus is configured to determine the target operational orbit using inter-satellite link data for orbit filtering in case that the target communication satellite is unable to receive the navigation signal, by: determining an initial operational state of a predicted orbit according to the initial orbit state quantity and the initial process noise covariance information in case that the target communication satellite is unable to receive the navigation signal; correcting the predicted orbit according to the inter-satellite link data in a Kalman filtering algorithm to obtain the target operational orbit; wherein the inter-satellite link data is used to determine an observation residual in the orbit filtering, and the observation residual represents a difference between an inter-satellite ranging observation value indicated by the inter-satellite link data and an inter-satellite ranging expected value of the predicted orbit.

[0009] Optionally, the apparatus is further configured to: determine that the target communication satellite is unable to receive the navigation signal in case of navigation signal denial; and determine that the target communication satellite is unable to receive the navigation signal in case that the target communication satellite is in an ungrounded state.

[0010] Optionally, the apparatus is configured to perform orbit integration on the target communication satellite according to the predetermined conservative force model coefficient, the target satellite position and the target non-conservative force model coefficient to obtain an integrated orbit, by: fitting the target non-conservative force model coefficient using a least square method to determine a target non-conservative force fitting coefficient; determining a dynamic model of the target communication satellite using the conservative force model coefficient and the target non-conservative force fitting coefficient; and calculating state information of the target communication satellite at any time based on the dynamic model and the target satellite position to obtain the integrated orbit, wherein the state information includes at least one of position data, clock data and velocity data of the target communication satellite.

[0011] Optionally, the apparatus is 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 a target non-conservative force model coefficient, by: determining the initial non-conservative force model parameters according to a type of the target communication satellite; and fitting the initial satellite position, the reference orbit and the initial non-conservative force model parameters using a least square method to obtain the target satellite position and the target non-conservative force model coefficient.

[0012] Optionally, the apparatus is configured to determine the initial non-conservative force model parameters according to the type of the target communication satellite by: in a case that the type of the target communication satellite is a low earth orbit type, the initial non-conservative force model parameters comprise at least one of an atmospheric drag model parameter, a solar radiation pressure model parameter, a magnetic force influence parameter; in a case that the type of the target communication satellite is a medium earth orbit type, the initial non-conservative force model parameters comprise at least one of the solar radiation pressure model parameter, a lunar gravity perturbation parameter, a third body gravity model parameter; in a case that the type of the target communication satellite is a high earth orbit type, the initial non-conservative force model parameters comprise at least one of an earth gravity field model parameter, the lunar gravity perturbation parameter, a solar gravity perturbation parameter.

[0013] Optionally, the apparatus is further configured to: in a case that the inter-satellite link data comprises at least one of a first geometric distance between the target communication satellite and any one of the other communication satellites, a first distance correction term, a first error correction term, and a first observation residual term 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 between the any one of the other communication satellites and the target communication satellite, a second distance correction term, a second error correction term, and a second observation residual term 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 a product of a light speed, a clock bias between a time of transmission of the signal by the target communication satellite and a time of reception of the signal by the any one of the other communication satellites; the second distance correction term is determined by a product of a light speed, a clock bias between a time of transmission of the signal by the any one of the other communication satellites and a time of reception of the signal by the target communication satellite; the first error correction term and the second error correction term are both determined by a satellite antenna phase center offset, a relativistic effect, an atmospheric delay; the first observation residual term and the second observation residual term are both determined by an observation noise generated in a process of obtaining the inter-satellite link data.

[0014] Optionally, the apparatus is further configured to: determine the first geometric distance according to a position of a transmission time of the signal transmitted by the target communication satellite and a position of a reception time of the signal by the any communication satellite; determine the second geometric distance according to a position of a transmission time of the signal transmitted by the any communication satellite and a position of a reception time of the signal by the target communication satellite; adjust 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 normalize the first geometric distance using a clock error of a predicted orbit generated in the orbit filtering; adjust 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 normalize the second geometric distance using the clock error of the predicted orbit; perform first-order Taylor expansion on the first geometric distance and the second geometric distance respectively, and retain 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; and determine the target operational orbit based on the first target geometric distance, the second target geometric distance, and the position error sequence.

[0015] Optionally, the apparatus is 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, by: determining the initial non-conservative force model parameters according to a type of the target communication satellite, wherein the type of the target communication satellite is any one of a low-orbit type, a medium-orbit type, and a high-orbit type.

[0016] Optionally, the apparatus is configured to perform orbit filtering on the target operational orbit using inter-satellite link data in a case where the target communication satellite is unable to receive the navigation signal, by: obtaining relative distances and relative clock errors between the target communication satellite and other communication satellites indicated by the inter-satellite link data, wherein the target communication satellite and the other communication satellites are in a same satellite constellation, in the case where the target communication satellite is unable to receive the navigation signal; and performing the orbit filtering based on the relative distances and the relative clock errors to determine the target operational orbit, wherein, in the process of the orbit filtering, an initial orbit state quantity is determined by the integral orbit, and an initial process noise covariance is determined by the process noise covariance information.

[0017] According to still another aspect of the embodiments of the present application, a computer readable storage medium is provided, which stores a computer program. The computer program is configured to execute the above method for determining a satellite operational orbit when running.

[0018] According to a further aspect of the embodiments of the present application, a computer program product or computer program is provided, which comprises computer instructions stored in a computer readable storage medium. A processor of a computer device reads the computer instructions from the computer readable storage medium, and the processor executes the computer instructions to enable the computer device to perform the method for determining a satellite orbit as described above.

[0019] According to a further aspect of the embodiments of the present application, an electronic device is also provided, which comprises a memory and a processor, wherein the memory stores a computer program, and the processor is configured to execute the computer program to perform the method for determining a satellite orbit as described above.

[0020] In the embodiments of the present application, in the case that the target communication satellite can receive navigation signals, an initial satellite position of the target communication satellite is obtained, and a reference orbit of the target communication satellite is determined based on space-time information of navigation satellites; orbit fitting is performed 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; orbit integration is performed 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; the integrated orbit and the reference orbit are compared to determine a position error sequence of the target communication satellite, and process noise covariance information under different integration lengths in the integrated orbit is determined based on the position error sequence; in the case that the target communication satellite cannot receive navigation signals, orbit filtering is performed using inter-satellite link data to determine a target orbit, which effectively suppresses the accumulation of overall rotation errors of a low-orbit constellation by comprehensively using non-conservative force model parameters, a position error sequence and inter-satellite link data, significantly improves the autonomous operation capability and orbit determination accuracy of a satellite network, meets the needs of real-time autonomous determination and robust service of an overall orbit of a satellite constellation in a confrontation scenario, and further solves the technical problem that satellites cannot operate according to correct orbits due to overall rotation of a satellite constellation. BRIEF DESCRIPTION OF DRAWINGS

[0021] The accompanying drawings, which are included to provide a further understanding of the present application and constitute a part of this application, illustrate certain illustrative embodiments of the present application and together with the description, serve to explain 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 flowchart of an optional method for determining a satellite orbit according to an embodiment of the present application;

[0024] Figure 3 is a flow diagram of a method for determining a satellite orbit according to an embodiment of the application;

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

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

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

[0028] Figure 7 is a structural diagram of an electronic device according to an embodiment of the application. DETAILED DESCRIPTION

[0029] In order to make the personnel in the technical field better understand the scheme of the application, the technical scheme in the embodiments of the application will be clearly and completely described below in conjunction with the drawings in the embodiments of the application. Obviously, the described embodiments are only a part of the embodiments of the application, rather than all the embodiments. Based on the embodiments in the application, all other embodiments obtained by a person of ordinary skill in the art without creative labor should fall within the scope of protection of the application.

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

[0031] The application will be described below in conjunction with the embodiments:

[0032] According to an aspect of the embodiments of the application, a method for determining a satellite orbit is provided. Optionally, in the present embodiment, the above-mentioned method for determining a satellite orbit can be applied in a hardware environment composed of a server 101 and a terminal device 103 as shown in the following figure: Figure 1 as shown in the following figure:Figure 1 As shown, the server 101 is connected with the terminal device 103 through a network, which can be used to provide services for the terminal device or the application program installed on the terminal device, and the application program 107 can be a video application program, an instant messaging application program, a browser application program, an education application program, a game application program, etc. The database 105 can be set on the server or independently of the server, which 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 and a wireless network, wherein the wired network includes a local area network, a metropolitan area network and a wide area network, and the wireless network includes Bluetooth, WIFI and other wireless communication networks. The terminal device 103 can be a terminal configured with an application program, 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 notebook computer, a tablet computer, a palm computer, a MID (Mobile Internet Device, mobile internet device), a PAD, a desktop computer, a smart television, a smart voice interaction device, a smart home appliance, a vehicle-mounted terminal, an aircraft, a virtual reality (Virtual Reality, VR for short) terminal, an augmented reality (Augmented Reality, AR for short) terminal, a mixed reality (Mixed Reality, MR for short) terminal and other computer devices. The above server can be a single server, a server cluster composed of multiple servers, or a cloud server.

[0033] In combination Figure 1 As shown, the above satellite orbit determination method can be executed by an electronic device, which can be a terminal device or a server. The above satellite orbit determination method can be respectively implemented 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 the present embodiment is not specifically limited. Alternatively, as an optional implementation, as shown in Figure 2 As shown, the above satellite orbit determination method includes:

[0035] S202, in the case that the target communication satellite can receive the navigation signal, an initial satellite position of the target communication satellite is acquired, and a reference orbit of the target communication satellite is determined based on space-time information of the navigation satellite, wherein the space-time information includes a position and a clock error of the navigation satellite, and the reference orbit is an orbit determined by the target communication satellite in 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 the three-dimensional spatial coordinates and other information of the target communication satellite at a specific time, and the target communication satellite can be a low-orbit satellite, a medium-orbit satellite, a high-orbit satellite, etc., and the present application does not make any limitation in this regard; the above space-time information refers to the accurate position and clock bias information of the navigation satellite (such as GPS or Beidou satellite) at a specific time; and the above reference orbit is an expected running orbit calculated and determined by the target communication satellite according to the space-time information provided by the navigation satellite, combined with its own dynamic model and observation data, during the autonomous orbit determination, including but not limited to the expected position and velocity vector of the target satellite during the autonomous orbit determination.

[0037] S204, orbit fitting is performed on the initial satellite position, the reference orbit and the initial non-conservative force model parameter to obtain a target satellite position and a target non-conservative force model coefficient.

[0038] Optionally, in the embodiments of the present application, the orbit fitting on the initial satellite position, the reference orbit and the initial non-conservative force model parameter refers to adjusting the satellite position and the dynamic model parameter using an optimization algorithm to minimize the difference between the observation value and the model prediction value; the target satellite position represents the currently optimal estimated position information of the target communication satellite; and the target non-conservative force model coefficient represents the non-conservative force parameter affecting the running orbit of the target communication satellite, such as solar radiation pressure, atmospheric drag, etc.

[0039] S206, orbit integration is performed on the target communication satellite according to the pre-determined conservative force model coefficient, the target satellite position and the target non-conservative force model coefficient to determine an integrated orbit of the target communication satellite.

[0040] Optionally, in the embodiments of the present application, the above orbit integration refers to a prediction process of the running orbit of the target communication satellite based on the dynamic model, using the conservative force model coefficient (such as gravity parameter), the target satellite position and the target non-conservative force model coefficient, and the integrated orbit is the result of the prediction, which reflects the predicted position and velocity of the satellite with respect to time considering the influence of all forces.

[0041] S208, the position error sequence of the target communication satellite is determined by comparing the integrated orbit and the reference orbit, and the process noise covariance information under different integration lengths in the integrated orbit is determined based on the position error sequence, wherein the position error sequence is used to indicate the change of the position deviation of the target communication satellite with respect to the target running orbit over time during the autonomous orbit determination.

[0042] Optionally, in the embodiments of the present application, the comparison refers to comparing the integral orbit with the reference orbit to determine the difference between the two orbits, i.e., the position error sequence; and the position error sequence contains the position deviation information of the target communication satellite relative to the target orbit at different time points, which is used to identify and correct the overall constellation rotation and drift error; and the process noise covariance information refers to a matrix describing the statistical characteristics of the prediction error of the satellite state (such as position, velocity, clock bias, etc.) in the Kalman filter or similar state estimation orbit filtering algorithm, which quantifies the state estimation uncertainty caused by incomplete dynamic model, external random disturbance (such as solar pressure, atmospheric drag, etc. non-conservative force) and measurement noise, etc. In the autonomous orbit determination process, it is used to adjust the weight of the prediction model and the observation data of the filter to achieve optimal state estimation. The process noise covariance information under different integration lengths reflects the uncertainty level of satellite state prediction over time, which is a key parameter for optimizing orbit calculation accuracy and robustness.

[0043] S210, in the case that the target communication satellite cannot receive navigation signals, using inter-satellite link data to perform orbit filtering to determine the target orbit.

[0044] Optionally, in the embodiments of the present application, the orbit filtering refers to the process of real-time updating and correcting the satellite orbit based on inter-satellite link data and position error sequence through a filtering algorithm when the target communication satellite loses external GNSS signals. The 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 target orbit refers to the real-time running track of the target communication satellite determined after the orbit filtering algorithm is corrected, which integrates the position error sequence and the 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 correct orbit even in complex environments.

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

[0046] Illustratively, the initial position of the target communication satellite is first accurately obtained, and then the high-precision space-time information (including position and clock bias) of the navigation satellite is used to determine the reference orbit of the target communication satellite; Next, the initial satellite position, the preset reference orbit, and the initial non-conservative force model parameters based on the satellite type are used for orbit fitting to obtain the optimized target satellite position and model coefficients reflecting the actual influence of non-conservative forces. These non-conservative forces can include but are not limited to solar pressure, atmospheric drag, etc.

[0047] Then, the target satellite position obtained by fitting the determined conservative force model coefficient (such as the earth gravity parameter) and the target non-conservative force model coefficient is used for orbit integration of the target communication satellite to predict the future running track of the target communication satellite, and an integrated orbit is obtained. By comparing the integrated orbit with the reference orbit, the position deviation of the target communication satellite in the autonomous orbit determination process, that is, the position error sequence, can be determined. Based on the position error sequence, the uncertainty level of the satellite state prediction is further analyzed, and the most suitable process noise covariance information under different integration lengths is determined through statistical analysis and an adaptive algorithm.

[0048] When the target communication satellite cannot receive the navigation signal, the inter-satellite link data is used for orbit filtering to determine the target running track. In the orbit filtering process, the initial orbit state quantity is 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 in the constellation. By combining the inter-satellite link data and the prior constraint (that is, the initial orbit state quantity and the process noise covariance information), the filtering algorithm can effectively suppress the overall rotation error of the constellation, maintain the stability and operation accuracy of the satellite network, and ensure that reliable services can still be provided in the confrontation or GNSS denial environment.

[0049] In an exemplary embodiment, Figure 3 is a flowchart of an optional satellite running track determination method 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 space-time information (including but not limited to the position, clock difference, etc. of the GNSS satellite (the above navigation satellite)) of the GNSS satellite in the current autonomous processing period, calculating the initial satellite position of the communication satellite at the current initial time and the reference orbit.

[0051] S2, performing orbit fitting (including but not limited to least square method algorithm, etc.) on the initial satellite position, the reference orbit, and the related non-conservative force (atmospheric resistance, earth albedo radiation pressure, solar pressure, etc. related to the type of satellite, for example, low orbit, high orbit, etc.), to obtain a more accurate satellite position, denoted as target satellite position and non-conservative force model coefficient.

[0052] S3, combining the conservative force model coefficient and the target non-conservative force fitting coefficient obtained by fitting the non-conservative force model coefficient, and using the digital integration method to perform orbit integration of the target communication satellite, which can be understood as using the current target satellite position to predict the satellite position at the next time, recursively obtaining the satellite position at multiple times, and obtaining the integrated orbit.

[0053] S4, compare the integral orbit with the reference orbit to obtain the position error of the target communication satellite at different times in the future, i.e. the position error sequence, and further perform detailed statistical analysis on the position error sequence, including calculating the mean, variance and autocorrelation function of the time sequence of the error, to identify the statistical characteristics of the error, then use an adaptive algorithm to adjust the process noise covariance information to realize dynamic updating of the process noise covariance information and better match the uncertainty of the satellite state prediction.

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

[0055] Among them, the above-mentioned inter-satellite link data is obtained by mutual observation between each communication satellite in the satellite constellation. Taking communication satellite A and communication satellite B as an example, it can be expressed as The formula represents that the inter-satellite link data includes various information (such as the geometric distance between the satellites, the satellite clock error of the communication satellite B at t1, the satellite clock error of the communication satellite A at t2, the error correction of the satellite antenna phase center, the relativistic effect and the atmospheric delay of the communication satellite A and the communication satellite B, the random noise or unmodeled residual in the observation, etc.) of the communication satellite A and the communication satellite B. The formula represents the geometric distance between the satellites. The formula represents the satellite clock error of the communication satellite B at t1. The formula represents the satellite clock error of the communication satellite A at t2. The formula represents the satellite clock error of the communication satellite A at t2. The formula represents the error correction of the satellite antenna phase center, the relativistic effect and the atmospheric delay of the communication satellite A and the communication satellite B. The formula represents the random noise or unmodeled residual in the observation.

[0056] By the embodiments of the present application, in the case that 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, wherein the space-time information includes the position and clock error of the navigation satellite, and the reference orbit is the running orbit determined by the target communication satellite itself in the autonomous orbit determination process; the initial satellite position, the reference orbit and the initial non-conservative force model parameters are subjected to orbit fitting to obtain the target satellite position and the target non-conservative force model coefficient, wherein the initial non-conservative force model parameters are determined by the type of the target communication satellite; the target communication satellite is subjected to orbit integration according to the pre-determined conservative force model coefficient, the target satellite position and the target non-conservative force model coefficient to determine the integral orbit of the target communication satellite; the integral orbit and the reference orbit are compared to determine the position error sequence of the target communication satellite, and the process noise covariance information under different integration time lengths in the integral orbit is determined based on the position error sequence, wherein the position error sequence is used to indicate the change of the position deviation of the target communication satellite relative to the target running orbit with time in the autonomous orbit determination process; in the case that the target communication satellite cannot receive navigation signals, the inter-satellite link data is used for orbit filtering to determine the target running orbit, wherein the initial orbit state quantity in the process of orbit filtering is determined by the integral orbit, 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 error 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-orbit constellation is effectively suppressed, the autonomous operation capability and the orbit determination precision of the satellite network are significantly improved, the demand for real-time autonomous determination of the overall orbit of the satellite constellation and stable service in the confrontation scene is met, and the technical problem that the satellite cannot run according to the correct orbit due to the overall offset rotation of the satellite constellation is solved.

[0057] As an optional solution, the above comparing the above integral orbit and the above reference orbit, determining the position error sequence of the above target communication satellite, and determining the process noise covariance information under different integration time lengths in the above integral orbit based on the above position error sequence, comprises: obtaining an integral orbit point corresponding to any time of the above integral orbit, and obtaining a reference orbit point corresponding to the above any time of the above reference orbit; determining error data corresponding to the above any time of the above target communication satellite according to the above integral orbit point and the above reference orbit point, wherein the error data comprises at least one of position error data and clock error data; generating the position error sequence according to the error data.

[0058] Optionally, in the embodiments of the present application, the integral orbit point position refers to the predicted position and velocity information of the target communication satellite calculated according to a preset time interval in the integral process; the reference orbit point position refers to the expected position and velocity of the satellite predicted based on initial conditions and a dynamic model in the same time interval; by calculating the difference between the integral orbit point position and the reference orbit point position, the error data of the 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, the target communication satellite obtains an integral orbit point position through orbit integration every 300 seconds in the autonomous orbit determination process, and a reference orbit determined according to GNSS ephemeris data also provides a set of reference orbit point positions at the same time point. By comparing the coordinates of the two orbit point positions at the 300-second time point, the position error data is calculated, the clock error data is determined by comparing the deviation of the clock difference, and the velocity error data is obtained by comparing the velocity vectors. Arranging these error data in chronological order constitutes the above-mentioned position error sequence describing the change of the satellite orbit deviation with time.

[0060] Through the embodiments of the present application, the precise monitoring and evaluation of the orbit error of the target communication satellite are realized by comparing the integral orbit and the reference orbit, achieving the purpose of accurately identifying the orbit deviation of the satellite when the satellite loses external navigation signals, and providing data support for subsequent orbit filtering and correction.

[0061] As an optional solution, the orbit filtering using inter-satellite link data to determine the target running orbit in the case that the target communication satellite cannot receive the navigation signal includes: determining an initial running state of a predicted orbit according to the initial orbit state quantity and the initial process noise covariance information in the case that the target communication satellite cannot receive the navigation signal; correcting the predicted orbit according to the inter-satellite link data according to the Kalman filtering algorithm to obtain the target running 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.

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

[0063] It should be noted that the Kalman filtering algorithm adjusts the orbit parameters of the satellite through two stages of state prediction and observation update, and realizes the minimization of errors, and the position error sequence is used to dynamically adjust the weight of the predicted orbit in the orbit filtering process, so as to ensure that the orbit filtering pays more attention to those data points with smaller errors and more accurate positions, and 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 predicted orbit.

[0064] In an exemplary embodiment, when the target communication satellite encounters a GNSS denial environment during autonomous orbit determination, resulting in the inability to receive external navigation signals, first, based on the obtained position error sequence, the weight of the predicted orbit in the orbit filtering process is adjusted, so that the prediction is closer to the true state of the satellite, then the observation residual of the inter-satellite ranging, that is, the difference between the actual ranging and the predicted value, is calculated using the inter-satellite link data, and the Kalman filtering is performed by comprehensively combining the two, to update the state vector of the satellite, including position, velocity and clock bias, etc. Through continuous prediction and update, the high-precision determination of the satellite orbit is maintained, and the accumulation of the overall rotation error of the constellation is effectively suppressed.

[0065] Through the embodiments of the present application, the strategy of fusing the position error sequence and the inter-satellite link data by Kalman filtering is adopted, the technical effect of accurately determining the running orbit of the target communication satellite in a complex environment where the target communication satellite cannot receive navigation signals is realized, and the purpose of effectively controlling the overall rotation error of the constellation and maintaining the autonomous operation capability of the satellite network is achieved.

[0066] As an optional solution, the above method further comprises: determining that the target communication satellite cannot receive the navigation signal when the navigation signal is denied; and determining that the target communication satellite cannot receive the navigation signal when the target communication satellite is in an ungrounded state.

[0067] Optionally, in the embodiments of the present application, the navigation signal rejection means that the target communication satellite is in a signal shielding environment caused by human or natural factors on the earth's surface or in the atmosphere, which causes it to be unable to receive signals of a global navigation satellite system (such as GPS, Beidou, Galileo, etc.); and the no ground support state means that the communication link between the satellite and the ground station is interrupted, and the satellite cannot obtain data or instructions transmitted from the ground.

[0068] For example, when the navigation signal receives signals of a global navigation satellite system (such as GPS, Beidou, Galileo, etc.) and the target communication satellite is in a no ground support state, the target communication satellite cannot receive the navigation signal;

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

[0070] In still another example, when the navigation signal is in a non-rejection state and the target communication satellite is in a no ground support state, the target communication satellite still cannot receive the navigation signal, that is, when the target communication satellite has a receiving blind area or signal acquisition obstacle, the target communication satellite is still determined to be unable to receive the navigation signal.

[0071] In general, when at least one of the navigation signal rejection and the target communication satellite in the no ground support state occurs, the target communication satellite is determined to be unable to receive the navigation signal.

[0072] It should be noted that the navigation signal rejection and the no ground support state can be caused by signal interference, natural phenomena (such as solar storm), equipment failure, satellite deployment strategy (such as optimization of relative positions between satellites), or specific operation requirements, and the scheme proposed in the present application is applicable to all these cases, and is not limited thereto.

[0073] In an exemplary embodiment, when the target communication satellite is performing a task, it encounters strong signal interference, which causes it to be unable to receive the navigation signal, and the autonomous running system of the target communication satellite immediately starts the navigation signal rejection detection mechanism, confirms the rejection state, and then, in the no ground support state, implements the satellite running orbit determination method proposed in the present application.

[0074] Through the embodiments of the present application, the signal rejection detection and the orbit filtering strategy in the no ground support state are adopted, the technical effect of autonomously determining the satellite running orbit under the extreme condition that the satellite loses external navigation signals and ground support is achieved, and the purpose of ensuring stable operation of the satellite constellation and providing high-precision navigation and positioning services in a complex environment is achieved.

[0075] As an optional solution, the orbit integration of the target communication satellite according to the predetermined conservative force model coefficient, the target satellite position and the target non-conservative force model coefficient comprises: fitting the target non-conservative force model coefficient by using a least square method to determine a target non-conservative force fitting coefficient; determining a dynamic model of the target communication satellite by using the conservative force model coefficient and the target non-conservative force fitting coefficient; and calculating state information of the target communication satellite at any time based on the dynamic model and the target satellite position to obtain the integral orbit, wherein the state information comprises at least one of position data, clock data and velocity data of the target communication satellite.

[0076] Optionally, in the embodiment of the present application, the conservative force model coefficient is used to describe the basic dynamic characteristics of satellite motion in the process of orbit integration, and the target non-conservative force model coefficient covers parameters of the influence of solar radiation pressure, atmospheric drag, non-spherical earth gravity field and other non-conservative forces on the satellite orbit, which can be determined by least square fitting; and the dynamic model can be understood as a mathematical model for describing the state of satellite motion based on the conservative force and non-conservative force model coefficients.

[0077] It should be noted that the coefficients of the conservative force and non-conservative force model can be different according to the specific type, operating environment and task requirements of the target communication satellite, the fitting algorithm can be other algorithms except least square 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 be different according to the specific implementation, and the present application does not make specific limitation.

[0078] By the embodiment of the present application, the orbit integration strategy combining the dynamic model based on the conservative force and non-conservative force model coefficients with the least square fitting is adopted, the technical effect of accurately determining the orbit of the target communication satellite in a complex environment is achieved, and the purpose of effectively controlling the overall rotation error and maintaining the high-precision orbit determination capability in the process of autonomous orbit determination of the satellite constellation is achieved.

[0079] As an optional solution, the orbit fitting of the initial satellite position, the reference orbit and the initial non-conservative force model parameter to obtain the target satellite position and the target non-conservative force model coefficient comprises: determining the initial non-conservative force model parameter according to the type of the target communication satellite; and fitting the initial satellite position, the reference orbit and the initial non-conservative force model parameter by using a least square method to obtain the target satellite position and the target non-conservative force model coefficient.

[0080] Optionally, in the embodiments of the present application, the 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 measurement results of the GNSS receiver on the satellite, and the reference orbit is a set of stable orbit paths calculated in advance according to factors such as satellite type, task requirements, and geodetic environment.

[0081] It should be noted that the type of the target communication satellite can include low-orbit satellites, medium-orbit satellites, high-orbit satellites, etc., and the selection and weight of the non-conservative force model parameters of each type of satellite will be different, so the determination method of the initial satellite position and the reference orbit can also vary according to the specific application scenario, such as using GNSS precise ephemeris data, ground radar measurement, or data provided by other satellite constellations. In addition, the use of the least squares method, such as the number of iterations, the convergence condition, the data window size, etc., can also be adjusted according to the implementation details, which is not limited in the present application.

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

[0083] Through the embodiments of the present application, the orbit fitting method combining the non-conservative force model parameters customized for the satellite type with the least squares fitting is used to achieve the technical effect of accurate estimation of the satellite orbit and the non-conservative force model parameters in the process of autonomous orbit determination of the satellite constellation, which eliminates the rank deficiency, ensures the stability and accuracy of the autonomous orbit determination algorithm, and effectively suppresses the overall rotation error of the constellation.

[0084] As an optional solution, the above determining the initial non-conservative force model parameters according to the type of the target communication satellite comprises: in the case that the type of the target communication satellite is a low-orbit type, the initial non-conservative force model parameters comprise at least one of an atmospheric drag model parameter, a solar radiation pressure model parameter, and a magnetic force influence parameter; in the case that the type of the target communication satellite is a medium-orbit type, the initial non-conservative force model parameters comprise at least one of the solar radiation pressure model parameter, a lunar gravitational perturbation parameter, and a third-body gravitational model parameter; and in the case that the type of the target communication satellite is a high-orbit type, the initial non-conservative force model parameters comprise at least one of an earth gravitational field model parameter, the 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 can be classified according to the height of the orbit on which the satellite runs, and specifically can be classified into a low-orbit type (such as LEO, Low Earth Orbit), a medium-orbit type (such as MEO, Medium Earth Orbit), and a high-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 force which are preset according to the type of the satellite and used to describe the influence of external force on the orbit of the satellite.

[0086] For example, in the case that the type of the target communication satellite is a low-orbit type, but not a medium-orbit type or a high-orbit type, the initial non-conservative force model parameters can comprise 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, in the case that the type of the target communication satellite is a medium-orbit type, but not a low-orbit type or a high-orbit type, the initial non-conservative force model parameters can comprise one or more of a solar radiation pressure model parameter, a lunar gravitational perturbation parameter, and a third-body gravitational model parameter.

[0088] In still another example, in the case that the type of the target communication satellite is a low-orbit type, but not a medium-orbit type or a high-orbit type, the initial non-conservative force model parameters can comprise one or more of an earth gravitational field model parameter, a lunar gravitational perturbation parameter, and a solar gravitational perturbation parameter.

[0089] In general, 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 to be obtained are not completely the same for different types of target communication satellites.

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

[0091] As an optional solution, the method further comprises: determining the first geometric distance according to a position of the transmission time of the signals transmitted by the target communication satellite and a position of the reception time of the signals received by the any one of the other communication satellites; determining the second geometric distance according to a position of the transmission time of the signals transmitted by the any one of the other communication satellites and a position of the reception time of the signals received by the target communication satellite; adjusting 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 normalizing the first geometric distance by using a clock bias of a predicted orbit generated in the orbit filtering; adjusting 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 normalizing the second geometric distance by using the clock bias of the predicted orbit; performing first-order Taylor expansion on the first geometric distance and the second geometric distance respectively, and retaining 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; and determining the target operational orbit based on the first target geometric distance, the second target geometric distance, and the position error sequence.

[0092] Specifically, Figure 4 is an inter-satellite link bidirectional observation schematic diagram of an optional satellite operational orbit determination method according to an embodiment of the present application, as Figure 4 shown, an inter-satellite link original observation equation can be expressed as:

[0093] .

[0094] .

[0095] where, and denote the actual reception time of the signal, and denote the actual transmission time of the signal, and denote the inter-satellite link observations of a communication satellite A transmitting and a communication satellite B receiving, and a communication satellite B transmitting and a communication satellite A receiving, respectively, in the same satellite constellation, assuming that the communication satellite A is the target communication satellite, corresponding to the first geometric distance, corresponding to the second geometric distance, corresponding to the first distance correction term, corresponding to the second distance correction term, denote the satellite clock bias of the communication satellite B at t1, denote the satellite clock bias of the communication satellite A at t1, denote the satellite clock bias of the communication satellite A at t2, denote the satellite clock bias of the communication satellite A at t2, denote the satellite clock bias of the communication satellite B at t2, denote the satellite clock bias of the communication satellite A at t2, denote the error corrections for the satellite antenna phase center, relativistic effects, atmospheric delays, etc. when the communication satellite A transmits and the communication satellite B receives, corresponding to the first error correction term, BA denote the error corrections for the satellite antenna phase center, relativistic effects, atmospheric delays, etc. when the communication satellite A transmits and the communication satellite B receives, corresponding to the second error correction term, denote the random noise or un-modeled residuals in the observations, and c is the speed of light, corresponding to the first observation residual term, corresponding to the second observation residual term.

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

[0097] .

[0098] .

[0099] After correcting various errors, the observed values are normalized using the predicted orbit clock difference, the inter-satellite link sampling interval is 300 seconds, the round-trip link interval is 1.5 seconds, the data processing interval is set to 300 seconds, and the bi-directional 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 time is not greater than 2 seconds.

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

[0101] .

[0102] .

[0103] wherein, 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 error of the actual signal transmission and reception time, and may be a communication satellite A or a communication satellite B.

[0104] wherein, the partial derivative of the geometric distance with respect to the satellite orbit is specifically as follows:

[0105] .

[0106] .

[0107] .

[0108] .

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

[0110] .

[0111] .

[0112] .

[0113] .

[0114] wherein, , T indicates time, x is a position vector of the satellite in an inertial system, v is a velocity vector of the satellite in the inertial system, p is a dynamic parameter, which is generally a solar radiation pressure model parameter, an atmospheric drag parameter, an empirical force model parameter, or the like, and denotes a state transition matrix, which can be obtained by integrating a variational equation, and denotes any time.

[0115] As an optional solution, the above initial satellite position, the above reference orbit, and the initial non-conservative force model parameter are orbit-fitted to obtain a target satellite position and a target non-conservative force model coefficient, including that the initial non-conservative force model parameter is determined according to a type of the target communication satellite, and the type of the target communication satellite is any one of a low-orbit type, a medium-orbit type, and a high-orbit type.

[0116] Optionally, in the embodiment of the application, the above orbit fitting can be performed by a numerical analysis method and a dynamic model, and the non-conservative force model parameter can include, but is not limited to, a solar radiation pressure parameter and an atmospheric drag parameter.

[0117] It should be noted that the above technical solutions of the application are applicable to communication satellites of different types of orbits, whether low-orbit, medium-orbit, or high-orbit, and the target non-conservative force model coefficient can be set according to the orbit environment of the satellite and the influence degree of the non-conservative force, so as to improve the orbit fitting accuracy, which is not limited in the application.

[0118] In an exemplary embodiment, taking the application scenario of a low-orbit satellite as an example, when the satellite constellation performs orbit solving by using inter-satellite link data in autonomous operation, the initial non-conservative force model parameter, including a solar radiation pressure model and an atmospheric drag model, can be set according to the characteristics of the low-orbit satellite in advance, and then the parameter is continuously optimized by an orbit fitting algorithm, so as to improve the accuracy and stability of the orbit solving.

[0119] According to the embodiment of the application, the initial non-conservative force model parameter set according to the type of the satellite is used for orbit fitting, high-precision orbit solving of the communication satellite in the case of navigation signal denial or without ground support is realized, and the purpose of autonomous orbit determination and effective control of the overall rotation error of the constellation is achieved.

[0120] As an optional solution, the orbit filtering using the inter-satellite link data to determine the target operational orbit in the case that the target communication satellite cannot receive the navigation signal comprises: obtaining the relative distance and the relative clock difference between the target communication satellite and other communication satellites in the same satellite constellation indicated by the inter-satellite link data in the case that the target communication satellite cannot receive the navigation signal; and performing the orbit filtering based on the relative distance and the relative clock difference to determine the target operational orbit, wherein, 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.

[0121] Optionally, in the embodiments of the present application, the inter-satellite link data refers to the relative distance and the relative clock difference data between the target communication satellite and other communication satellites in the same satellite constellation obtained through bidirectional link communication, including but not limited to the mutual distance measurement value and the time synchronization difference value exchanged in real time by the on-board equipment, which can be used for autonomous orbit filtering in the case that the target satellite loses the navigation signal, wherein the on-board equipment refers to the communication antenna, the distance measuring device, the time synchronization module, 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, which can be used to process the inter-satellite link data to determine the target operational orbit, and the present application does not limit this.

[0123] In an exemplary embodiment, taking the application scenario of medium orbit communication satellites as an example, when the target communication satellite temporarily cannot receive the navigation signal due to entering a certain area, the relative distance and the 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 the relative clock difference, combined with the initial orbit state quantity and the process noise covariance obtained by the previous orbit integration, real-time orbit updating is performed through the extended Kalman filtering algorithm, ensuring the accuracy and stability of the target satellite operational orbit during the navigation signal interruption period.

[0124] Through the embodiments of the present application, the orbit filtering using the inter-satellite link data realizes the autonomous determination of the target communication satellite operational orbit in the case that the navigation signal cannot be received, and achieves the purpose of maintaining the overall operational performance of the satellite constellation and providing continuous navigation service.

[0125] In an exemplary embodiment, the embodiments of the present application can be applied to the field of satellite navigation technology, and realize a satellite constellation autonomous orbit determination overall rotation error control method, wherein the satellite orbit error is a key element of the satellite space reference, and its accuracy directly determines the service performance of the satellite, and the high-precision navigation enhancement service puts forward a very high centimeter-level accuracy requirement on the orbit information of the low earth orbit satellite.

[0126] In the prior art, the low earth orbit constellation orbit autonomous determination based on inter-satellite link can be realized through inter-satellite cooperative control and mutual acquisition of state information, and the autonomous orbit determination of the constellation is generally carried out by using the inter-satellite distance observation value. However, the 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 directly affect the orbit determination accuracy of the constellation, and the longer the time, the greater the orbit determination error, which makes the high-precision navigation enhancement service face serious degradation risk.

[0127] Therefore, the embodiments of the present application can realize a satellite constellation autonomous orbit determination overall rotation error control method suitable for on-board processing, which can effectively suppress the overall rotation error of the constellation in the case of GNSS denial or lack of ground support for the satellite, significantly improve the autonomous solution accuracy of the low earth orbit constellation orbit, and meet the real-time autonomous determination and stable service requirements of the low earth orbit constellation overall orbit in the future confrontation scene, so as to enhance the autonomous operation capability and improve the comprehensive operation and maintenance level of the system.

[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 coefficient is constrained, and after the constellation rotation error of the filter solution orbit relative to the reference orbit is estimated, the orbit parameters obtained by the autonomous orbit determination are directly corrected, so as to realize the suppression of the overall rotation error of the constellation, and improve the autonomous operation capability of the satellite network, including but not limited to:

[0129] S1, acquiring the space-time information (including but not limited to the position of the GNSS satellite, the clock error, etc.) of the GNSS satellite (the above navigation satellite) in the current autonomous processing period by using the target communication satellite, and calculating the initial satellite position and reference orbit of the communication satellite at the current initial time, for example, acquiring the GNSS precise ephemeris data file in the current autonomous processing period by using the target communication satellite, and calculating the initial satellite position and reference orbit.

[0130] S2, performing orbit fitting (including but not limited to least square method algorithm, etc.) on the initial satellite position, reference orbit and related non-conservative force (atmospheric drag, earth albedo radiation pressure, solar radiation pressure, etc., related to the type of satellite, for example, low earth orbit, high earth orbit, etc.), to obtain more accurate satellite position, denoted as target satellite position and non-conservative force model coefficient.

[0131] S3, in combination with the conservative force model coefficient and the target non-conservative force fitting coefficient obtained by fitting the non-conservative force model coefficient, the orbit of the target communication satellite is integrated by using a digital integration method, which can be understood as predicting the satellite position at the next time using the current target satellite position, recursively obtaining the satellite position at multiple times, and obtaining the integral orbit.

[0132] S4, comparing the integral orbit with the reference orbit, the position error of the target communication satellite at different times in the future is obtained, that is, the position error sequence.

[0133] S5, in the GNSS denial or no ground support state, the position error sequence and the inter-satellite link data are used for orbit filtering (including but not limited to Kalman filtering), to determine the target orbit of the communication satellite, that is, using the dynamic model parameters of the orbit fitting in the previous period as the prior constraint information of the autonomous orbit determination algorithm, adjusting the weight of the predicted orbit in the orbit filtering, and suppressing the accumulation of the overall rotation error of the low-orbit constellation.

[0134] Through the embodiments of the present application, the prior orbit parameter constraint is introduced to suppress the overall rotation error of the low-orbit constellation, the dynamic model parameters of the orbit fitting in the previous period are used as the prior constraint information of the autonomous orbit determination algorithm, the rank deficiency of the autonomous orbit determination is eliminated, and the autonomous orbit determination algorithm can be correctly and stably solved, so that the accumulation of the overall rotation error of the low-orbit constellation is effectively suppressed, and the autonomous solution precision of the low-orbit constellation orbit clock difference is significantly improved.

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

[0136] It should be noted that, for the foregoing method embodiments, in order to simply describe, they are all expressed as a series of action combinations, but those skilled in the art should know that the present application is not limited to the action order described, because according to the present application, certain steps can be performed in other order or simultaneously. Secondly, those skilled in the art should know that the embodiments described in the specification all belong to preferred embodiments, and the actions and modules involved are not necessarily required by the present application.

[0137] According to another aspect of the embodiments of the present application, a satellite orbit determination device for implementing the satellite orbit determination method described above is also provided. As shown in the figure, Figure 5 the device comprises:

[0138] The acquisition module 502 is configured to acquire an initial satellite position of the target communication satellite when the target communication satellite is capable of receiving the navigation signal, and determine a reference orbit of the target communication satellite based on space-time information of the navigation satellite, wherein the space-time information comprises a position and a clock error of the navigation satellite, and the reference orbit is an operation orbit determined by the target communication satellite in the autonomous orbit determination process.

[0139] The fitting module 504 is 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.

[0140] The integration module 506 is configured to perform orbit integration on the target communication satellite according to predetermined 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] The comparison module 508 is 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 under different integration time lengths in the integrated orbit based on the position error sequence, wherein the position error sequence is used to indicate a change of a position deviation of the target communication satellite relative to the target operation orbit over time in the autonomous orbit determination process.

[0142] The determination module 510 is configured to perform orbit filtering using inter-satellite link data to determine the target operation orbit when the target communication satellite is incapable of receiving the navigation signal.

[0143] As an optional solution, the device is configured to compare the integrated orbit and the reference orbit to determine the position error sequence of the target communication satellite, and determine the process noise covariance information under different integration time lengths in the integrated orbit based on the position error sequence in the following manner: acquiring an integrated orbit point corresponding to the integrated orbit at any time, and acquiring a reference orbit point corresponding to the reference orbit at any time; determining error data corresponding to the target communication satellite at any time according to the integrated orbit point and the reference orbit point, wherein the error data comprises at least one of position error data and clock error data; and generating the position error sequence according to the error data.

[0144] As an optional solution, the device is further configured to determine the target operational orbit using the inter-satellite link data for orbit filtering when the target communication satellite is unable to receive the navigation signal by: determining an initial operational state of a predicted orbit according to the initial orbit state quantity and the initial process noise covariance information when the target communication satellite is unable to receive the navigation signal; and correcting the predicted orbit according to the inter-satellite link data according to a Kalman filtering algorithm to obtain the target operational orbit; wherein the inter-satellite link data is used to determine an observation residual in the orbit filtering, and the observation residual represents a difference between an inter-satellite ranging observation value indicated by the inter-satellite link data and an inter-satellite ranging expected value of the predicted orbit.

[0145] As an optional solution, the device is further configured to determine that the target communication satellite is unable to receive the navigation signal when the navigation signal is blocked, and determine that the target communication satellite is unable to receive the navigation signal when the target communication satellite is in a no-ground-support state.

[0146] As an optional solution, the device is further configured to determine the target operational orbit using the inter-satellite link data for orbit filtering when the target communication satellite is unable to receive the navigation signal by: determining an initial operational state of a predicted orbit according to the initial orbit state quantity and the initial process noise covariance information when the target communication satellite is unable to receive the navigation signal; and correcting the predicted orbit according to the inter-satellite link data according to a Kalman filtering algorithm to obtain the target operational orbit; wherein the inter-satellite link data is used to determine an observation residual in the orbit filtering, and the observation residual represents a difference between an inter-satellite ranging observation value indicated by the inter-satellite link data and an inter-satellite ranging expected value of the predicted orbit.

[0147] As an optional solution, the device is further configured to determine the target operational orbit using the inter-satellite link data for orbit filtering when the target communication satellite is unable to receive the navigation signal by: determining an initial operational state of a predicted orbit according to the initial orbit state quantity and the initial process noise covariance information when the target communication satellite is unable to receive the navigation signal; and correcting the predicted orbit according to the inter-satellite link data according to a Kalman filtering algorithm to obtain the target operational orbit; wherein the inter-satellite link data is used to determine an observation residual in the orbit filtering, and the observation residual represents a difference between an inter-satellite ranging observation value indicated by the inter-satellite link data and an inter-satellite ranging expected value of the predicted orbit.

[0148] As an optional solution, the apparatus is further configured to determine the initial non-conservative force model parameters according to the type of the target communication satellite in the following manner: in the case that the type of the target communication satellite is a low-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; in the case that the type of the target communication satellite is a medium-orbit type, the initial non-conservative force model parameters include at least one of a solar radiation pressure model parameter, a lunar gravity perturbation parameter, and a third body gravity model parameter; in the case that the type of the target communication satellite is a high-orbit type, the initial non-conservative force model parameters include at least one of an earth gravity field model parameter, a lunar gravity perturbation parameter, and a solar gravity perturbation parameter.

[0149] As an optional solution, the apparatus is further configured to: in the case that the inter-satellite link data includes a signal transmitted by the target communication satellite to any one of the other communication satellites, at least one of a first geometric distance between the target communication satellite and the any one of the other communication satellites, a first distance correction term, a first error correction term, and a first observation residual term; and in the case that the inter-satellite link data includes a signal transmitted by the any one of the other communication satellites to the target communication satellite, at least one of a second geometric distance between the any one of the other communication satellites and the target communication satellite, a second distance correction term, a second error correction term, and a second observation residual term; wherein the first distance correction term is determined by a product of the speed of light, a clock bias between a time of transmission of the signal by the target communication satellite and a time of reception of the signal by the any one of the other communication satellites; the second distance correction term is determined by a product of the speed of light, a clock bias between a time of transmission of the signal by the any one of the other communication satellites and a time of reception of the signal by the target communication satellite; the first error correction term and the second error correction term are each determined by a satellite antenna phase center offset, a relativistic effect, and an atmospheric delay; and the first observation residual term and the second observation residual term are each determined by an observation noise generated in the process of obtaining the inter-satellite link data.

[0150] As an optional solution, the device is further configured to: determine the first geometric distance according to the position of the transmission time of the signal transmitted by the target communication satellite and the position of the reception time of any communication satellite; determine the second geometric distance according to the position of the transmission time of the signal transmitted by any communication satellite and the position of the reception time of the target communication satellite; 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 normalize 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 normalize 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, and retain the first-order term to 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 operational orbit based on the first target geometric distance, the second target geometric distance and the position error sequence.

[0151] As an optional solution, the device is configured to perform orbit fitting on the initial satellite position, the reference orbit and the initial non-conservative force model parameter to obtain the target satellite position and the target non-conservative force model coefficient by the following manner: the initial non-conservative force model parameter is determined according to the type of the target communication satellite, wherein the type of the target communication satellite is any one of a low-orbit type, a medium-orbit type and a high-orbit type.

[0152] As an optional solution, the device is configured to perform orbit filtering on the target operational orbit by using the inter-satellite link data in the case that the target communication satellite cannot receive the navigation signal by the following manner: in the case that the target communication satellite cannot receive the navigation signal, the relative distance and the relative clock error between the target communication satellite and other communication satellites are obtained according to the inter-satellite link data, wherein the target communication satellite and the other communication satellites are in the same satellite constellation; and the target operational orbit is determined by performing orbit filtering based on the relative distance and the relative clock error, wherein in the process of the orbit filtering, the initial orbit state quantity is determined by the integral 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 or a part of a computer program with a predetermined function, and works with other related parts to achieve a predetermined target, and can be implemented entirely or partially 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 an overall module or unit that includes the functions of the module or unit.

[0154] As to the apparatus in the above-mentioned embodiments, the specific manners in which various modules perform operations have been described in detail in the embodiments related to the method, and thus will not be described in detail here.

[0155] According to an aspect of the present application, there is provided a computer program product comprising a computer program.

[0156] The above-mentioned sequence numbers of the embodiments of the present application are merely for description, and do not represent advantages or disadvantages of the embodiments.

[0157] Figure 6 A computer system structure block diagram 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 bring any limitation to the functions and use range 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 programs stored in a read-only memory 602 (ROM) or loaded from a storage portion 608 into a random access memory 603 (RAM). In the random access memory 603, various programs and data required for system operation are also stored. The central processing unit 601, the read-only memory 602 and the random access memory 603 are connected to each other through a bus 604. An input / output interface 605 (I / O interface) is also connected to the bus 604.

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

[0161] In particular, according to the embodiments of the present application, the processes described in each method flowchart can be implemented as a computer software program. For example, the embodiments of the present application include a computer program product comprising a computer program carried on a computer readable medium, the computer program containing program code for executing the methods shown in the flowcharts. In such embodiments, the computer program can be downloaded and installed from a network through the communication section 609, and / or installed from the detachable 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 performed.

[0162] In such embodiments, the computer program can be downloaded and installed from a network through the communication section 609, and / or installed from the detachable 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 performed.

[0163] According to still another aspect of the embodiments of the present application, an electronic device for implementing the above-mentioned method for determining a satellite orbit is also provided, which can be the terminal device or the server as shown. Figure 1 The present embodiment takes the terminal device as an example to illustrate. As shown in the figure, the electronic device comprises a memory 702 and a processor 704, the memory 702 stores a computer program, and the processor 704 is configured to execute the steps in any of the above-mentioned method embodiments through the computer program. Figure 7

[0164] Optionally, in the present embodiment, the above-mentioned electronic device can be located in at least one of the network devices in the computer network.

[0165] Optionally, in the present embodiment, the above-mentioned processor can be configured to execute the method in each embodiment of the present application through the computer program.

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

[0167] ​​​​The memory 702 can be used to store software programs and modules, such as program instructions / modules corresponding to the method and device for determining a satellite orbit according to embodiments of the present application. The processor 704 executes various functions and data processing by running the software programs and modules stored in the memory 702, that is, implements the method for determining a satellite orbit as described above. The memory 702 can include a high-speed random access memory, and can further include a non-volatile memory, such as one or more magnetic storage devices, flash memories, or other non-volatile solid-state memories. In some examples, the memory 702 can further include a memory disposed remotely with respect to the processor 704, which can be connected to the terminal through a network. Examples of the network include, but are not limited to, the Internet, an intranet, a local area network, a mobile communication network, and a combination thereof. Specifically, the memory 702 can be, but is not limited to, used to store information carried by navigation signals and information carried by communication signals. As an example, as shown in Figure 7 The memory 702 can include, but is not limited to, the acquisition module 502, the fitting module 504, the integration module 506, the comparison module 508, and the determination module 510 in the device for determining a satellite orbit as described above. In addition, other module units in the device for determining a satellite orbit can also be included, which will not be described in detail in this example.

[0168] Optionally, the transmission device 706 is used to receive or send data via a network. Specific examples of the network can include wired networks and wireless networks. In an example, 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 to communicate with the Internet or a local area network. In an example, the transmission device 706 is a radio frequency (Radio Frequency, RF) module, which is used to communicate with the Internet in a wireless manner.

[0169] In addition, the electronic device further includes a display 708 for displaying the positioning information, and a connection bus 710 for connecting various module components in the electronic device.

[0170] In other embodiments, the terminal device or the server can be a node in a distributed system, where the distributed system can be a blockchain system, which can be a distributed system formed by the plurality of nodes communicating through a network. The nodes can form a peer-to-peer network, and any form of computing device, such as a server, a terminal, or an electronic device, can become a node in the blockchain system by joining the peer-to-peer network.

[0171] According to an aspect of the present application, a computer readable storage medium is provided, and a processor of an electronic device reads the computer instructions from the computer readable storage medium, and the processor executes the computer instructions to enable the electronic device to perform the satellite orbit determination method provided in the various optional implementations of the satellite orbit determination aspect.

[0172] Optionally, in the embodiment, the computer readable storage medium can be configured to store the computer instructions for executing the method in the embodiments of the present application.

[0173] Optionally, in the embodiment, a person skilled in the art can understand that all or part of the steps in the various methods of the above embodiments can be completed by programs instructing the hardware related to the terminal device, and the programs can be stored in a computer readable storage medium, and the storage medium can include a flash disk, a read-only memory (ROM), a random access memory (RAM), a magnetic disk or an optical disk, etc.

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

[0175] The integrated units in the above embodiments, if realized in the form of software function units and sold or used as independent products, can be stored in the computer readable storage medium. Based on such understanding, the technical solutions of the present application essentially or the parts that make contributions to the prior art or the whole or part of the technical solutions can be embodied in the form of software products, and the computer software products are stored in the storage medium, including a plurality of instructions for enabling one or more electronic devices to execute all or part of the steps of the methods described in the embodiments of the present application.

[0176] In the above embodiments of the present application, the description of each embodiment has its own focus, and the parts not described in detail in a certain embodiment can be referred to the related description of other embodiments.

[0177] In the several embodiments of the present application, it should be understood that the disclosed application programs can be implemented in other ways. Among them, the above-mentioned device embodiments are only schematic, for example, the division of the units is only a logical function division, and actual implementation can have another division manner, for example, a plurality of units or components can be combined or integrated into another system, or some features can be ignored or not executed. In addition, the coupling or direct coupling or communication connection between the shown or discussed mutual elements can be indirect coupling or communication connection through some interfaces, units or modules, and can be electrical or other forms.

[0178] The units described as separate components may or may not be physically separate, and the components displayed as units may or may not be physical units, that is, may be located in one place, or may be distributed to multiple network units. Part or all of the units can be selected according to actual needs to achieve the purpose of the embodiment scheme.

[0179] In addition, each functional unit in each embodiment of the present application can be integrated in one processing unit, or each unit can exist physically alone, or two or more units can be integrated in one unit. The integrated unit can be realized in the form of hardware or in the form of a software functional unit.

[0180] The above is only the preferred embodiment of the present application, and it should be pointed out that for ordinary skilled persons in the art, without departing from the principles of the present application, a number of improvements and refinements can be made, which should be considered as the protection scope of the present application.

Claims

1. A method for determining a satellite orbit, characterized in that: include: obtaining an initial satellite position of the target communication satellite when the target communication satellite is capable of receiving a navigation signal, and determining a reference orbit of the target communication satellite based on spatiotemporal information of the navigation satellite, wherein the spatiotemporal information includes the position and clock error of the navigation satellite, and the reference orbit is an orbit determined autonomously by the target communication satellite during autonomous orbit determination; 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 a predetermined conservative force model coefficient, the target satellite position, and the target non-conservative force model coefficient 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 times in the integrated orbit based on the position error sequence, wherein the position error sequence is used to indicate how the position deviation of the target communication satellite relative to the target orbit changes over time during the autonomous orbit determination process; In the case that the target communication satellite cannot receive the navigation signal, orbit filtering is performed using intersatellite link data to determine the target operating orbit.

2. The method according to claim 1, characterized in that The comparing the integrated orbit and the reference orbit to determine a position error sequence of the target communication satellite, and determining process noise covariance information at different integration times in the integrated orbit based on the position error sequence, includes: Obtaining the integral orbit point position corresponding to the integral orbit at any time, and obtaining the reference orbit point position corresponding to the reference orbit at any time; Determining error data corresponding to the target communication satellite at any time according to the integrated orbit point position and the reference orbit point position, wherein the error data includes at least one of position error data and clock error data; The position error sequence is generated based on the error data.

3. The method according to claim 1, characterized in that When the target communication satellite cannot receive the navigation signal, performing orbit filtering using intersatellite link data to determine the target orbit includes: In the case where the target communication satellite cannot receive the navigation signal, determining an initial operating state of the predicted orbit according to the initial orbit state quantity and initial process noise covariance information; Correcting the predicted orbit using the intersatellite link data according to a Kalman filter algorithm to obtain the target orbit; The intersatellite link data is used to determine an observation residual in the process of orbit filtering, and the observation residual represents a difference between an intersatellite ranging observation value indicated by the intersatellite link data and an expected intersatellite ranging value of the predicted orbit.

4. The method according to claim 1, wherein The method further comprises: determining, upon the navigation signal denial, that the target communications satellite is unable to receive the navigation signal; When the target communication satellite is in a state without ground support, it is determined that the target communication satellite is unable to receive the navigation signal.

5. The method according to claim 1, wherein The step of performing orbit integration on the target communication satellite according to the predetermined conservative force model coefficient, the target satellite position, and the target non-conservative force model coefficient to obtain an integrated orbit includes: Fitting the target non-conservative force model coefficient using the least squares method to determine the target non-conservative force fitting coefficient; Determining a dynamic model of the target communication satellite using the conservative force model coefficients and the target non-conservative force fitting coefficients; The state information of the target communication satellite at any moment is calculated based on the dynamic model and the target satellite position 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.

6. 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: Determining the initial non-conservative force model parameters according to the type of the target communication satellite; The initial satellite position, the reference orbit and the initial non-conservative force model parameters are fitted using a least squares method to obtain the target satellite position and the target non-conservative force model coefficients.

7. The method according to claim 6, characterized in that Determining the initial non-conservative force model parameters according to the type of the target communication satellite includes: In a case where the target communication satellite is of a low-orbit type, the initial non-conservative force model parameters include at least one of an atmospheric drag model parameter, a solar light pressure model parameter, and a magnetic influence parameter; In a case where the target communication satellite is of a medium-orbit type, the initial non-conservative force model parameters include at least one of the solar pressure model parameters, the lunar gravitational perturbation parameters, and the third-body gravitational model parameters; In the case where the target communication satellite is of a high-orbit type, the initial non-conservative force model parameters include at least one of the Earth's 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 comprises: The intersatellite 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, wherein the target communication satellite and the other communication satellites are in the same satellite constellation; 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 communication satellite and the target communication satellite when the any one communication satellite transmits a signal to the target communication satellite; The first range correction term is determined by the product of the speed of light and the clock deviation between the time when the target communication satellite transmits the signal and the time when any communication satellite receives the signal; The second range 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 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, relativistic effect, and atmospheric delay; The first observation residual term and the second observation residual term are both determined by observation noise generated in the process of acquiring the intersatellite link data.

9. The method according to claim 8, characterized in that The method further comprises: Determining the first geometric distance based on the position of the target communication satellite at the time of transmitting the signal and the position of any one of the communication satellites at the time of receiving the signal; determining the second geometric distance based on the position of the any one of the communication satellites at the time of transmitting the signal and the position of the target communication satellite at the time of receiving the signal; The first geometric distance is adjusted using at least one of the first range correction term, the first error correction term, and the first observation residual term, and the first geometric distance is normalized using the clock error of the predicted orbit generated in the orbit filtering; the second geometric distance is adjusted using at least one of the second range correction term, the second error correction term, and the second observation residual term, and the second geometric distance is normalized 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 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 the first target geometric distance and the second target geometric distance at a reference time, respectively; The target operating trajectory is determined 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, wherein the type of the target communication satellite is any one of a low orbit type, a medium orbit type, and a high orbit type.

11. The method according to claim 1, wherein When the target communication satellite cannot receive the navigation signal, performing orbit filtering using intersatellite link data to determine the target orbit includes: In a case where the target communication satellite cannot receive the navigation signal, obtaining a relative distance and a relative clock error between the target communication satellite and other communication satellites indicated by the intersatellite link data, wherein the target communication satellite and the other communication satellites are in the same satellite constellation; The orbit filtering is performed based on the relative distance and the relative clock error to determine the target orbit, wherein, during the orbit filtering process, an initial orbit state quantity is determined by the integrated orbit, and an initial process noise covariance is determined by the process noise covariance information.

12. A device for determining a satellite orbit, characterized in that: include: an acquisition module, configured to acquire an initial satellite position of a target communication satellite when the target communication satellite is capable of receiving a navigation signal, and determine a reference orbit of the target communication satellite based on spatiotemporal information of the navigation satellite, wherein the spatiotemporal information includes the position and clock error of the navigation satellite, and the reference orbit is an orbit determined by the target communication satellite during autonomous orbit determination; A fitting module is used 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 a target non-conservative force model coefficient; an integration module, configured to perform orbit integration on the target communication satellite according to a predetermined conservative force model coefficient, the target satellite position, and the target non-conservative force model coefficient, to determine an integrated orbit of the target communication satellite; a comparison module, configured to compare the integrated orbit with the reference orbit to determine a position error sequence of the target communication satellite, and determine process noise covariance information at different integration times in the integrated orbit based on the position error sequence, wherein the position error sequence is used to indicate how the position deviation of the target communication satellite relative to the target orbit changes over time during the autonomous orbit determination process; The determination module is used to use intersatellite link data to perform orbit filtering and determine the target 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, wherein the computer program can be executed by an electronic device to perform the method according to any one of claims 1 to 11.

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

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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