A method for recovering a Beidou satellite maneuvering orbit

Abstract

The application discloses a Beidou satellite maneuvering orbit recovery method, which comprises the following steps: S1, giving initial satellite state information; S2, detecting whether satellite orbit maneuvering occurs for each ephemeris based on current satellite state information; if yes, entering step S4; if no, entering step S3; S3, updating state parameter values of all satellites and returning to step S2; S4, processing the maneuvering orbit by adopting combined short-arc orbit determination and grouping filtering method, and realizing maneuvering orbit recovery. The combined short-arc orbit determination and grouping filtering method realizes quick recovery of the Beidou satellite maneuvering orbit, so as to guarantee the navigation, positioning and timing service performance of the Beidou system in the case that the Beidou satellite occurs orbit maneuvering. The method can not only effectively improve the orbit determination precision in the case that the satellite occurs orbit maneuvering, but also greatly improve the reliability of the global service capability of the Beidou No. 3 system.

Description

Technical Field

[0001] This invention belongs to the field of satellite orbit recovery technology, specifically relating to a method for recovering the maneuvering orbit of BeiDou satellites by combining short-arc orbit determination and group filtering. Background Technology

[0002] Due to gravitational disturbances from celestial bodies such as the Earth, Sun, and Moon, as well as various other influences, satellites in orbit will gradually deviate from their intended positions. When the deviation exceeds acceptable limits, it becomes necessary to perform orbital maneuvers to maintain the satellite's nominal position. During a maneuver, the satellite's position may shift by several kilometers. Due to the irregular thrust typically involved, orbital maneuvers invalidate previous dynamic models. Furthermore, it often takes time to accurately re-determine the maneuvered satellite's orbit after a maneuver. Orbital maneuvers severely impact the effectiveness of BDS satellite operations, leading to a decrease in the accuracy of end-user positioning, navigation, and timing services. To effectively ensure the service performance of the BDS-3 system and the availability of maneuvered satellites, rapid recovery of the maneuvered orbit is crucial.

[0003] Currently, the main methods used for rapid recovery of satellite maneuvering orbits include maneuvering reconstruction, adaptive Kalman filtering, and kinematic orbit determination methods. Although a large amount of research has been conducted on satellite maneuvering orbit recovery, due to the rapid motion and favorable geometric conditions of low Earth orbit (LEO) satellites, the available methods are primarily applicable to LEO satellites. Summary of the Invention

[0004] In view of the above-mentioned shortcomings in the prior art, the BeiDou satellite maneuver orbit recovery method provided by the present invention solves the problem that the existing BeiDou-3 system cannot quickly restore the accuracy of the maneuver orbit during the precise orbit determination process.

[0005] To achieve the above-mentioned objectives, the technical solution adopted by this invention is: a method for restoring the maneuvering orbit of a BeiDou satellite, comprising the following steps:

[0006] S1. Given initial satellite state information;

[0007] S2. Based on the current satellite status information, detect whether satellite orbital maneuvers have occurred at each epoch;

[0008] If so, proceed to step S4;

[0009] If not, proceed to step S3;

[0010] S3. Update the status parameter values ​​of all satellites and return to step S2;

[0011] S4. The combined short-arc tracing method and group filtering method are used to process the maneuver track to achieve track restoration.

[0012] Furthermore, in step S1, the initial satellite state information includes the state parameter values ​​of all satellites in the initial epoch, the given prior random information, the short arc orbital period length, and the maneuver orbit deviation threshold.

[0013] The state parameter values ​​include the satellite's three-dimensional position, three-dimensional velocity, transmit / receive hardware delay, and solar radiation pressure.

[0014] The given prior random information includes the prior variance of all satellite state parameters, the prior variance of observations, and the prior process noise of the state equation.

[0015] The length of the short arc orbit determination period is an integer multiple of the orbit determination epoch time interval.

[0016] Furthermore, in step S2, the method for detecting whether a satellite orbital maneuver has occurred in the first epoch includes the following sub-steps:

[0017] S21. Based on the given initial satellite state information, construct the satellite orbit state equation, update the orbit determination filter time for all satellites, and then determine the predicted satellite state information for the next epoch.

[0018] S22. Sum the two-way observations between all satellite pairs to obtain the orbit determination observations L for all satellite pairs. (1) ;

[0019] S23. Based on satellite prediction status information Determine the observation value L based on the orbit (1) Calculate all orbits to determine the observation value L (1) Filtered innovation value Among them, A (1) The design matrix represents the observation equation;

[0020] S24. Convert each filtered information value The filtered innovation value is compared with the maneuvering trajectory deviation threshold, and those values ​​exceeding the threshold are stored in the innovation set V0. The result is then compared with the maneuvering satellite SV. i The set of all new information about the observations Does it meet the requirements?

[0021] If so, a satellite orbital maneuver will occur, and the process will proceed to step S4;

[0022] If not, no satellite orbital maneuver has occurred, proceed to step S3.

[0023] Furthermore, in step S21, the constructed satellite orbital state equation is:

[0024]

[0025]

[0026] In the formula, Φ represents the forecast status information of all satellites at time t1. 1,0 Represents the state transition matrix. This represents the state parameter values ​​of all satellites in the initial epoch. express Corresponding satellite prediction state variance value, This represents the prior variance of all satellite state parameters. Φ 1,0 The transpose matrix, Σ W This represents the variance of the prior process of the state equation.

[0027] Furthermore, step S22 includes the following sub-steps:

[0028] S22-1. Within the first epoch, any two satellites are selected to form a satellite pair, including the first satellite and the second satellite.

[0029] S22-2, The observation value composed of the ranging signal transmitted by the first satellite and the ranging signal received by the second satellite is taken as the positive observation value;

[0030] The observation value composed of the ranging signal transmitted by the second satellite and the ranging signal received by the first satellite is used as the reverse observation value;

[0031] S22-3. Add the forward and reverse observations to obtain the orbit determination observations between satellite pairs;

[0032] S22-4. Repeat steps S22-1 and S22-3 to obtain orbit determination observations for all satellite pairs within the first epoch.

[0033] Furthermore, in step S3, the formula for updating the state parameter values ​​of all satellites is:

[0034]

[0035]

[0036]

[0037]

[0038] In the formula, L (1) Indicates all orbital determination observations, A (1) The design matrix represents the observation equation. K represents the post-hoc state parameters of the satellite. (1) This represents the gain matrix of the filter. This represents the variance of the predicted satellite state at time t1, where Σ represents the prior variance of the observed values. This represents the forecast status information of all satellites at time t1. Let I represent the posterior variance matrix of the satellite, and let I represent the identity matrix.

[0039] Furthermore, step S4 includes the following sub-steps:

[0040] S41. During the p-th to q-1th epochs in which the orbital maneuver occurs, the orbits of the maneuvering satellite and the non-maneuvering satellite are initially determined using a group filtering method.

[0041] S42. The short-arc method is used to accurately determine the orbit of the maneuvering satellite;

[0042] S43. For the q-th epoch, the group filtering method is used again to determine the orbits of the maneuvering satellite and the non-maneuvering satellite respectively;

[0043] S44. Calculate the innovation value of all observations corresponding to the maneuvering satellite within the q-th epoch, and determine whether there is at least one innovation value greater than the maneuvering orbit deviation threshold.

[0044] If so, proceed to step S45;

[0045] If not, then complete the restoration of the maneuver track;

[0046] S45. In the next epoch, continue executing step S43 until the maneuver orbit is restored.

[0047] Further, step S41 specifically includes:

[0048] Within epths p to q-1 of the orbital maneuver, all observations related to the maneuvering satellite are grouped into the first group, and the remaining observations are grouped into the second group.

[0049] The state parameters are solved for the two sets of observations respectively, and all the observations in the first set within the time range are saved to determine the orbits of the maneuvering satellite and the non-maneuvering satellite.

[0050] Further, step S42 specifically includes:

[0051] The three-dimensional position and velocity parameters of the maneuvering satellite in the p-th epoch are solved using the observations in the first group from the p-th to the q-1-th epochs. The least squares batch processing is then used to estimate the solved parameters, thereby determining the orbit of the maneuvering satellite.

[0052] Among them, the estimated three-dimensional position and velocity parameters Represented as:

[0053]

[0054]

[0055] In the formula, L1 (p.q-1) A represents the observations of the mobile satellite group. (p,q-1) The design matrix represents the grouped observations of all maneuvering stars from epoch p to q-1.

[0056] The beneficial effects of this invention are as follows:

[0057] 1) The method of the present invention can achieve rapid recovery of maneuver orbits by relying solely on two-way inter-satellite observations, without relying on ground observation stations at all.

[0058] 2) When satellite orbital maneuvers occur, the method of this invention can ensure that the orbit determination accuracy of non-motorized satellites is completely unaffected by the maneuvering orbit.

[0059] 3) The method of the present invention can restore the orbital accuracy of a maneuvering satellite to the level before the maneuver occurred within 12 hours. Attached Figure Description

[0060] Figure 1 The flowchart of the BeiDou satellite maneuver orbit restoration method provided by the present invention. Detailed Implementation

[0061] The specific embodiments of the present invention are described below to enable those skilled in the art to understand the present invention. However, it should be understood that the present invention is not limited to the scope of the specific embodiments. For those skilled in the art, various changes are obvious as long as they are within the spirit and scope of the present invention as defined and determined by the appended claims. All inventions utilizing the concept of the present invention are protected.

[0062] This invention provides a method for restoring the maneuvering orbit of a BeiDou satellite, such as... Figure 1 As shown, it includes the following steps:

[0063] S1. Given initial satellite state information;

[0064] S2. Based on the current satellite status information, detect whether satellite orbital maneuvers have occurred at each epoch;

[0065] If so, proceed to step S4;

[0066] If not, proceed to step S3;

[0067] S3. Update the status parameter values ​​of all satellites and return to step S2;

[0068] S4. The combined short-arc tracing method and group filtering method are used to process the maneuver track to achieve track restoration.

[0069] In step S1 of this embodiment of the invention, the initial satellite state information includes the state parameter values ​​of all satellites at the initial epoch. Given prior random information, the short-arc fixed orbit time period length dT, and the maneuver orbit deviation threshold v0;

[0070] Among them, the status parameter values ​​include the satellite's three-dimensional position, three-dimensional velocity, transmit / receive hardware delay, and solar radiation pressure;

[0071] Given prior random information including the prior variance of all satellite state parameters Prior variance S of observations W And the prior process noise of the state equation; where the prior variance values ​​are all given by humans;

[0072] The length of the short arc orbit determination period is an integer multiple of the orbit determination epoch time interval.

[0073] In step S2 of this embodiment of the invention, the method for detecting whether a satellite orbital maneuver has occurred in the first epoch (from time t0 to time t1) includes the following sub-steps:

[0074] S21. Based on the given initial satellite state information, construct the satellite orbit state equation, update the orbit determination filter time for all satellites, and then determine the predicted satellite state information for the next epoch.

[0075] S22. Sum the two-way observations between all satellite pairs to obtain the orbit determination observations L for all satellite pairs. (1) ;

[0076] S23. Based on satellite prediction status information Determine the observation value L based on the orbit (1) Calculate all orbits to determine the observation value L (1) Filtered innovation value Among them, A (1) The design matrix represents the observation equation;

[0077] S24. Convert each filtered information value The filtered innovation value is compared with the maneuvering trajectory deviation threshold, and those values ​​exceeding the threshold are stored in the innovation set V0. The result is then compared with the maneuvering satellite SV. i The set of all new information about the observations Does it meet the requirements?

[0078] If so, a satellite orbital maneuver will occur, and the process will proceed to step S4;

[0079] If not, no satellite orbital maneuver has occurred, proceed to step S3.

[0080] In step S21 of this embodiment, the constructed satellite orbital state equation is:

[0081]

[0082]

[0083] In the formula, Φ represents the forecast status information of all satellites at time t1. 1,0 Represents the state transition matrix. This represents the state parameter values ​​of all satellites in the initial epoch. express Corresponding satellite prediction state variance value, This represents the prior variance of all satellite state parameters. Φ 1,0 The transpose matrix, Σ W This represents the variance of the prior process of the state equation.

[0084] In this embodiment, step S22 includes the following sub-steps:

[0085] S22-1. Within the first epoch, any two satellites are selected to form a satellite pair, including the first satellite SV. A Second satellite SV B ;

[0086] S22-2, the first satellite SV A Transmit ranging signal and second satellite SV B The observations composed of the received ranging signals are used as the forward observations D. AB ;

[0087] The second satellite SV B Transmit ranging signal and first satellite SV A The observations composed of the received ranging signals are used as the reverse observations D. BA ;

[0088] S22-3, The positive observation value D AB and reverse observation D BA Adding them together, we obtain the orbital determination observation value D between the satellite pairs. AB +D BA ;

[0089] S22-4. Repeat steps S22-1 and S22-3 to obtain orbit determination observations for all satellite pairs within the first epoch.

[0090] In this embodiment, "innovation" refers to the specific numerical value of the filtered innovation, which is a technical term in Kalman filtering.

[0091] In step S3 of this embodiment of the invention, the formula for updating the state parameter values ​​of all satellites is:

[0092]

[0093]

[0094]

[0095]

[0096] In the formula, L (1) Indicates all orbital determination observations, A (1) The design matrix represents the observation equation. K represents the post-hoc state parameters of the satellite. (1) This represents the gain matrix of the filter. This represents the variance of the predicted satellite state at time t1, where Σ represents the prior variance of the observed values. This represents the forecast status information of all satellites at time t1. Let I represent the posterior variance matrix of the satellite, and let I represent the identity matrix.

[0097] In step S4 of this embodiment, if If established, the short-arc orbit determination method and group filtering method are used to process the maneuver orbit; assuming the p-th epoch satellite SV is detected. i An orbital maneuver occurs, and the initial time of this epoch is T1. T1 is the starting time of the short-arc fixed orbital segment. Furthermore, if the short-arc fixed orbital segment is defined as dT, then the ending time of the short-arc fixed orbital segment is T2 = T1 + dT. Since dT is an integer multiple of the orbital epoch time interval, let T2 be the starting time of the q-th epoch.

[0098] Based on this, step S4 of the present invention includes the following sub-steps:

[0099] S41. During the p-th to q-1th epochs in which the orbital maneuver occurs, the orbits of the maneuvering satellite and the non-maneuvering satellite are initially determined using a group filtering method.

[0100] S42. The short-arc method is used to accurately determine the orbit of the maneuvering satellite;

[0101] S43. For the q-th epoch, the group filtering method is used again to determine the orbits of the maneuvering satellite and the non-maneuvering satellite respectively;

[0102] S44. Calculate the innovation value of all observations corresponding to the maneuvering satellite within the q-th epoch, and determine whether there is at least one innovation value greater than the maneuvering orbit deviation threshold.

[0103] If so, proceed to step S45;

[0104] If not, then complete the restoration of the maneuver track;

[0105] S45. In the next epoch, continue executing step S43 until the maneuver orbit is restored.

[0106] In this embodiment, step S41 specifically involves:

[0107] During epth to q-1 of the orbital maneuver, all satellites related to the maneuvering satellite SV will be... i The relevant observations are designated as the first group Obs1, and the remaining observations are designated as the second group Obs2;

[0108] The state parameters are solved for the two sets of observations respectively, and all the observations in the first set of Obs1 within the time range are saved to determine the orbits of the motorized and non-motorized satellites.

[0109] Specifically, the state parameters of the two sets of observations are solved using the methods described in steps S21, S22, and S24 above.

[0110] In this embodiment, step S42 specifically includes:

[0111] Using the observations L1 from the first group of Obs1 in the p-th to q-1th epochs (p,q-1) Solve for the three-dimensional position and velocity parameters of the maneuvering satellite at the p-th epoch. The least squares batch processing method is used to estimate the solution parameters, thereby determining the orbit of the maneuvering satellite; in this process, the maneuvering satellite SV i The light pressure parameter is considered a known value, and the estimated value of the first epoch is used. The state parameters of other satellites are also considered known values, and the estimated values ​​obtained in step S41 are used.

[0112] Among them, the estimated three-dimensional position and velocity parameters Represented as:

[0113]

[0114]

[0115] In the formula, L1 (p.q-1) A represents the observations of the mobile satellite group. (p,q-1) The design matrix represents the grouped observations of all maneuvering stars from epoch p to q-1.

[0116] In this embodiment, the accuracy of the maneuvering star orbit determined by the group filtering in step S41 is insufficient. Therefore, the short arc method in step S42 is used to further optimize the maneuvering star orbit to improve the accuracy. Only the posterior variance of the maneuvering star state parameters obtained in step S41 is retained in the maneuvering orbit information obtained in step S42.

[0117] Specific embodiments have been used to illustrate the principles and implementation methods of this invention. The descriptions of the embodiments above are only for the purpose of helping to understand the method and core ideas of this invention. At the same time, for those skilled in the art, there will be changes in the specific implementation methods and application scope based on the ideas of this invention. Therefore, the content of this specification should not be construed as a limitation of this invention.

[0118] Those skilled in the art will recognize that the embodiments described herein are intended to help the reader understand the principles of the invention, and should be understood that the scope of protection of the invention is not limited to such specific statements and embodiments. Those skilled in the art can make various other specific modifications and combinations based on the technical teachings disclosed in this invention without departing from the spirit of the invention, and these modifications and combinations are still within the scope of protection of this invention.

Claims

1. A method for recovering a Beidou satellite maneuvering orbit, characterized in that, The method comprises the following steps: S1, providing initial satellite state information; S2, detecting whether a satellite orbit maneuver occurs at each epoch based on current satellite state information; If yes, go to step S4; If no, go to step S3; S3, updating state parameter values of all satellites and returning to step S2; S4, processing the maneuvering orbit by using a joint short-arc method orbit determination and grouping filtering method to realize maneuvering orbit recovery; In step S1, the initial satellite state information comprises state parameter values of all satellites at an initial epoch, given prior random information, a short-arc method orbit determination time interval length, and a maneuvering orbit deviation threshold value; The state parameter values comprise satellite three-dimensional positions, three-dimensional velocities, transceiver hardware delays, and solar radiation pressure; The given prior random information comprises prior variances of all satellite state parameters, prior variances of observation values, and prior process noises of state equations; The short-arc method orbit determination time interval length is an integer multiple of an orbit determination epoch time interval; In step S2, the method for detecting whether a satellite orbit maneuver occurs at a first epoch comprises the following steps: S21, based on the given satellite initial state information, constructing satellite orbit state equation to determine orbit filtering time update of all satellites, and further determining satellite predicted state information of next ephemeris ; S22, adding all the bi-directional observations between satellites to obtain orbit determination observations of all satellite pairs ; S23. computing a filter innovation value for all orbit determination observations and satellite predicted state information ; wherein represents a design matrix of the observation equation;​​ S24, store each filtered innovation value in the innovation set compared to the maneuver orbit bias threshold and the filtered innovation values greater than the maneuver orbit bias threshold are stored in the innovation set and determine whether the maneuver satellite related to the observation satisfies ;​ If yes, a satellite orbit maneuver occurs, and go to step S4; If no, a satellite orbit maneuver does not occur, and go to step S3; In step S3, the formula for updating state parameter values of all satellites is: wherein, represents all the orbit-determined observations, represents the design matrix of the observation equation, represents the satellite a posteriori state parameters, represents the gain matrix of the filter, represents represents the satellite state prediction variance values corresponding to all the satellite prediction state information at the time instant, represents the observation value prior variance, represents represents all the satellite prediction state information at the time instant, represents the satellite a posteriori variance matrix, represents the identity matrix; Step S4 comprises the following steps: S41、In the first to the epoch, the orbit of the maneuvering satellite and the non-maneuvering satellite is preliminarily determined respectively by using the grouping filtering method. S42, accurately determining the orbit of a maneuvering satellite by using a short-arc method; S43、For the first epoch, re-adopt the grouping filtering method to determine the orbit of the maneuverable satellite and the non-maneuverable satellite respectively; S44, Calculate the... Within each epoch, the information values ​​of all observations corresponding to the maneuvering satellite are determined, and it is determined whether there is at least one information value greater than the maneuvering orbit deviation threshold. If yes, go to step S45; If no, complete maneuvering orbit recovery; S45, in a next epoch, continue to execute step S43 until maneuvering orbit recovery is completed.

2. The method of claim 1, wherein the BeiDou satellite maneuver orbit recovery is characterized by, In step S21, the constructed satellite orbit state equation is: wherein represents all satellite predicted state information at time represents state transition matrix, represents state parameter values of all satellites at initial epoch, represents corresponding satellite predicted state variance values, represents all satellite state parameter prior variances, represents transpose of represents state equation prior process variance.

3. The method of claim 1, wherein, Step S22 comprises the following steps: S22-1, in a first epoch, arbitrarily selecting two satellites to form a satellite pair, comprising a first satellite and a second satellite; S22-2, taking observation values formed by a first satellite transmitting a ranging signal and a second satellite receiving the ranging signal as forward observation values; Taking observation values formed by the second satellite transmitting a ranging signal and the first satellite receiving the ranging signal as reverse observation values; S22-3, adding the forward observation values and the reverse observation values to obtain orbit determination observation values between the satellite pair; S22-4, repeating steps S22-1 and S22-3 to obtain orbit determination observation values of all satellite pairs in the first epoch.

4. The method of claim 1, wherein, Step S41 is specifically: The first to the epochs, all observations related to the maneuvering satellite are taken as the first set, and the remaining observations are taken as the second set; Solve for the state parameters of the two sets of observations respectively, and then... To the All observations in the first group within each epoch are preserved, thereby determining the orbits of both motorized and non-motorized satellites.

5. The method of claim 4, wherein, Step S42 is specifically: Using the first To the The observations in the first group of epochs are used to solve for the maneuvering satellite in the epoch. The three-dimensional position and velocity parameters of each epoch are obtained, and the least squares batch processing is used to estimate the solution parameters, thereby determining the orbit of the maneuvering satellite; wherein the estimated three-dimensional position and velocity parameters are represented as: wherein denotes the observations of the maneuvering satellite group, denotes the observations of the maneuvering satellite group at the denotes the observations of the maneuvering satellite group at the denotes the design matrix of all observations of the maneuvering satellite group at the first to the last epoch.

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