Beidou satellite maneuvering orbit recovery method

Abstract

The invention discloses a Beidou satellite maneuvering orbit recovery method. The method comprises the steps that S1, initial satellite state information is given; s2, based on the current satellite state information, detecting whether satellite orbital maneuver occurs in each epoch one by one; if yes, entering the step S4; if not, entering the step S3; s3, updating the state parameter values of all satellites, and returning to the step S2; and S4, processing the maneuvering orbit by adopting a combined short arc method orbit determination and packet filtering method to realize maneuvering orbit recovery. According to the method, rapid recovery of the maneuvering orbit of the Beidou satellite is realized by combining short arc method orbit determination and packet filtering, so that the navigation, positioning and time service performance of a Beidou system can be ensured under the condition that orbit maneuvering of the Beidou satellite occurs. According to the method, the orbit determination precision under the orbit maneuver condition of the satellite can be effectively improved, and the reliability of the global service capability of the Beidou No.3 system can be greatly improved.

Description

Technical Field

[0001] The invention belongs to the technical field of satellite orbit recovery, and in particular relates to a Beidou satellite maneuvering orbit recovery method combining short arc orbit determination and group filtering. Background Art

[0002] Due to gravitational disturbances from celestial bodies such as the Earth, the Sun and the Moon, as well as various other influences, satellites in orbit will gradually deviate from their intended positions. When the deviation exceeds an acceptable range, it is necessary to perform orbital maneuvers to maintain the nominal position of the satellite. During the maneuver, the position of the satellite may move by several kilometers. Due to the irregular thrust usually involved, the occurrence of orbital maneuvers will invalidate the previous dynamic model. In addition, it often takes a certain amount of time to accurately re-determine the orbit of the maneuvering satellite after the maneuver occurs. Orbital maneuvers will seriously affect the effectiveness of BDS satellite operations, resulting in a decrease in the accuracy of positioning, navigation and timing services for end users. In order to effectively ensure the service performance of the BDS-3 system and the availability of maneuvering satellites, it is crucial to achieve rapid recovery of the maneuvering orbit.

[0003] The current methods for rapid recovery of satellite maneuvering orbits mainly include maneuvering reconstruction, adaptive Kalman filtering, and kinematic orbit determination methods. Although a large number of studies have been conducted on satellite maneuvering orbit recovery, the currently available methods are mainly applicable to low-orbit satellites due to their rapid motion and favorable geometric conditions. Summary of the invention

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

[0005] In order to achieve the above-mentioned invention object, the technical solution adopted by the present invention is: a Beidou satellite maneuvering orbit recovery method, comprising the following steps:

[0006] S1, given initial satellite status information;

[0007] S2. Based on the current satellite status information, detect whether satellite orbit maneuvers occur in each epoch one by one;

[0008] If yes, 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. Use the combined short arc orbit determination and group filtering method to process the maneuvering orbit and realize maneuvering orbit recovery.

[0012] Furthermore, in step S1, the initial satellite state information includes state parameter values ​​of all satellites in the initial epoch, given a priori random information, the length of the short arc regular orbit period, and the maneuvering orbit deviation threshold;

[0013] The state parameter values ​​include the satellite's three-dimensional position, three-dimensional velocity, transceiver hardware delay, and solar pressure;

[0014] The given a priori random information includes a priori variances of all satellite state parameters, a priori variances of observation values, and a priori 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 orbit maneuver occurs in the first epoch comprises the following sub-steps:

[0017] S21, based on the given satellite initial state information, construct the satellite orbit state equation to update the filter time for all satellite orbits, and then determine the satellite prediction state information for the next epoch

[0018] S22, add the two-way observation values ​​between all satellite pairs to obtain the orbit determination observation value L of all satellite pairs (1) ;

[0019] S23. Based on satellite prediction status information and orbit determination observation value L (1) , calculate all orbits to determine the observed value L (1) The filtered innovation value Among them, A (1) The design matrix representing the observation equation;

[0020] S24, each filter innovation value Compare it with the maneuvering orbit deviation threshold, and store the filtered innovation value greater than the maneuvering orbit deviation threshold into the innovation set V0, and determine whether it is consistent with the maneuvering satellite SV i The set of all innovations about the observations Is it satisfied?

[0021] If yes, then the satellite orbit maneuver occurs and the process goes to step S4;

[0022] If not, no satellite orbit maneuver occurs, and the process goes to step S3.

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

[0024]

[0025]

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

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

[0028] S22-1. In the first epoch, arbitrarily select two satellites to form a satellite pair, including a first satellite and a second satellite;

[0029] S22-2, taking the observation value consisting of the ranging signal transmitted by the first satellite and the ranging signal received by the second satellite as the forward observation value;

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

[0031] S22-3, adding the forward observation value and the reverse observation value to obtain the orbit determination observation value between the satellite pairs;

[0032] S22-4. Repeat steps S22-1 and S22-3 to obtain orbit determination observation values ​​of all satellite pairs in 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] Where, L (1) represents all orbital determination observations, A (1) represents the design matrix of the observation equation, represents the satellite post-verification state parameter, K (1) represents the gain matrix of the filter, represents the satellite state prediction variance corresponding to all satellite forecast state information at time t1, Σ represents the prior variance of the observation value, Indicates the forecast status information of all satellites at time t1, represents the satellite posterior variance matrix, and I represents the identity matrix.

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

[0040] S41, within the pth to q-1th epochs where the orbital maneuver occurs, preliminarily determining the orbits of the maneuvering satellite and the non-maneuvering satellite respectively by using a group filtering method;

[0041] S42. Use the short arc method to accurately determine the orbit of a maneuvering satellite;

[0042] S43, for the qth epoch, re-adopting the grouping filtering method to respectively determine the orbits of the maneuvering satellite and the non-maneuvering satellite;

[0043] S44, calculating the innovation values ​​of all observation values ​​corresponding to the maneuvering satellite in the qth epoch, and determining whether there is at least one innovation value greater than the maneuvering orbit deviation threshold;

[0044] If yes, proceed to step S45;

[0045] If not, complete the maneuver track recovery;

[0046] S45. In the next epoch, continue to execute step S43 until the maneuvering orbit recovery is completed.

[0047] Furthermore, the step S41 is specifically as follows:

[0048] In the epochs p to q-1 where the orbital maneuver occurs, all observations related to the maneuvering satellite are taken as the first group, and the remaining observations are taken as the second group;

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

[0050] Furthermore, the step S42 is specifically as follows:

[0051] The three-dimensional position and velocity parameters of the maneuvering satellite at the pth epoch are solved by using the observation values ​​in the first group from the pth to the q-1th epochs, and the solved parameters are estimated by using the least squares batch processing to determine the orbit of the maneuvering satellite;

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

[0053]

[0054]

[0055] Where, L1 (p.q-1) represents the observation value of the maneuvering satellite group, A (p,q-1) The design matrix representing the observation values ​​of all maneuvering satellite groups from the pth to the q-1th epoch.

[0056] The beneficial effects of the present invention are:

[0057] 1) The method of the present invention can realize the rapid recovery of the maneuvering orbit only by relying on the inter-satellite two-way observation value, and can be completely independent of the ground observation station.

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

[0059] 3) The method of the present invention can restore the orbit accuracy of the maneuvering satellite to the level before the maneuver occurs within 12 hours. BRIEF DESCRIPTION OF THE DRAWINGS

[0060] Figure 1 A flow chart of the Beidou satellite maneuvering orbit recovery method provided by the present invention. DETAILED DESCRIPTION

[0061] The specific implementation modes of the present invention are described below so that those skilled in the art can understand the present invention. However, it should be clear that the present invention is not limited to the scope of the specific implementation modes. For those of ordinary skill in the art, as long as various changes are within the spirit and scope of the present invention as defined and determined by the attached claims, these changes are obvious, and all inventions and creations utilizing the concept of the present invention are protected.

[0062] The embodiment of the present invention provides a Beidou satellite maneuvering orbit recovery method, such as Figure 1 As shown, the following steps are included:

[0063] S1, given initial satellite status information;

[0064] S2. Based on the current satellite status information, detect whether satellite orbit maneuvers occur in each epoch one by one;

[0065] If yes, 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. Use the combined short arc orbit determination and group filtering method to process the maneuvering orbit and realize maneuvering orbit recovery.

[0069] In step S1 of the embodiment of the present invention, the initial satellite state information includes the state parameter values ​​of all satellites in the initial epoch Given a priori random information, the length of the short arc fixed orbit period dT and the maneuvering orbit deviation threshold v0;

[0070] Among them, the state parameter values ​​include the satellite's three-dimensional position, three-dimensional velocity, transmission and reception hardware delay, and solar pressure;

[0071] Given a priori random information including the prior variance of all satellite state parameters Observation prior variance S W and the prior process noise of the state equation; wherein the prior variance values ​​are all given artificially;

[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 the embodiment of the present invention, the method for detecting whether a satellite orbit maneuver occurs in the first epoch (from time t0 to time t1) includes the following sub-steps:

[0074] S21, based on the given satellite initial state information, construct the satellite orbit state equation to update the filter time for all satellite orbits, and then determine the satellite prediction state information for the next epoch

[0075] S22, add the two-way observation values ​​between all satellite pairs to obtain the orbit determination observation value L of all satellite pairs (1) ;

[0076] S23. Based on satellite prediction status information and orbit determination observation value L (1) , calculate all orbits to determine the observed value L (1) The filtered innovation value Among them, A (1) The design matrix representing the observation equation;

[0077] S24, each filter innovation value Compare it with the maneuvering orbit deviation threshold, and store the filtered innovation value greater than the maneuvering orbit deviation threshold into the innovation set V0, and determine whether it is consistent with the maneuvering satellite SV i The set of all innovations about the observations Is it satisfied?

[0078] If yes, then the satellite orbit maneuver occurs and the process goes to step S4;

[0079] If not, no satellite orbit maneuver occurs, and the process goes to step S3.

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

[0081]

[0082]

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

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

[0085] S22-1. In the first epoch, two satellites are randomly selected to form a satellite pair, including the first satellite SV A and the second satellite SV B ;

[0086] S22-2, the first satellite SV A Transmit ranging signal and second satellite SV B The observation value composed of the received ranging signal is taken as the forward observation value D AB ;

[0087] The second satellite SV B Transmit ranging signal and first satellite SV A The observation value composed of the received ranging signal is used as the reverse observation value D BA ;

[0088] S22-3, the forward observation value D AB and reverse observation D BA Add together to get the orbit 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 observation values ​​of all satellite pairs in the first epoch.

[0090] In this embodiment, “innovation” refers to the specific value of the filtered innovation, which is a professional term in Kalman filtering.

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

[0092]

[0093]

[0094]

[0095]

[0096] Where, L (1) represents all orbital determination observations, A (1) represents the design matrix of the observation equation, represents the satellite post-verification state parameter, K (1) represents the gain matrix of the filter, represents the satellite state prediction variance corresponding to all satellite forecast state information at time t1, Σ represents the prior variance of the observation value, Indicates the forecast status information of all satellites at time t1, represents the satellite posterior variance matrix, and I represents the identity matrix.

[0097] In step S4 of this embodiment, if If the pth epoch satellite SV is detected, the short arc orbit determination and group filtering method are used to process the maneuvering orbit. i An orbital maneuver occurs, and the initial time of this epoch is T1. T1 is the starting time of the short arc statutory orbit segment. In addition, the short arc statutory orbit segment is set as dT, and the end time of the short arc statutory orbit segment is T2 = T1 + dT. Since dT is an integer multiple of the orbit epoch time interval, T2 is set as the starting time of the qth epoch.

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

[0099] S41, within the pth to q-1th epochs where the orbital maneuver occurs, preliminarily determining the orbits of the maneuvering satellite and the non-maneuvering satellite respectively by using a group filtering method;

[0100] S42. Use the short arc method to accurately determine the orbit of a maneuvering satellite;

[0101] S43, for the qth epoch, re-adopting the grouping filtering method to respectively determine the orbits of the maneuvering satellite and the non-maneuvering satellite;

[0102] S44, calculating the innovation values ​​of all observation values ​​corresponding to the maneuvering satellite in the qth epoch, and determining whether there is at least one innovation value greater than the maneuvering orbit deviation threshold;

[0103] If yes, proceed to step S45;

[0104] If not, complete the maneuver track recovery;

[0105] S45. In the next epoch, continue to execute step S43 until the maneuvering orbit recovery is completed.

[0106] Step S41 in this embodiment is specifically as follows:

[0107] During the pth to q-1th epochs of orbital maneuvers, all SVs associated with the maneuvering satellites i The relevant observations are taken as the first group Obs1, and the remaining observations are taken as the second group Obs2;

[0108] The state parameters of the two groups of observations are solved respectively, and all the observations of the first group Obs1 within the time range are saved to determine the orbits of the maneuvering satellite and the non-maneuvering satellite.

[0109] The state parameters of the two groups of observation values ​​are solved according to the methods in the aforementioned steps S21, S22 and S24.

[0110] In this embodiment, step S42 is specifically as follows:

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

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

[0113]

[0114]

[0115] Where, L1 (p.q-1) represents the observation value of the maneuvering satellite group, A (p,q-1) The design matrix representing the observation values ​​of all maneuvering satellite groups from the pth to the q-1th epoch.

[0116] In this embodiment, the accuracy of the maneuvering satellite orbit determined by group filtering in step S41 is insufficient, so on this basis, the short arc method of S42 is used to further optimize the maneuvering satellite orbit to improve the accuracy. Only the maneuvering orbit information obtained in step S42 retains the posterior variance value of the maneuvering satellite state parameter obtained in S41.

[0117] The present invention uses specific embodiments to illustrate the principles and implementation methods of the present invention. The description of the above embodiments is only used to help understand the method of the present invention and its core idea. At the same time, for those skilled in the art, according to the idea of ​​the present invention, there will be changes in the specific implementation methods and application scope. In summary, the content of this specification should not be understood as a limitation on the present invention.

[0118] Those skilled in the art will appreciate that the embodiments described herein are intended to help readers understand the principles of the present invention, and should be understood that the protection scope of the present invention is not limited to such specific statements and embodiments. Those skilled in the art can make various other specific variations and combinations that do not deviate from the essence of the present invention based on the technical revelations disclosed by the present invention, and these variations and combinations are still within the protection scope of the present invention.

Claims

1. A Beidou satellite maneuvering orbit recovery method, characterized in that: The following steps are involved: S1, given initial satellite status information; S2. Based on the current satellite status information, detect whether satellite orbit maneuvers occur in each epoch one by one; If yes, proceed to step S4; If not, proceed to step S3; S3, update the status parameter values ​​of all satellites, and return to step S2; S4. Use the combined short arc orbit determination and group filtering method to process the maneuvering orbit and realize maneuvering orbit recovery.

2. The BeiDou satellite maneuvering orbit recovery method according to claim 1, characterized in that: In step S1, the initial satellite state information includes state parameter values ​​of all satellites in the initial epoch, given a priori random information, the length of the short arc orbit determination period, and the maneuvering orbit deviation threshold; The state parameter values ​​include the satellite's three-dimensional position, three-dimensional velocity, transceiver hardware delay, and solar pressure; The given a priori random information includes a priori variances of all satellite state parameters, a priori variances of observation values, and a priori process noise of the state equation; The length of the short-arc orbit determination period is an integer multiple of the orbit determination epoch time interval.

3. The Beidou satellite maneuvering orbit recovery method according to claim 2, characterized in that: In step S2, the method for detecting whether a satellite orbit maneuver occurs in the first epoch comprises the following sub-steps: S21, based on the given satellite initial state information, construct the satellite orbit state equation to update the filter time for all satellite orbits, and then determine the satellite prediction state information for the next epoch S22, add the two-way observation values ​​between all satellite pairs to obtain the orbit determination observation value L of all satellite pairs (1) ; S23. Based on satellite prediction status information and orbit determination observation value L (1) , calculate all orbits to determine the observed value L (1) The filtered innovation value Among them, A (1) The design matrix representing the observation equation; S24, each filter innovation value Compare it with the maneuvering orbit deviation threshold, and store the filtered innovation value greater than the maneuvering orbit deviation threshold into the innovation set V0, and determine whether it is consistent with the maneuvering satellite SV i The set of all innovations about the observations Is it satisfied? If yes, then the satellite orbit maneuver occurs and the process goes to step S4; If not, no satellite orbit maneuver occurs, and the process goes to step S3.

4. The Beidou satellite maneuvering orbit recovery method according to claim 3 is characterized in that: In step S21, the satellite orbit state equation constructed is: In the formula, represents the forecast status information of all satellites at time t1, Φ 1,0 represents the state transfer matrix, represents the state parameter values ​​of all satellites in the initial epoch, express Corresponding to the satellite prediction state variance value, represents the prior variance of all satellite state parameters, Represents Φ 1,0 The transposed matrix, Σ W represents the prior process variance of the state equation.

5. The Beidou satellite maneuvering orbit recovery method according to claim 3, characterized in that: The step S22 comprises the following sub-steps: S22-1. In the first epoch, arbitrarily select two satellites to form a satellite pair, including a first satellite and a second satellite; S22-2, taking the observation value consisting of the ranging signal transmitted by the first satellite and the ranging signal received by the second satellite as the forward observation value; Taking the observation value composed of the ranging signal transmitted by the second satellite and the ranging signal received by the first satellite as the reverse observation value; S22-3, adding the forward observation value and the reverse observation value to obtain the orbit determination observation value between the satellite pairs; S22-4. Repeat steps S22-1 and S22-3 to obtain orbit determination observation values ​​of all satellite pairs in the first epoch.

6. The Beidou satellite maneuvering orbit recovery method according to claim 3, characterized in that: In step S3, the formula for updating the state parameter values ​​of all satellites is: Where, L (1) represents all orbital determination observations, A (1) represents the design matrix of the observation equation, represents the satellite post-verification state parameter, K (1) represents the gain matrix of the filter, represents the satellite state prediction variance corresponding to all satellite forecast state information at time t1, Σ represents the prior variance of the observation value, Indicates the forecast status information of all satellites at time t1, represents the satellite posterior variance matrix, and I represents the identity matrix.

7. The Beidou satellite maneuvering orbit recovery method according to claim 1, characterized in that: The step S4 comprises the following sub-steps: S41, within the pth to q-1th epochs where the orbital maneuver occurs, preliminarily determining the orbits of the maneuvering satellite and the non-maneuvering satellite respectively by using a group filtering method; S42. Use the short arc method to accurately determine the orbit of a maneuvering satellite; S43, for the qth epoch, re-adopting the grouping filtering method to respectively determine the orbits of the maneuvering satellite and the non-maneuvering satellite; S44, calculating the innovation values ​​of all observation values ​​corresponding to the maneuvering satellite in the qth epoch, and determining whether there is at least one innovation value greater than the maneuvering orbit deviation threshold; If yes, proceed to step S45; If not, complete the maneuver track recovery; S45. In the next epoch, continue to execute step S43 until the maneuvering orbit recovery is completed.

8. The Beidou satellite maneuvering orbit recovery method according to claim 7, characterized in that: The step S41 is specifically as follows: In the epochs p to q-1 where the orbital maneuver occurs, all observations related to the maneuvering satellite are taken as the first group, and the remaining observations are taken as the second group; The state parameters of the two groups of observations are solved respectively, and all the observations in the first group within the time range are saved to determine the orbits of the maneuvering satellite and the non-maneuvering satellite.

9. The Beidou satellite maneuvering orbit recovery method according to claim 8, characterized in that: The step S42 is specifically as follows: The three-dimensional position and velocity parameters of the maneuvering satellite at the pth epoch are solved by using the observation values ​​in the first group from the pth to the q-1th epochs, and the solved parameters are estimated by using the least squares batch processing to determine the orbit of the maneuvering satellite; Among them, the estimated three-dimensional position and velocity parameters It is expressed as: Where L1 (p.q-1) represents the observation value of the maneuvering satellite group, A (p,q-1) The design matrix representing the observation values ​​of all maneuvering satellite groups from the pth to the q-1th epoch.

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