Adaptive identification method and system for absolute navigation filtering

Abstract

The invention discloses a self-adaptive identification method and system for absolute navigation filtering, and the method carries out the parameter judgment of the calculation result of each step in real time in the calculation process, avoids the introduction of wrong input observed quantity data, guarantees the real-time performance and reliability of an attitude and orbit control system, and improves the recognition accuracy of the attitude and orbit control system. Multiple times of adaptive identification of absolute navigation filtering in single shot time are realized, and the correctness of formation control absolute navigation filtering results is improved.

Description

Technical Field

[0001] The invention belongs to the technical field of satellite formation navigation and control, and in particular relates to an adaptive identification method and system for absolute navigation filtering. Background Art

[0002] Spacecraft formation flying is a new spacecraft space operation mode that emerged in the late 1980s with the development of micro-satellites. Compared with a single spacecraft, satellite formation flying has very prominent advantages and has been favored by major space powers around the world since the concept was born.

[0003] In order to ensure the formation configuration, it is necessary to introduce the control calculation method of absolute navigation and relative navigation, and absolute navigation filtering is an indispensable part of it. At present, when calculating the absolute navigation filter of the attitude and orbit control system, the observation quantity is directly introduced as input to complete the filter calculation of absolute navigation. However, since the observation quantity is introduced for calculation in each shot, once the observation quantity has a large deviation, the calculated orbital parameter mean instant root error will occur, which will affect the calculation of the absolute navigation of the satellite, and also affect the orbit control of other satellites and formation control. Summary of the invention

[0004] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art, to provide an adaptive identification method and system for absolute navigation filtering, to realize a multiple adaptive cycle identification method for absolute navigation filtering under a single shot, and to improve the correctness and reliability of the absolute navigation EKF filtering calculation results.

[0005] The objective of the present invention is achieved through the following technical solutions: an adaptive identification method for absolute navigation filtering, comprising: step S1: setting the current beat filter abnormality flag ErrFlag to 1, and setting the current beat abnormality count ErrCnt to 0; step S2: if the current beat filter observation is valid, then calculating the modulus of the difference between the observation position and the previous beat filter output position in the J2000 coordinate system, if the position difference modulus meets the preset threshold range, then the current beat filter abnormality flag is set ErrFlag to 0; step S3: if the current beat abnormality count ErrCnt = 0, then using the orbital time as At the filtering moment, the current beat filter output state variable and the state error covariance matrix are obtained according to the filtering moment, the previous beat filter output state variable and the state error covariance matrix; if the current beat abnormal count ErrCnt is not equal to 0, the filtering moment is recalculated, and the current beat filter output state variable and the state error covariance matrix are obtained according to the recalculated filtering moment, the previous beat filter output state variable and the state error covariance matrix; step S4: obtain the position and speed according to the current beat filter output state variable and the state error covariance matrix, and obtain the position modulus, the speed modulus, the position and speed cross multiplication modulus according to the position and speed Value and the transfer matrix from the J2000 coordinate system to the orbital coordinate system. If the position modulus does not meet the preset position modulus threshold range or the speed V modulus does not meet the preset speed modulus threshold range or the position and speed cross multiplication modulus does not meet the preset cross multiplication threshold range, then set ErrFlag to 1 and return to step S1, otherwise ErrFlag is 0; step S5: according to the position speed, position modulus, speed modulus, position and speed cross multiplication modulus and the transfer matrix from the J2000 coordinate system to the orbital coordinate system, obtain the instantaneous root of the filtering time. If the instantaneous root of the filtering time does not meet the preset instantaneous root threshold range of the filtering time, Then set ErrFlag to 1 and return to step S1, otherwise ErrFlag is 0; step S6: obtain the instantaneous root of the orbit calculation time according to the instantaneous root of the filtering time. If the instantaneous root of the orbit calculation time does not meet the preset instantaneous root threshold range of the orbit calculation time, set ErrFlag to 1 and return to step S1, otherwise ErrFlag is 0; step S7: obtain the flat root of the orbit calculation time according to the instantaneous root of the orbit calculation time. If the flat root of the orbit calculation time does not meet the preset flat root threshold range of the orbit calculation time, set ErrFlag to 1 and return to step S1, otherwise ErrFlag is 0.

[0006] In the above-mentioned adaptive identification method of absolute navigation filtering, the recalculated filtering time is the previous filtering time plus the single-shot time.

[0007] In the above-mentioned adaptive identification method of absolute navigation filtering, the transfer matrix from the J2000 coordinate system to the orbital coordinate system is:

[0008]

[0009] Among them, i is the instantaneous root of the satellite's orbital inclination; Ω is the instantaneous root of the satellite's ascending node's right ascension; u is the instantaneous root of the satellite's latitude argument.

[0010] In the above-mentioned adaptive identification method of absolute navigation filtering, the instantaneous root at the filtering moment includes the semi-major axis at the filtering moment, the eccentricity at the filtering moment and the orbital inclination at the filtering moment.

[0011] In the above-mentioned adaptive identification method of absolute navigation filtering, the instantaneous root at the filtering moment does not satisfy the preset instantaneous root threshold range at the filtering moment: the semi-major axis at the filtering moment does not satisfy the preset semi-major axis threshold range at the filtering moment, or the eccentricity at the filtering moment does not satisfy the preset eccentricity threshold range at the filtering moment, or the orbital inclination at the filtering moment does not satisfy the preset orbital inclination threshold range at the filtering moment.

[0012] In the above-mentioned adaptive identification method of absolute navigation filtering, the instantaneous root of orbit calculation time includes the semi-major axis of orbit calculation time, the eccentricity of orbit calculation time and the orbit inclination of orbit calculation time.

[0013] In the above-mentioned adaptive identification method of absolute navigation filtering, the instantaneous root of the orbit calculation moment does not satisfy the preset instantaneous root threshold range of the orbit calculation moment when: the semi-major axis of the orbit calculation moment does not satisfy the preset semi-major axis threshold range of the orbit calculation moment, or the eccentricity of the orbit calculation moment does not satisfy the preset eccentricity threshold range of the orbit calculation moment, or the orbit inclination of the orbit calculation moment does not satisfy the preset orbit inclination threshold range of the orbit calculation moment.

[0014] In the above-mentioned adaptive identification method of absolute navigation filtering, the mean root of orbit calculation time includes the mean semi-major axis of orbit calculation time, the mean eccentricity of orbit calculation time and the mean orbit inclination of orbit calculation time.

[0015] In the above-mentioned adaptive identification method of absolute navigation filtering, the flat root at the orbit calculation time does not satisfy the preset flat root threshold range at the orbit calculation time when: the flat semi-major axis at the orbit calculation time does not satisfy the preset flat semi-major axis threshold range at the orbit calculation time, or the flat eccentricity at the orbit calculation time does not satisfy the preset flat eccentricity threshold range at the orbit calculation time, or the flat orbit inclination at the orbit calculation time does not satisfy the preset flat orbit inclination threshold range at the orbit calculation time.

[0016] An adaptive identification system for absolute navigation filtering comprises: a first module, used to set the current beat filter abnormality flag ErrFlag to 1, and set the current beat abnormality count ErrCnt to 0; a second module, used to calculate the modulus of the difference between the position of the observed quantity and the position output of the previous beat filter in the J2000 coordinate system if the current beat filter observation quantity is valid, and if the modulus of the position difference meets the preset threshold range, the current beat filter abnormality flag ErrFlag is set to 0; a third module, used to use the orbit time as the filtering time if the current beat abnormality count ErrCnt is 0, and obtain the current beat filter output state variable and state error covariance matrix according to the filtering time, the previous beat filter output state variable and state error covariance matrix; if the current beat abnormality count ErrCnt is not equal to 0, recalculate The filtering moment is used to obtain the current beat filter output state variable and the state error covariance matrix according to the recalculated filtering moment, the previous beat filter output state variable and the state error covariance matrix; the fourth module is used to obtain the position and speed according to the current beat filter output state variable and the state error covariance matrix, and obtain the position modulus, velocity modulus, position and velocity cross product modulus and the transfer matrix from the J2000 coordinate system to the orbit coordinate system according to the position and velocity; the fifth module is used to obtain the instantaneous root of the filtering moment according to the position velocity, the position modulus, the velocity modulus, the position and velocity cross product modulus and the transfer matrix from the J2000 coordinate system to the orbit coordinate system; the sixth module is used to obtain the instantaneous root of the orbit calculation moment according to the instantaneous root of the filtering moment; the seventh module is used to obtain the flat root of the orbit calculation moment according to the instantaneous root of the orbit calculation moment.

[0017] Compared with the prior art, the present invention has the following beneficial effects:

[0018] (1) During the calculation process of the present invention, the parameters of each step of the calculation process and the result are judged in real time to avoid the introduction of erroneous input observation data, thereby ensuring the real-time and reliability of the attitude and orbit control system, realizing multiple adaptive identifications of the absolute navigation filter within a single shot time, and improving the correctness of the absolute navigation filter results of formation control;

[0019] (2) The process of the present invention is clear, the logic of the identification method is clear, and the algorithm is simple to implement;

[0020] (3) The present invention can avoid navigation filtering result errors caused by the introduction of observation data anomalies;

[0021] (4) The present invention only needs to be judged and completed in a single shot, and is easy to implement on the satellite and operate on the ground. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Various other advantages and benefits will become apparent to those of ordinary skill in the art by reading the detailed description of the preferred embodiments below. The accompanying drawings are only for the purpose of illustrating the preferred embodiments and are not to be considered as limiting the present invention. Moreover, the same reference symbols are used throughout the accompanying drawings to represent the same components. In the accompanying drawings:

[0023] Figure 1 It is a flow chart of an adaptive identification method of absolute navigation filtering provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0024] Exemplary embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. Although exemplary embodiments of the present disclosure are shown in the accompanying drawings, it should be understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided in order to enable a more thorough understanding of the present disclosure and to be able to fully convey the scope of the present disclosure to those skilled in the art. It should be noted that, in the absence of conflict, the embodiments of the present invention and the features described in the embodiments can be combined with each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0025] Figure 1 FIG. 1 is a flow chart of an adaptive identification method for absolute navigation filtering provided by an embodiment of the present invention. Figure 1 As shown, the adaptive identification method of the absolute navigation filter includes:

[0026] Step S1: set the current beat filter abnormality flag ErrFlag to 1, and set the current beat abnormality count ErrCnt to 0;

[0027] Step S2: If the current filter observation is valid, set the current observation flag StateFlag to 0, and the filter beat EKFCnt ≥ 1, then calculate the modulus of the difference between the observation position and the previous filter output position in the J2000 coordinate system, and if the position difference modulus meets the preset threshold range, then set the current filter abnormal flag ErrFlag to 0;

[0028] Step S3: If the current beat abnormal count ErrCnt = 0, use the track time as the filtering time, and output the state variable X according to the filtering time and the previous beat filtering output k-1 and the state error covariance matrix P k-1 Get the current filter output state variable X k and the state error covariance matrix P k ; If the current beat abnormal count ErrCnt is not equal to 0, then set the current beat observation quantity flag StateFlag to 1, recalculate the filtering time = the previous beat filtering time + single beat time, and output the state variable X according to the recalculated filtering time and the previous beat filtering time k-1and the state error covariance matrix P k-1 Get the current filter output state variable X k and the state error covariance matrix P k ;

[0029] Step S4: Filter and output state variable X according to the current beat k and the state error covariance matrix P k Get the position and speed, and get the position R module value, speed V module value, position and speed cross product R×V module value and the transfer matrix from the J2000 coordinate system to the track coordinate system according to the position and speed. If the position R module value does not meet the preset position module value threshold range or the speed V module value does not meet the preset speed module value threshold range or the position and speed cross product R×V module value does not meet the preset cross product threshold range, set ErrFlag to 1 and return to step S1, otherwise ErrFlag is 0;

[0030] Step S5: According to the position speed, the module value R, the module value V, the cross product of the position and the speed R×V module value and A oi Matrix, get the instantaneous root of the filtering moment, if the instantaneous root of the filtering moment does not meet the preset threshold range of the instantaneous root of the filtering moment, set ErrFlag to 1, return to step S1, otherwise ErrFlag is 0;

[0031] Step S6: Obtain the orbit calculation instant root according to the filtering instant root. If the orbit calculation instant root does not meet the preset orbit calculation instant root threshold range, set ErrFlag to 1 and return to step S1; otherwise, ErrFlag is 0.

[0032] Step S7: Obtain the orbit calculation time flat root according to the orbit calculation time instant root. If the orbit calculation time flat root does not meet the preset orbit calculation time flat root threshold range, set ErrFlag to 1 and return to step S1; otherwise, ErrFlag is 0.

[0033] This embodiment can be applied to the satellite attitude and orbit control subsystem. First, the observation data of position and velocity are introduced, and the filtering result of this shot is obtained by combining the filtering data of the previous shot. In the filtering calculation process of this shot, each calculation step will judge the result data introduced by the observation data. If the result of the observation data calculation cannot meet the threshold range, the filtering data of the previous shot is used to recursively obtain the filtering data of this shot to ensure the correctness of the filtering data of this shot.

[0034] Specifically, the absolute navigation filter adaptive identification method of this embodiment includes the following steps:

[0035] The processing methods of absolute navigation include using ground-injected orbit parameters for absolute navigation, using GNSS for absolute navigation, determining the orbit calculation method (instantaneous root), calculating orbit-related parameters (orbital angular velocity, geographic longitude and latitude, solar vector, illumination linear velocity, two-dimensional guidance angle and angular velocity), etc.

[0036] The input of absolute orbit calculation is: the absolute position, velocity, data validity mark, GNSS time of the main and auxiliary satellites, and the absolute navigation filter parameters X0 and P0. The initial value of the filter output state variable X0 is the first GNSS available data when the absolute navigation starts filtering, and the state error covariance matrix P0 is diag([1 1 1 1 1 1]).

[0037] Absolute orbit calculation output: orbit calculation time, absolute position, velocity, number of absolute orbit instantaneous flat roots, and absolute navigation data identification.

[0038] Use the standard EKF filtering algorithm to perform absolute navigation filtering operations on each beat. Determine the EKF algorithm filtering time: the first beat of the algorithm uses the packet time; the subsequent algorithm, if the data is valid, uses the packet time, otherwise it is (previous beat filtering time + single beat time).

[0039] From the previous state Predict state quantity

[0040] The state quantity is The expression of dX=f(X) is as follows:

[0041]

[0042] Where:

[0043] μ—constant, 3.986005×10 14 m 3 / s 2 ;

[0044] a J2 , a J3 , a J4 —Perturbation accelerations of J2, J3 and J4, m / s 2 ;

[0045] r—position modulus of the satellite in the J2000.0 inertial coordinate system, m;

[0046] X—Satellite position velocity in J2000.0 inertial coordinate system, m / s.

[0047] EKF filtering calculation process (6 calculation steps for each beat), before the calculation process starts (set the current beat exception flag ErrFlag to 0, set the current beat exception count ErrCnt to 0, and set the step Step to 0), Step<=6 or ErrCnt<=1 to perform the following cycle.

[0048] Step 1: If the current observation is valid, convert the current observation from WGS84 to J2000. If the filter time EKFCnt ≥ 1, calculate the modulus of the position difference between the observation position in J2000 and the position output of the previous filter. If the modulus of the position difference ≥ the set threshold, set the filter.

[0049] Step 2: If the current beat ErrCnt = 0, use the orbit time in the orbit data packet as the filtering time, and output the state variable X through the previous beat filtering k-1 and the state error covariance matrix P k-1 Output the current filter output state variable X k and the state error covariance matrix P k ; Otherwise, set the current observation invalid and recalculate the filtering time = the previous filtering time + the single-shot time, and output the state variable X through the previous filtering. k-1 and the state error covariance matrix P k-1 Output the current filter output state variable X k and the state error covariance matrix P k .

[0050] Step 3: Use the current beat filter output X from step 2 k and P k Get the position and speed, and calculate the modulus R, modulus V, modulus R×V and A based on the position and speed oi Matrix, if the position modulus does not meet the preset position modulus threshold range or the speed V modulus does not meet the preset speed modulus threshold range or the position and speed cross product modulus does not meet the preset cross product threshold range, then set ErrFlag to 1 and return to step 1, otherwise it is 0.

[0051] Among them, A oi is the transfer matrix from the J2000.0 coordinate system to the orbital coordinate system,

[0052]

[0053] Among them, i is the instantaneous root of the orbital inclination of the local star; Ω is the instantaneous root of the right ascension of the ascending node of the local star; u is the instantaneous root of the latitude argument of the local star.

[0054] Step 4: Use the position, velocity, position modulus R, velocity modulus V, modulus R×V, and A from step 3 oiMatrix, calculate the instantaneous root of the filtering moment, if the instantaneous root of the filtering moment does not meet the preset instantaneous root threshold range of the filtering moment, set the current beat filtering abnormality flag ErrFlag to 1, and return to step 1, otherwise it is 0.

[0055] The instant root at the filtering moment includes the semi-major axis at the filtering moment, the eccentricity at the filtering moment, and the orbital inclination at the filtering moment. The instant root at the filtering moment does not meet the preset threshold range of the instant root at the filtering moment: the semi-major axis at the filtering moment does not meet the preset threshold range of the semi-major axis at the filtering moment, or the eccentricity at the filtering moment does not meet the preset threshold range of the eccentricity at the filtering moment, or the orbital inclination at the filtering moment does not meet the preset threshold range of the orbital inclination at the filtering moment.

[0056] Step 5: Obtain the orbit calculation instant root according to the filtered instant root in step 4. If the orbit calculation instant root does not meet the preset orbit calculation instant root threshold range, set ErrFlag to 1 and return to step 1, otherwise it is 0.

[0057] Step 6: Obtain the orbit calculation time root according to the orbit calculation time instant root in step 5. If the orbit calculation time root does not meet the preset orbit calculation time root threshold range, set ErrFlag to 1 and return to step 1, otherwise set it to 0.

[0058] After each step is executed, the result needs to be judged: if ErrFlag is 1, first set ErrFlag to 0, and then judge if (Step ≥ 1 and ErrCnt = 0 and (EKFCnt ≥ 1, and the current observation is valid)), then set ErrCnt = 1 and set the step count Step to 1; if ErrFlag is 0, then increase the step count by 1.

[0059] Steps 1 to 6 are single-shot multiple EKF filter calculation methods for absolute navigation. After each step is executed, the result needs to be judged: if ErrFlag is 1, first set ErrFlag to 0, and then judge if [Step ≥ 1 and ErrCnt = 0 and (EKFCnt ≥ 1 and the current shot observation is valid)], then set ErrCnt = 1 and set Step to 1; if ErrFlag is 0, then add 1 to the step count.

[0060] Step 1 to Step 6, when ErrCnt≤1: if the current shot observation is valid, set the no observation recursive count EKFDtCnt to 0; otherwise, when EKFDtCnt is accumulated and judged to be greater than the threshold range, set the current shot EKF filter flag to 0 and set the "data valid count" to 0; when the above ErrCnt=2: set the current shot EKF filter flag to 0 and set the "data valid count" to 0.

[0061] In step 1, the EKF filtering algorithm is initialized: the first beat of the EKF filtering is to initialize the EKF filtering algorithm, clear the algorithm-related counters, set the 30-min external filtering warning flag and the 60-min external filtering fault flag to 0, and set the absolute navigation stability flag to 0.

[0062] Step 1 to step 6: before starting the EKF filter calculation under a single shot, set ErrFlag to 0, set the current shot abnormality count ErrCnt to 0, and set the step count Step to 1. If Step<=6 or ErrCnt<=1, perform steps 1 to 6.

[0063] The present embodiment also provides an adaptive identification system for absolute navigation filtering, which includes: a first module, which is used to set the current beat filter abnormality flag ErrFlag to 1 and the current beat abnormality count ErrCnt to 0; a second module, which is used to calculate the modulus of the difference between the position of the observation value and the position output of the previous beat filter in the J2000 coordinate system if the current beat filter observation value is valid, and if the modulus of the position difference meets the preset threshold range, the current beat filter abnormality flag sets ErrFlag to 0; a third module, which is used to use the orbit time as the filtering time if the current beat abnormality count ErrCnt=0, and obtain the current beat filter output state variable and state error covariance matrix according to the filtering time, the previous beat filter output state variable and the state error covariance matrix; if the current beat abnormality count ErrCnt is not equal to 0, recalculate the filtering time, and obtain the current filter output state variable and state error covariance matrix according to the recalculated filtering time, the previous filter output state variable and the state error covariance matrix; the fourth module is used to obtain the position and speed according to the current filter output state variable and the state error covariance matrix, and obtain the position modulus, velocity modulus, position and velocity cross-product modulus and the transfer matrix from the J2000 coordinate system to the orbit coordinate system according to the position and velocity; the fifth module is used to obtain the instantaneous root of the filtering time according to the position and velocity, the position modulus, the velocity modulus, the position and velocity cross-product modulus and the transfer matrix from the J2000 coordinate system to the orbit coordinate system; the sixth module is used to obtain the instantaneous root of the orbit calculation time according to the instantaneous root of the filtering time; the seventh module is used to obtain the flat root of the orbit calculation time according to the instantaneous root of the orbit calculation time.

[0064] During the calculation process of this embodiment, parameter judgment of each step of the calculation process and result is performed in real time to avoid the introduction of erroneous input observation data, thereby ensuring the real-time and reliability of the attitude and orbit control system, realizing multiple adaptive identifications of absolute navigation filtering within a single shot time, and improving the correctness of the absolute navigation filtering results of formation control; the process of this embodiment is clear, the logic of the identification method is clear, and the algorithm is simple to implement; this embodiment can avoid erroneous navigation filtering results caused by the introduction of abnormal observation data; this embodiment only needs to be judged and completed in a single shot, which is easy to implement on the satellite and operate on the ground.

[0065] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. An adaptive identification method for absolute navigation filtering, characterized in that include: Step S1: set the current beat filter abnormality flag ErrFlag to 1, and set the current beat abnormality count ErrCnt to 0; Step S2: If the current filter observation is valid, calculate the modulus of the difference between the observation position and the previous filter output position in the J2000 coordinate system. If the modulus of the position difference meets the preset threshold range, set the current filter abnormality flag ErrFlag to 0; Step S3: If the current beat abnormality count ErrCnt=0, use the track time as the filtering time, and obtain the current beat filter output state variable and state error covariance matrix according to the filtering time, the previous beat filter output state variable and the state error covariance matrix; if the current beat abnormality count ErrCnt is not equal to 0, recalculate the filtering time, and obtain the current beat filter output state variable and state error covariance matrix according to the recalculated filtering time, the previous beat filter output state variable and the state error covariance matrix; Step S4: obtain the position and velocity according to the current beat filter output state variable and the state error covariance matrix, obtain the position modulus, velocity modulus, position and velocity cross multiplication modulus and the transfer matrix from the J2000 coordinate system to the track coordinate system according to the position and velocity, if the position modulus does not meet the preset position modulus threshold range or the velocity V modulus does not meet the preset velocity modulus threshold range or the position and velocity cross multiplication modulus does not meet the preset cross multiplication threshold range, set ErrFlag to 1 and return to step S1, otherwise ErrFlag is 0; Step S5: According to the position speed, the position modulus, the speed modulus, the position and speed cross product modulus, and the transfer matrix from the J2000 coordinate system to the orbital coordinate system, the instantaneous root of the filtering moment is obtained. If the instantaneous root of the filtering moment does not meet the preset threshold range of the instantaneous root of the filtering moment, ErrFlag is set to 1, and the process returns to step S1, otherwise ErrFlag is 0; Step S6: Obtain the orbit calculation instant root according to the filtering instant root. If the orbit calculation instant root does not meet the preset orbit calculation instant root threshold range, set ErrFlag to 1 and return to step S1; otherwise, ErrFlag is 0. Step S7: Obtain the orbit calculation time flat root according to the orbit calculation time instant root. If the orbit calculation time flat root does not meet the preset orbit calculation time flat root threshold range, set ErrFlag to 1 and return to step S1; otherwise, ErrFlag is 0.

2. The method for adaptive identification of absolute navigation filtering according to claim 1, characterized in that: The recalculated filtering time is the previous filtering time plus the single-shot time.

3. The method for adaptive identification of absolute navigation filtering according to claim 1, characterized in that: The transfer matrix from the J2000 coordinate system to the orbital coordinate system is: Among them, i is the instantaneous root of the satellite's orbital inclination; Ω is the instantaneous root of the satellite's ascending node's right ascension; u is the instantaneous root of the satellite's latitude argument.

4. The method for adaptive identification of absolute navigation filtering according to claim 1, characterized in that: The instantaneous root at the filtering moment includes the semi-major axis at the filtering moment, the eccentricity at the filtering moment, and the orbital inclination at the filtering moment.

5. The method for adaptive identification of absolute navigation filtering according to claim 4, characterized in that: The instantaneous root of the filtering moment does not meet the preset threshold range of the instantaneous root of the filtering moment: The semi-major axis at the filtering moment does not meet the preset semi-major axis threshold range at the filtering moment, or the eccentricity at the filtering moment does not meet the preset eccentricity threshold range at the filtering moment, or the orbital inclination at the filtering moment does not meet the preset orbital inclination threshold range at the filtering moment.

6. The method for adaptive identification of absolute navigation filtering according to claim 1, characterized in that: The instantaneous root of orbit calculation time includes the semi-major axis at the orbit calculation time, the eccentricity at the orbit calculation time and the orbit inclination at the orbit calculation time.

7. The method for adaptive identification of absolute navigation filtering according to claim 6, characterized in that: The orbit calculation moment root does not meet the preset orbit calculation moment root threshold range: The semi-major axis at the orbit calculation time does not meet the preset semi-major axis threshold range at the orbit calculation time, or the eccentricity at the orbit calculation time does not meet the preset eccentricity threshold range at the orbit calculation time, or the orbit inclination at the orbit calculation time does not meet the preset orbit inclination threshold range at the orbit calculation time.

8. The method for adaptive identification of absolute navigation filtering according to claim 1, characterized in that: The mean root of orbital calculation time includes the mean semi-major axis of orbital calculation time, the mean eccentricity of orbital calculation time and the mean orbital inclination of orbital calculation time.

9. The method for adaptive identification of absolute navigation filtering according to claim 8, characterized in that: The orbit calculation time flat root does not meet the preset orbit calculation time flat root threshold range: The mean semi-major axis at the orbit calculation time does not meet the preset mean semi-major axis threshold range at the orbit calculation time, or the mean eccentricity at the orbit calculation time does not meet the preset mean eccentricity threshold range at the orbit calculation time, or the mean orbit inclination at the orbit calculation time does not meet the preset mean orbit inclination threshold range at the orbit calculation time.

10. An adaptive identification system for absolute navigation filtering, characterized in that include: The first module is used to set the current beat filter exception flag ErrFlag to 1 and the current beat exception count ErrCnt to 0; The second module is used to calculate the modulus of the difference between the position of the observation value and the position output of the previous filter in the J2000 coordinate system if the current filter observation value is valid. If the modulus of the position difference meets the preset threshold range, the current filter abnormality flag is set to 0. The third module is used to use the orbit time as the filtering time if the current beat abnormality count ErrCnt=0, and obtain the current beat filter output state variable and state error covariance matrix according to the filtering time, the previous beat filter output state variable and the state error covariance matrix; if the current beat abnormality count ErrCnt is not equal to 0, recalculate the filtering time, and obtain the current beat filter output state variable and state error covariance matrix according to the recalculated filtering time, the previous beat filter output state variable and the state error covariance matrix; The fourth module is used to obtain the position and velocity according to the current beat filter output state variable and the state error covariance matrix, and obtain the position modulus, velocity modulus, position and velocity cross product modulus and the transfer matrix from the J2000 coordinate system to the orbital coordinate system according to the position and velocity; The fifth module is used to obtain the instantaneous root of the filtering time according to the position velocity, the position modulus, the velocity modulus, the position and velocity cross product modulus, and the transfer matrix from the J2000 coordinate system to the orbital coordinate system; The sixth module is used to obtain the instantaneous root of orbit calculation time according to the instantaneous root of filtering time; The seventh module is used to obtain the mean root of the orbit calculation time according to the instantaneous root of the orbit calculation time.

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