An adaptive identification method and system for absolute navigation filtering

Abstract

The application discloses an adaptive identification method and system of absolute navigation filtering, wherein the method performs parameter judgment on the calculation result of each step in real time during the calculation process, avoids introducing incorrect input observation data, guarantees the real-time performance and reliability of the attitude and orbit control system, realizes multiple adaptive identifications of the absolute navigation filtering within a single exposure time, and improves the correctness of the absolute navigation filtering result of the formation control.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of satellite formation navigation and control, and particularly relates to an adaptive identification method and system of absolute navigation filtering. BACKGROUND

[0002] Spacecraft formation flight is a new spacecraft space operation mode emerging in the late 1980s along with the development of microsatellites. Compared with a single spacecraft, satellite formation flight has very outstanding advantages, and has been favored by space powers around the world since its inception.

[0003] In order to ensure the formation configuration, the control calculation method of absolute navigation and relative navigation needs to be introduced, and the absolute navigation filtering is an indispensable link. At present, when the absolute navigation filtering of the attitude and orbit control system is calculated, the observation quantity is directly introduced as the input to complete the filtering calculation of the absolute navigation. However, since the observation quantity is introduced for calculation every time, once the observation quantity has a large deviation, the errors of the orbit parameter plane instantaneous root are all present, which affects the calculation of the absolute navigation of the star, and also affects the orbit control of the other star and the formation control. SUMMARY

[0004] The technical problem solved by the application is to overcome the deficiencies of the prior art, and provide an adaptive identification method and system of absolute navigation filtering, which realizes the multiple adaptive cyclic identification method of the absolute navigation filtering under a single shot, and improves the correctness and reliability of the calculation result of the absolute navigation EKF filtering.

[0005] The application is achieved by the following technical scheme: an adaptive identification method of absolute navigation filtering, comprising: step S1: setting a current filter abnormality flag ErrFlag as 1 and a current filter abnormality count ErrCnt as 0; step S2: if a current filter observation is valid, calculating a modulus value of a position difference between an observation position in a J2000 coordinate system and a last filter output position, and setting the current filter abnormality flag ErrFlag as 0 if the modulus value of the position difference meets a preset threshold range; step S3: if the current filter abnormality count ErrCnt is 0, using an orbit time as a filter time, obtaining a current filter output state variable and a state error covariance matrix according to the filter time, a last filter output state variable and a state error covariance matrix; if the current filter abnormality count ErrCnt is not 0, recalculating the filter time, and obtaining the current filter output state variable and the state error covariance matrix according to the recalculated filter time, the last filter output state variable and the state error covariance matrix; step S4: obtaining a position and a velocity according to the current filter output state variable and the state error covariance matrix, obtaining a position modulus value, a velocity modulus value, a position and velocity cross product modulus value and a transition matrix from the J2000 coordinate system to an orbit coordinate system according to the position and the velocity, setting the ErrFlag as 1 and returning to step S1 if the position modulus value does not meet a preset position modulus value threshold range, the velocity modulus value does not meet a preset velocity modulus value threshold range or the position and velocity cross product modulus value does not meet a preset cross product threshold range, or setting the ErrFlag as 0; step S5: obtaining a filter time instant root according to the position and velocity, the position modulus value, the velocity modulus value, the position and velocity cross product modulus value and the transition matrix from the J2000 coordinate system to the orbit coordinate system, setting the ErrFlag as 1 and returning to step S1 if the filter time instant root does not meet a preset filter time instant root threshold range, or setting the ErrFlag as 0; step S6: obtaining an orbit calculation time instant root according to the filter time instant root, setting the ErrFlag as 1 and returning to step S1 if the orbit calculation time instant root does not meet a preset orbit calculation time instant root threshold range, or setting the ErrFlag as 0; step S7: obtaining an orbit calculation time instant plane root according to the orbit calculation time instant root, setting the ErrFlag as 1 and returning to step S1 if the orbit calculation time instant plane root does not meet a preset orbit calculation time instant plane root threshold range, or setting the ErrFlag as 0.

[0006] In the adaptive identification method of absolute navigation filtering, the recalculated filter time is the last filter time plus a single filter time.

[0007] In the adaptive identification method of absolute navigation filtering, the transition matrix from the J2000 coordinate system to the orbit coordinate system is:

[0008]

[0009] Wherein, i is the instantaneous root of the orbit inclination; Ω is the instantaneous root of the ascending node right ascension; and u is the instantaneous root of the latitude amplitude.

[0010] In the adaptive identification method of the absolute navigation filter, the filter time instantaneous root comprises a filter time semi-major axis, a filter time eccentricity and a filter time orbit inclination.

[0011] In the adaptive identification method of the absolute navigation filter, the filter time instantaneous root does not satisfy a preset filter time instantaneous root threshold range, i.e., the filter time semi-major axis does not satisfy a preset filter time semi-major axis threshold range, or the filter time eccentricity does not satisfy a preset filter time eccentricity threshold range, or the filter time orbit inclination does not satisfy a preset filter time orbit inclination threshold range.

[0012] In the adaptive identification method of the absolute navigation filter, the orbit calculation time instantaneous root comprises an orbit calculation time semi-major axis, an orbit calculation time eccentricity and an orbit calculation time orbit inclination.

[0013] In the adaptive identification method of the absolute navigation filter, the orbit calculation time instantaneous root does not satisfy a preset orbit calculation time instantaneous root threshold range, i.e., the orbit calculation time semi-major axis does not satisfy a preset orbit calculation time semi-major axis threshold range, or the orbit calculation time eccentricity does not satisfy a preset orbit calculation time eccentricity threshold range, or the orbit calculation time orbit inclination does not satisfy a preset orbit calculation time orbit inclination threshold range.

[0014] In the adaptive identification method of the absolute navigation filter, the orbit calculation time plane root comprises an orbit calculation time plane semi-major axis, an orbit calculation time plane eccentricity and an orbit calculation time plane orbit inclination.

[0015] In the adaptive identification method of the absolute navigation filter, the orbit calculation time plane root does not satisfy a preset orbit calculation time plane root threshold range, i.e., the orbit calculation time plane semi-major axis does not satisfy a preset orbit calculation time plane semi-major axis threshold range, or the orbit calculation time plane eccentricity does not satisfy a preset orbit calculation time plane eccentricity threshold range, or the orbit calculation time plane orbit inclination does not satisfy a preset orbit calculation time plane orbit inclination threshold range.

[0016] An adaptive identification system of absolute navigation filtering comprises: a first module for setting a current filtering exception flag ErrFlag to 1 and a current filtering exception count ErrCnt to 0; a second module for calculating a modulus value of a position difference between an observation position in a J2000 coordinate system and a last filtering output position if a current filtering observation is valid, and setting the current filtering exception flag ErrFlag to 0 if the modulus value of the position difference meets a preset threshold range; a third module for using an orbit time as a filtering time, obtaining a current filtering output state variable and a state error covariance matrix according to the filtering time, a last filtering output state variable and a state error covariance matrix if the current filtering exception count ErrCnt is 0; recalculating the filtering time and obtaining the current filtering output state variable and the state error covariance matrix according to the recalculated filtering time, the last filtering output state variable and the state error covariance matrix if the current filtering exception count ErrCnt is not equal to 0; a fourth module for obtaining a position and a velocity according to the current filtering output state variable and the state error covariance matrix, obtaining a position modulus value, a velocity modulus value, a position and velocity cross product modulus value and a transition matrix from the J2000 coordinate system to an orbit coordinate system according to the position and the velocity; a fifth module for obtaining a filtering time instant root according to the position, the velocity, the position modulus value, the velocity modulus value, the position and velocity cross product modulus value and the transition matrix from the J2000 coordinate system to the orbit coordinate system; a sixth module for obtaining an orbit calculation time instant root according to the filtering time instant root; and a seventh module for obtaining an orbit calculation time plane root according to the orbit calculation time instant root.

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

[0018] (1) The present application judges the parameters of each step of the calculation process and the results in real time during the calculation process, avoids introducing incorrect input observation data, ensures the real-time performance and reliability of the attitude and orbit control system, realizes multiple adaptive identification of the absolute navigation filtering within a single shot time, and improves the correctness of the absolute navigation filtering result of the formation control;

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

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

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

[0022] Various other advantages and benefits will become apparent to those of ordinary skill in the art upon reading the following detailed description of the preferred embodiments with reference made to the accompanying drawings. The drawings are for purposes of illustration only and are not intended to be limiting in

[0023] Figure 1 is a flow chart of the adaptive identification method of absolute navigation filtering provided by the embodiments of the present application. DETAILED DESCRIPTION

[0024] Exemplary embodiments of the present disclosure will be described herein below with reference to the accompanying drawings. Although exemplary embodiments of the present disclosure are shown in the drawings, it is understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the present disclosure to those skilled in the art. It should be noted that the embodiments in the present application and the features in the embodiments can be combined with each other without conflict. The present application will be described in detail below with reference to the accompanying drawings and in conjunction with the embodiments.

[0025] Figure 1 is a flow chart of the adaptive identification method of absolute navigation filtering provided by the embodiments of the present application. As shown in Figure 1 , the adaptive identification method of absolute navigation filtering comprises:

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

[0027] Step S2: if the current filter observation is valid, setting the current observation flag StateFlag to 0, and the filter shot count EKFCnt≥1, calculating the modulus value of the difference between the observation position in the J2000 coordinate system and the last filter output position, and if the modulus value of the position difference meets the preset threshold range, setting the current filter abnormality flag ErrFlag to 0;

[0028] Step S3: if the current filter abnormality count ErrCnt=0, using the orbit time as the filter time, and obtaining the current filter output state variable X k-1 and the state error covariance matrix P k-1 according to the filter time, the last filter output state variable X k and the state error covariance matrix P k ; if the current filter abnormality count ErrCnt is not equal to 0, setting the current observation flag StateFlag to 1, recalculating the filter time = the last filter time + the single shot time, and obtaining the current filter output state variable X k-1and state error covariance matrix P k-1 get the current filter output state variable X k and state error covariance matrix P k

[0029] Step S4: according to the current filter output state variable X k and state error covariance matrix P k get the position and velocity, and according to the position and velocity, get the position R modulus value, velocity V modulus value, position and velocity cross product R x V modulus value, and transfer matrix of J2000 coordinate system to orbital coordinate system, if the position R modulus value does not satisfy the preset position modulus value threshold range or the velocity V modulus value does not satisfy the preset velocity modulus value threshold range or the position and velocity cross product R x V modulus value does not satisfy the preset cross product threshold range, set ErrFlag as 1, return to step S1, otherwise ErrFlag is 0;

[0030] Step S5: according to the position and velocity, modulus value R, modulus value V, position and velocity cross product R x V modulus value, and A oi matrix, get the filter time instant root, if the filter time instant root does not satisfy the preset filter time instant root threshold range, set ErrFlag as 1, return to step S1, otherwise ErrFlag is 0;

[0031] Step S6: according to the filter time instant root, get the orbit calculation time instant root, if the orbit calculation time instant root does not satisfy the preset orbit calculation time instant root threshold range, set ErrFlag as 1, return to step S1, otherwise ErrFlag is 0;

[0032] Step S7: according to the orbit calculation time instant root, get the orbit calculation time instant flat root, if the orbit calculation time instant flat root does not satisfy the preset orbit calculation time instant flat root threshold range, set ErrFlag as 1, return to step S1, otherwise ErrFlag is 0.

[0033] The embodiment can be applied to a satellite attitude and orbit control subsystem, first introduce the position and velocity observation data, and combine the filter data of the last filter to get the filter result of this filter. In the filter calculation process of this filter, the result data introduced by the observation at each calculation step is judged, if the result calculated by the observation data cannot satisfy the threshold range, the filter data of this filter is obtained by using the filter data of the last filter, and the correctness of the filter data of this filter is ensured

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

[0035] ​The processing mode of absolute navigation includes using ground injection orbit parameters for absolute navigation, using GNSS for absolute navigation, determining orbit calculation mode (instant root), orbit related parameter calculation (orbit angular velocity, geographic longitude and latitude, sun vector, light line velocity, two-dimensional guide angle and angular velocity), etc.

[0036] Absolute orbit calculation input: primary and secondary star GNSS absolute position, velocity, data valid identification, GNSS time, absolute navigation filter parameter X0, P0. Among them, the filter output state variable initial value X0 is the first GNSS available data at the beginning of absolute navigation filtering, and the state error covariance matrix is P0.

[0037] Absolute orbit calculation output: orbit calculation time, absolute position, velocity, absolute orbit instant root number, absolute navigation data identification.

[0038] The standard EKF filter algorithm is used for absolute navigation filtering operation of each shot. The EKF algorithm filtering time is determined: the first shot of the algorithm uses the packet time; for subsequent algorithms, if the data is valid, the packet time is used, otherwise, the (previous shot filtering time + single shot time) is used.

[0039] The state of the previous shot is State quantity prediction is performed

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

[0041]

[0042] In the formula:

[0043] μ - constant, 3.986005 * 10 14 m 3 / s 2 ;

[0044] a J2 , a J3 , a J4 - J2, J3 and J4 term perturbation acceleration, m / s 2 ;

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

[0046] X - satellite position and velocity in J2000.0 inertial coordinate system, m / s.

[0047] EKF filtering calculation process (6 calculation steps per beat), before the calculation process (set the current beat exception identification ErrFlag to 0, set the current beat exception count ErrCnt to 0, and set the step Step to 0), Step<=6 or ErrCnt<=1, the following loop is performed.

[0048] Step 1: If the current observation is valid, convert the current observation from WGS84 to J2000, and if the filter beat EKFCnt is greater than or equal to 1, calculate the modulus of the difference between the observation position in J2000 and the last beat filter output position, and if the modulus of the position difference is greater than or equal to 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 filter time, and use the last beat filter output state variable X k-1 and state error covariance matrix P k-1 output the current beat filter output state variable X k and state error covariance matrix P k ; otherwise, set the current observation as invalid and recalculate the filter time = last beat filter time + single beat time, and use the last beat filter output state variable X k-1 and state error covariance matrix P k-1 output the current beat filter output state variable X k and state error covariance matrix P k .

[0050] Step 3: Use the current beat filter output X k and P k to get the position and velocity, and calculate the modulus R, modulus V, modulus R x V, and A oi matrix, if the position modulus does not satisfy the preset position modulus threshold range or the velocity V modulus does not satisfy the preset velocity modulus threshold range or the position and velocity cross product modulus does not satisfy the preset cross product threshold range, set ErrFlag to 1, return to step 1, otherwise set ErrFlag to 0.

[0051] where A oi is the transformation matrix from J2000.0 coordinate system to orbit coordinate system,

[0052]

[0053] where i is the instantaneous root of the inclination of the orbit, Ω is the instantaneous root of the right ascension of the ascending node, and u is the instantaneous root of the amplitude angle of the latitude.

[0054] Step 4: Use the position, velocity, position modulus R, velocity modulus V, modulus R x V, and A oiThe matrix is used to calculate the filter time instant root. If the filter time instant root does not satisfy a preset filter time instant root threshold range, a current shot filter abnormality flag ErrFlag is set to 1, and step 1 is returned. Otherwise, the ErrFlag is 0.

[0055] The filter time instant root includes a filter time instant semi-major axis, a filter time instant eccentricity and a filter time instant orbital inclination. The filter time instant root not satisfying the preset filter time instant root threshold range means that the filter time instant semi-major axis does not satisfy a preset filter time instant semi-major axis threshold range, or the filter time instant eccentricity does not satisfy a preset filter time instant eccentricity threshold range, or the filter time instant orbital inclination does not satisfy a preset filter time instant orbital inclination threshold range.

[0056] Step 5: The orbit calculation time instant root is obtained according to the filter time instant root of step 4. If the orbit calculation time instant root does not satisfy a preset orbit calculation time instant root threshold range, the ErrFlag is set to 1, and step 1 is returned. Otherwise, the ErrFlag is 0.

[0057] Step 6: The orbit calculation time instant plane root is obtained according to the orbit calculation time instant root of step 5. If the orbit calculation time instant plane root does not satisfy a preset orbit calculation time instant plane root threshold range, the ErrFlag is set to 1, and step 1 is returned. Otherwise, the ErrFlag is 0.

[0058] At the end of each step, a result judgment needs to be performed. If the ErrFlag is 1, the ErrFlag is first set to 0, and then if (Step≥1 and ErrCnt=0 and (EKFCnt≥1 and current shot observation is valid)), the ErrCnt is set to 1 and the step count Step is set to 1. If the ErrFlag is 0, the step count is increased by 1.

[0059] Steps 1-6 are a single-shot multiple-time EKF filter calculation method for absolute navigation. At the end of each step, a result judgment needs to be performed. If the ErrFlag is 1, the ErrFlag is first set to 0, and then if [Step≥1 and ErrCnt=0 and (EKFCnt≥1 and current shot observation is valid)], the ErrCnt is set to 1 and the Step is set to 1. If the ErrFlag is 0, the step count is increased by 1.

[0060] Steps 1-6, when ErrCnt≤1: If the current shot observation is valid, the no-observation recursive count EKFDtCnt is set to 0. Otherwise, the EKFDtCnt is accumulated and is greater than a threshold range, the current shot EKF filter flag is set to 0, and the “data valid count” is set to 0. When the above ErrCnt=2: the current shot EKF filter flag is set to 0, and the “data valid count” is set to 0.

[0061] In step 1, the EKF filtering algorithm is initialized: the EKF filtering first beat, the EKF filtering algorithm is initialized, the algorithm related counter is cleared, the 30 min outside filtering early warning flag and the 60 min outside filtering failure flag are set to 0, and the absolute navigation stability flag is set to 0.

[0062] In steps 1-6, before starting the EKF filtering calculation under single beat, the ErrFlag is set to 0, the current beat abnormality count ErrCnt is set to 0, and the step count Step is set to 1, if Step<=6 or ErrCnt<=1, steps 1-6 are performed.

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

[0064] In the calculation process of the embodiment, the parameter judgment of each calculation process and result is performed in real time, so that the introduction of incorrect input observation data is avoided, the real-time performance and reliability of the attitude and orbit control system are ensured, the multiple adaptive identifications of the absolute navigation filtering in a single beat time are realized, and the correctness of the formation control absolute navigation filtering result is improved; the process of the embodiment is clear, the identification method is logical, and the algorithm is simple to realize; the navigation filtering result error caused by the introduction of abnormal observation data can be avoided by the embodiment; the embodiment only needs to be judged under a single beat, and is easy to implement on a satellite and operate on the ground.

[0065] Although the present application has been disclosed with reference to the preferred embodiments, it is not intended to limit the present application, and any person skilled in the art can make possible changes and modifications to the technical solutions of the present application using the disclosed methods and technical contents without departing from the spirit and scope of the present application. Therefore, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present application without departing from the technical solutions of the present application shall fall within the protection scope of the present application.

Claims

1. A method of adaptive identification of an absolute navigation filter, characterized in that The method comprises the following steps: Step S1: setting a current filter abnormality flag ErrFlag as 1 and a current filter abnormality count ErrCnt as 0; Step S2: if a current filter observation is valid, calculating a modulus value of a difference between an observation position in a J2000 coordinate system and a last filter output position, and setting the current filter abnormality flag ErrFlag as 0 if the modulus value of the position difference meets a preset threshold range; Step S3: if the current filter abnormality count ErrCnt is 0, using an orbit time as a filter time, and obtaining a current filter output state variable and a state error covariance matrix according to the filter time, a last filter output state variable and a state error covariance matrix; if the current filter abnormality count ErrCnt is not 0, recalculating the filter time, and obtaining the current filter output state variable and the state error covariance matrix according to the recalculated filter time, the last filter output state variable and the state error covariance matrix; Step S4: obtaining a position and a velocity according to the current filter output state variable and the state error covariance matrix, obtaining a position modulus value, a velocity modulus value, a position and velocity cross product modulus value and a transfer matrix from the J2000 coordinate system to an orbit coordinate system according to the position and the velocity, setting the ErrFlag as 1 and returning to step S1 if the position modulus value does not meet a preset position modulus value threshold range, the velocity modulus value does not meet a preset velocity modulus value threshold range or the position and velocity cross product modulus value does not meet a preset cross product threshold range, or setting the ErrFlag as 0; Step S5: obtaining a filter time instant root according to the position and velocity, the position modulus value, the velocity modulus value, the position and velocity cross product modulus value and the transfer matrix from the J2000 coordinate system to the orbit coordinate system, setting the ErrFlag as 1 and returning to step S1 if the filter time instant root does not meet a preset filter time instant root threshold range, or setting the ErrFlag as 0; Step S6: obtaining an orbit calculation time instant root according to the filter time instant root, setting the ErrFlag as 1 and returning to step S1 if the orbit calculation time instant root does not meet a preset orbit calculation time instant root threshold range, or setting the ErrFlag as 0; Step S7: obtaining an orbit calculation time plane root according to the orbit calculation time instant root, setting the ErrFlag as 1 and returning to step S1 if the orbit calculation time plane root does not meet a preset orbit calculation time plane root threshold range, or setting the ErrFlag as 0.

2. The method of adaptive identification of the absolute navigation filter according to claim 1, characterized in that: The recalculated filter time is the last filter time plus a single shot time.

3. The method of claim 1, wherein: The transfer matrix from the J2000 coordinate system to the orbit coordinate system is: wherein i is an instant root of an orbit inclination, Ω is an instant root of an ascending node right ascension of the sun, and u is an instant root of an amplitude of the sun's latitude.

4. The method of claim 1, wherein: The filter time instant root comprises a filter time semi-major axis, a filter time eccentricity and a filter time orbit inclination.

5. The method of adaptive identification of the absolute navigation filter according to claim 4, characterized in that: The filter time instant root does not meet the preset filter time instant root threshold range if the filter time semi-major axis does not meet a preset filter time semi-major axis threshold range, the filter time eccentricity does not meet a preset filter time eccentricity threshold range or the filter time orbit inclination does not meet a preset filter time orbit inclination threshold range. ​ 6. The method of claim 1, wherein: The orbit calculation time instant moment root comprises an orbit calculation time instant semi-major axis, an orbit calculation time instant eccentricity and an orbit calculation time instant orbit inclination.

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

8. The method of claim 1, wherein: The orbit calculation time instant plane root comprises an orbit calculation time instant plane semi-major axis, an orbit calculation time instant plane eccentricity and an orbit calculation time instant plane orbit inclination.

9. The method of adaptive identification of the absolute navigation filter according to claim 8, characterized in that: The orbit calculation time instant plane root does not satisfy a preset orbit calculation time instant plane root threshold range. The orbit calculation time instant plane semi-major axis does not satisfy a preset orbit calculation time instant plane semi-major axis threshold range, or the orbit calculation time instant plane eccentricity does not satisfy a preset orbit calculation time instant plane eccentricity threshold range, or the orbit calculation time instant plane orbit inclination does not satisfy a preset orbit calculation time instant plane orbit inclination threshold range.

10. An adaptive identification system for absolute navigation filtering, characterized by The method comprises the following steps: A first module is configured to set a current beat filter abnormality identifier ErrFlag to 1 and a current beat abnormality counter ErrCnt to 0. A second module is configured to calculate a modulus value of a difference between an observation position in a J2000 coordinate system and a last beat filter output position if a current beat filter observation is valid, and set the current beat filter abnormality identifier ErrFlag to 0 if the modulus value of the position difference satisfies a preset threshold range. A third module is configured to use an orbit time instant as a filter time instant, obtain a current beat filter output state variable and a state error covariance matrix according to the filter time instant, a last beat filter output state variable and a state error covariance matrix if the current beat abnormality counter ErrCnt is 0, and obtain a current beat filter output state variable and a state error covariance matrix according to a recalculated filter time instant, a last beat filter output state variable and a state error covariance matrix if the current beat abnormality counter ErrCnt is not equal to 0. A fourth module is configured to obtain a position and a velocity according to the current beat filter output state variable and the state error covariance matrix, obtain a position modulus value, a velocity modulus value, a position and velocity cross product modulus value and a transfer matrix from a J2000 coordinate system to an orbit coordinate system according to the position and the velocity. A fifth module is configured to obtain a filter time instant moment root according to the position and velocity, the position modulus value, the velocity modulus value, the position and velocity cross product modulus value and the transfer matrix from the J2000 coordinate system to the orbit coordinate system. A sixth module is configured to obtain an orbit calculation time instant moment root according to the filter time instant moment root. A seventh module is configured to obtain an orbit calculation time instant plane root according to the orbit calculation time instant moment root.

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