CN112836339B - Navigation satellite orbit extrapolation software design method - Google Patents
Navigation satellite orbit extrapolation software design method Download PDFInfo
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Abstract
The invention provides a navigation satellite orbit extrapolation software design method, which comprises the following steps: the orbit data preprocessing module receives the number of the observing and controlling injection orbit, the navigation satellite ephemeris and astronomical navigation attitude data as orbit data to be calculated; the track data backup and recovery module performs data backup on the number of the track to be observed and controlled; the track source preprocessing self-checking module performs self-checking on track data to be calculated according to the check code and performs mutual checking on the track data to be calculated; the orbit self-diagnosis module selects the number of the observing and controlling injection orbit according to an orbit self-diagnosis algorithm, and the navigation satellite ephemeris or astronomical navigation attitude data is used as orbit extrapolation calculation data; the orbit extrapolation calculation module obtains real-time orbit data of the satellite by using the orbit extrapolation calculation data, and the real-time orbit data is used for closed-loop calculation of a satellite attitude control system; and the track source main selection switching module performs autonomous selection and switching of the track data to be calculated according to the track priority, the track switch state and the track data self-diagnosis result.
Description
Technical Field
The invention relates to the technical field of navigation satellites, in particular to a method for designing extrapolation software of navigation satellite orbits.
Background
The main function of the navigation satellite is to provide all-weather and high-precision navigation and time service for global users, and has high requirements on the continuity of the service. The high-reliability satellite orbit of the navigation satellite is maintained, so that not only can the gesture and pointing precision of the navigation satellite be ensured, but also the ground navigation receiving terminal is required to be ensured to obtain high-precision positioning. The stability of the existing navigation satellite can not fully meet the requirements, so that research on a method for improving the stable operation capability of the navigation satellite is urgently needed.
Disclosure of Invention
The invention aims to provide a design method of navigation satellite orbit extrapolation software, which aims to solve the problem that the stability of the existing navigation satellite is difficult to improve.
In order to solve the technical problems, the invention provides a navigation satellite orbit extrapolation software design method, which comprises the following steps:
the orbit data preprocessing module is configured to receive the number of the observing and controlling injection orbit, the navigation satellite ephemeris and astronomical navigation attitude data as orbit data to be calculated;
the track data backup and recovery module is configured to backup data for the number of the track to be observed and controlled;
the track source preprocessing self-checking module is configured to perform self-checking on track data to be calculated according to the check code and perform mutual checking on the track data to be calculated;
the orbit origin diagnosis module is configured to select the number of observing and controlling the upper-injection orbit according to an orbit origin diagnosis algorithm, and the navigation satellite ephemeris or astronomical navigation attitude data is used as orbit extrapolation calculation data;
the orbit extrapolation calculation module is configured to obtain real-time orbit data of the satellite by using the orbit extrapolation calculation data, and the real-time orbit data is used for closed-loop calculation of the satellite attitude control system;
the track source main selection switching module is configured to perform autonomous selection and switching of track data to be calculated according to the track source priority, the track switch state and the track data self-diagnosis result.
Optionally, in the method for designing the software for extrapolating the orbit of the navigation satellite, the number of the orbit to be observed and controlled is a first orbit source, including the number of the flat orbit to be observed and controlled on the ground;
the navigation satellite ephemeris is a second orbit source comprising obtaining a position velocity under WGS-84 from the loaded ephemeris;
the astronomical navigation attitude data is a third orbit source and comprises orbit data for performing astronomical orbit calculation according to satellite attitude sensitive information.
Optionally, in the method for designing the software for extrapolating the orbit of the navigation satellite,
when the number of the orbit-determining ground injection is flat, carrying out orbit calculation according to the load ephemeris or carrying out orbit extrapolation calculation through satellite attitude sensitive information;
and the first track source, the second track source and the third track source are subjected to priority management, and an optimal track is selected by a track source data mutual checking mechanism so as to realize uninterrupted switching among different track sources and ensure that track data input to gesture control is continuously and effectively.
Optionally, in the method for designing the software for extrapolating the orbit of the navigation satellite,
the working modes of the navigation satellite orbit extrapolation software comprise a ground control mode and an on-board autonomous mode;
under a ground control mode, selecting a first track source through measuring and controlling an uplink instruction;
in the autonomous on-board mode, the navigation satellite orbit extrapolation software performs autonomous selection and switching according to the orbit source priority, the orbit switching state and the orbit data self-diagnosis result.
Optionally, in the method for designing the software for extrapolating the orbit of the navigation satellite, the orbit data to be calculated is subjected to local data backup and remote data backup;
after receiving the orbit data of the measurement and control betting, the navigation satellite orbit extrapolation software respectively performs remote data backup in the EEPROM of the spaceborne computer and the bus data terminal under the spaceborne computer;
diagnosing and selecting by adopting a three-in-two strategy, and storing the number of the measurement and control injection tracks into main data, backup data and default data;
after the track number is injected into the ground, the time stamp and the data validity check are carried out on the injected data, the check is successful, and then the corresponding backup position data are updated according to the storage logic.
Optionally, in the method for designing the software for extrapolating the orbit of the navigation satellite,
the ephemeris data are generated for a single loading machine of a navigation satellite, the ephemeris and the loading time t are read from a loading task processor through a 1553 bus every cycle by a satellite-borne computer, and the J2000 position and the J2000 speed are calculated by using the Beidou loading;
calculating a reference time product second t1:
t 1 ＝WNgro*604800+t oe
calculating time deviation:
t k ＝t-t oe
wherein t is the current Beidou, t oe Reference time for ephemeris; considering the beginning or end of the week transform, if tk is greater than 302400, tk is subtracted by 604800; if tk is less than-302400, then tk is added to 604800;
calculating a straight-up point angle:
A 0 ＝27906100+ΔA
n＝n 0 +Δn, whereμ＝3.986004418e 14
M k ＝M 0 +nt k ，
Calculating a close point angle:
M k ＝E k -esinE k
the iteration method is used for calculation, and the maximum number of iterations is 60, and each iteration is calculated as follows:
E k+1 ＝E k -(E k -e*sinE k -M k )/(1-e*cosE k )
if E k+1 -E k ≤1e -6 Ending the iteration;
calculating phi k ：
Φ k ＝V k +ω, wherein,
calculating δu k 、δr k 、δi k ：
Calculation u k 、r k 、i k ：
Calculated geodetic position XYZ:
wherein,
calculated geodesic velocities Vx, vy, vz:
wherein omega e ＝7.292115E-5rad/s
Wherein,
ground-fixed system position and speed-J2000 position and speed;
r J ＝(G·N·P) T ·r W wherein r is J For the J2000 position r W Is a ground fixation system position;
wherein,speed J2000>Is the ground fastening speed->
ω e ＝[0,0,ω e ]Is the earth rotation rate vector.
Optionally, in the method for designing the software for extrapolating the orbit of the navigation satellite,
adopting a J2 perturbation model to extrapolate the root number according to the dynamics principle of the satellite; the extrapolation process comprises Kepler root conversion and utilization of t 0 Number of time-of-day squaresCalculating t i Time of day of the number of roots>By t i Time of day of the number of roots>Calculating t i Instantaneous root number sigma at moment, utilization t i Calculating t by instantaneous root sigma of moment i Position and velocity data of time satellite in J2000 coordinate system
By t 0 Number of time-of-day squaresCalculate the number of roots at time t +.>
Wherein: />
Using the number of squares at time tCalculating the instantaneous root sigma (a, i, omega, ζ, eta, lambda) at the moment t:
wherein DeltaSigma s The expression of (2) is as follows:
if the track extrapolation semi-major axis check enable flag is enabled, checking the transient root number semi-major axis as follows:
|a×6378.137-28125|≤△a
wherein: the default value of delta a is 1000km, and the delta a is changed through uploading data;
if the conditions are met, checking to be correct, setting an extrapolation effective mark of the measurement and control track as effective, and converting the instantaneous root number into position speed data; checking errors, setting an extrapolation valid mark of the measurement and control track as invalid, and keeping the last value of the extrapolated track data; if the error is checked for 5 times continuously, setting the reference point use mark as an old reference point, and if only one group exists, switching to a default extrapolation point;
if the track extrapolation semi-long axis verification enabling mark is forbidden, an extrapolation track valid mark is set to be valid;
calculating coordinates of satellite at t moment in J2000 coordinate system by using instantaneous root number sigma at t momentAnd speed->The units are people respectivelyA guard length unit and a guard length unit/guard time unit;
wherein,
p＝a
the calculation method of u is as follows:
ω＝a tan 2(-η,ξ)
M＝λ-ω
f＝M+2e sin M
u＝ω+f
wherein atan2 represents a two-dimensional arctangent function;
coordinates are setAnd speed->Units are converted to meters and meters per second.
Optionally, in the method for designing the software for extrapolating the orbit of the navigation satellite, the satellite sensitive quaternion and the ground sensitive geocentric vector are utilized, and the simplified orbit model is filtered through an unscented Kalman filter to obtain the current inertial system position and speed vector of the satellite;
acquiring a current measured value, and determining a quaternion Q of a attitude-determining inertial system of the star sensor bi Conversion to a gesture matrix R bi By R bi And a geodetic attitude-determining geodetic vector E b Calculating the measured value y of the geocentric vector of the inertial system at the current moment k ：
Sigma points were calculated:
p pair P x,k-1 The matrix is subjected to Cholesky decomposition to obtain P x,k-1 Square root G of the array:
wherein pij is P x,k-1 Elements i, j E [1,6 ]]；
Establishing a time propagation equation, and calculating a predicted value of the Sigma point according to a system state equation;
taking a two-body model represented by six numbers as a state equation, and calculating a predicted value of the state quantity:
calculating a predicted value of the covariance matrix:
calculating a predicted value of the observed quantity:
the observed quantity is a satellite inertial system geocentric vector, and is obtained by using the satellite position inverse number in the state quantity and normalizing;
calculating a predicted value of the output quantity:
measuring an update equation, and calculating a gain array:
updating the covariance matrix:
updating state quantity:
optionally, in the method for designing the software for extrapolating the orbit of the navigation satellite, a recursive algorithm using a two-body model represented by six numbers as a state equation is as follows:
converting position and speed into six track numbers
Wherein->
Calculating the six numbers of the next period:
wherein the method comprises the steps of
Ek is calculated by using an iterative method, and is iterated for 60 times at most, and each iteration is calculated as follows: ek (Ek) k+1 ＝M+e*sin(Ek k ) The method comprises the steps of carrying out a first treatment on the surface of the If E k+1 -E k ≤1e -6 Ending the iteration;
conversion of six track numbers into position and speed
Wherein->
Wherein->
In the design method of the navigation satellite orbit extrapolation software provided by the invention, the orbit data preprocessing module is used for receiving the measurement and control of the number of the injection orbits, the ephemeris of the navigation satellites and the astronomical navigation attitude data as orbit data to be calculated, so that the orbit data is provided by multiple orbit sources and multiple channels, and the stability and the instantaneity (continuity) of extrapolation calculation are ensured; the track data backup and recovery module is used for carrying out data backup on the number of the track to be observed and controlled and the number of the track to be injected, so that the reliability of data storage is realized; the track source preprocessing self-checking module performs self-checking on track data to be calculated according to the check codes and performs mutual checking on the track data to be calculated, so that the optimal track source is selected, the data source which cannot pass the checking due to data errors is eliminated, and the accuracy of track extrapolation calculation is improved; the orbit self-diagnosis module selects the number of the injection orbits to measure and control according to the orbit self-diagnosis algorithm, the navigation satellite ephemeris or astronomical navigation attitude data is used as orbit extrapolation calculation data, the orbit extrapolation calculation module utilizes the orbit extrapolation calculation data to obtain real-time orbit data of the satellite, the real-time orbit data is used for closed-loop calculation of a satellite attitude control system, the optimal orbit source is selected, and the orbit extrapolation calculation is more stable, accurate and real-time; in addition, the track source main selection switching module is used for automatically selecting and switching the track data to be calculated according to the track source priority, the track switch state and the track data self-diagnosis result, so that the flexibility and the high efficiency of the whole scheme are embodied.
In summary, the invention adopts a software design method of multi-track source data fusion, autonomous diagnosis and autonomous selection of the optimal track, and aims at three track data: and observing and controlling the track data of the uploading, the ephemeris track data, utilizing satellite star sensitive quaternion and earth sensitive geocentric vector to filter the track data of the simplified track model through an Unscented Kalman Filter (UKF), carrying out data preprocessing and backup in a local and remote storage mode, simultaneously carrying out autonomous diagnosis on the track data, and finally selecting the optimal track data. The invention solves the problems of continuity, stability and reliability of the satellite-borne orbit extrapolation data of the navigation satellite, and improves the satellite attitude pointing precision.
Drawings
FIG. 1 is a schematic diagram of a data selection process for software for extrapolating orbit of a navigation satellite according to an embodiment of the present invention;
FIG. 2 is a schematic diagram of the overall flow of software for extrapolating the orbit of a navigation satellite according to an embodiment of the present invention;
FIG. 3 is a schematic diagram of a self-diagnostic process for software orbit data for extrapolating orbit data of a navigation satellite according to an embodiment of the present invention;
FIG. 4 is a schematic diagram of a software orbit source data backup process for extrapolating the orbit of a navigation satellite according to an embodiment of the invention.
Detailed Description
The method for designing the extrapolation software of the navigation satellite orbit according to the present invention will be described in further detail with reference to the accompanying drawings and the specific embodiments. Advantages and features of the invention will become more apparent from the following description and from the claims. It should be noted that the drawings are in a very simplified form and are all to a non-precise scale, merely for convenience and clarity in aiding in the description of embodiments of the invention.
In addition, features of different embodiments of the invention may be combined with each other, unless otherwise specified. For example, a feature of the second embodiment may be substituted for a corresponding feature of the first embodiment, or may have the same or similar function, and the resulting embodiment would fall within the disclosure or scope of the disclosure.
The invention provides a design method of navigation satellite orbit extrapolation software, which aims to solve the problem that the stability of the existing navigation satellite is difficult to improve.
In order to realize the above idea, the invention provides a navigation satellite orbit extrapolation software design method, which comprises the following steps: as shown in fig. 2, the track data preprocessing module receives the number of the observing and controlling injection tracks, the navigation satellite ephemeris and astronomical navigation attitude data as track data to be calculated; the track data backup and recovery module performs data backup on the number of the track to be observed and controlled and the number of the track to be injected; the track source preprocessing self-checking module performs self-checking on the track data to be calculated according to the check code and performs mutual checking on the track data to be calculated; the track source diagnosis module selects the number of the track to be measured and controlled and the satellite ephemeris or astronomical navigation attitude data to be used as track extrapolation calculation data according to a track source diagnosis algorithm; the orbit extrapolation calculation module utilizes the orbit extrapolation calculation data to obtain real-time orbit data of the satellite, and the real-time orbit data is used for closed-loop calculation of a satellite attitude control system; and the track source main selection switching module performs autonomous selection and switching of the track data to be calculated according to the track source priority, the track switch state and the track data self-diagnosis result.
The design method of the navigation satellite orbit extrapolation software is provided with the navigation satellite orbit extrapolation software, and the main function of the method is to obtain real-time orbit data of the satellite by using an extrapolation or autonomous orbit determination method for closed-loop calculation of the satellite attitude control software. The orbit data used for orbit extrapolation has 3 sources, and the number of the injected orbit, the ephemeris of the navigation satellite and the astronomical navigation attitude data are measured and controlled.
The navigation satellite orbit extrapolation software mainly has several functions, namely orbit data input interface management, orbit data backup and recovery functions, orbit source preprocessing and self-checking functions, orbit source diagnosis functions, orbit extrapolation calculation functions and orbit source main selection switching functions.
The navigation satellite orbit extrapolation software is mainly characterized in that: based on satellite software functional design, three main direct sources of on-board orbit data are: the measurement and control system determines the number of the orbit ground injection, obtains the position speed under WGS-84 from the load ephemeris, and performs astronomical orbit calculation orbit data according to the satellite attitude sensitive information. When orbit data is injected into the orbit ground without the measurement and control system, orbit calculation can be carried out by means of load ephemeris or orbit extrapolation calculation can be carried out by means of astronomical navigation orbit. And adopting a priority management strategy for the three track sources, and selecting an optimal track for use by a track source data mutual checking mechanism. The uninterrupted switching among different track sources is realized, and the track data input to the gesture control is ensured to be continuously and effectively. A strategy of track source diagnosis is adopted to ensure the correctness and usability of the selected track source data.
As shown in fig. 1, orbit source data, measurement and control uploading orbit data, ephemeris data generated by a navigation task processor, and attitude data acquired by a satellite attitude sensor are acquired first. And secondly, carrying out data stream verification on the track source data, and backing up the track source data to a local unit and a remote unit respectively. And carrying out data availability check, data preprocessing, data processing and check root number generation on the track source data again. And in addition, the mutual verification processing of the source data is carried out on the track source processed data. And finally, according to the mutual checking result of the track source data, adopting an autonomous diagnosis selection strategy to select the optimal track source data to enter the gesture control closed loop calculation.
As shown in fig. 2, first the track computation software initializes the relevant data structures: orbit operation mode, control mode, measurement and control orbit parameter, astronomical orbit parameter and ephemeris orbit parameter; after the software is initialized, checking the important data area data and recovering the track backup data; calculating the hour angles of the julian century, the julian day and the star; track data preprocessing; track source diagnostic selection; track extrapolation calculation; instruction processing;
as shown in fig. 3, after start-up, the initialization track data structure includes track algorithm switch status, track algorithm priority, algorithm self-check, mutual check flags. Track source data preprocessing outputs a track source data self-checking effective mark; judging whether the track source verification effective number is more than or equal to 2, if so, carrying out mutual verification on the track source data with the effective self-verification, and judging whether the mutual verification times are more than 2, otherwise, judging whether the current working mode is autonomous on the satellite, if so, judging whether the reserved track data are effective, otherwise, adopting a default track; judging whether the reserved track data is valid or not, if yes, checking the track with valid self-checking by adopting the reserved track according to the track source priority, and if not, adopting a default track; and if the number of the mutual check times is more than 2, selecting a track with effective mutual check according to the priority of the track source, otherwise, judging whether the current working mode is autonomous on the satellite. After the steps are completed, track calculation is performed, and the process is finished.
The track extrapolation mode of operation design includes: the working modes of the navigation satellite orbit extrapolation software are a ground control mode and an on-board autonomous mode. And under the ground control mode, the track source is selected through measuring and controlling the uplink instruction. In the autonomous on-board mode, the satellite-borne orbit extrapolation software performs autonomous selection and switching according to the orbit source priority, the orbit switching state and the orbit data self-diagnosis result.
The track source data backup design includes: as shown in fig. 4, the track data adopts a local and remote data backup method. And after the satellite-borne software receives the track data of the measurement and control uploading, respectively carrying out remote data backup in an EEPROM of the satellite-borne computer and a bus data terminal under the satellite-borne computer. The measurement and control track number is a group of flat number determined by the ground measurement and control track, and ground injection is needed according to the existing ground measurement and control resources. As the primary track source in ground mode of operation. The three-taking-two strategy is adopted for diagnosis and selection, and the measurement and control track storage is divided into three types, namely main part, backup and default. In the specific embodiment, after the track number is injected into the ground, firstly, time stamp and data validity check are carried out on the injected data, the check is successful, and then, corresponding backup position data are updated according to the storage logic.
The orbit source ephemeris orbit extrapolation calculation process comprises the following steps: as shown in fig. 2, ephemeris data is generated for a single navigation satellite load machine, the ephemeris and load time t (Zhou Namiao count in the beidou) are read from the load task processor through a 1553 bus every cycle by the spaceborne computer, and the J2000 position and speed are calculated by using the load beidou.
Calculate the reference time product seconds t1
t 1 ＝WNgro*604800+t oe
Calculating time bias
t k ＝t-t oe
Wherein t is the current Beidou, t oe Reference time for ephemeris. tk must take into account the beginning or end of the weekly transformation, i.e.: if tk is greater than 302400, tk is subtracted by 604800; if tk is less than-302400, tk is added to 604800.
Calculating the angle of the point at the nearest point
A 0 ＝27906100+ΔA
n＝n 0 Wherein, the sum of the values of +Deltan,μ＝3.986004418e 14
M k ＝M 0 +nt k ，
calculating the angle of the closest point
M k ＝E k -esinE k
The iteration method is used for calculation, and the maximum number of iterations is 60, and each iteration is calculated as follows:
E k+1 ＝E k -(E k -e*sinE k -M k )/(1-e*cosE k )
if E k+1 -E k ≤1e -6 The iteration ends.
Calculating phi k
Φ k ＝V k +ω, wherein,
calculating δu k 、δr k 、δi k
Calculation u k 、r k 、i k
Calculated geodetic position XYZ
Wherein,
calculated geodesic velocities Vx, vy, vz
Wherein omega e ＝7.292115E-5rad/s
Wherein,
ground-fixed system position and speed-J2000 position and speed; r is (r) J ＝(G·N·P) T ·r W Wherein r is J For the J2000 position r W Is a ground fixation system position;wherein (1)>Speed J2000>Is the ground fastening speed->ω e ＝[0,0,ω e ]Is the earth rotation rate vector; the ephemeris data format is shown in Table 1:
TABLE 1
/>
The track source measurement and control track extrapolation calculation process comprises the following steps: and (5) performing root extrapolation by adopting a J2 perturbation model according to the dynamics principle of the satellite. The extrapolation process mainly comprises Kepler root conversion and utilization of t 0 Number of time-of-day squaresCalculating t i Time of day of the number of roots>By t i Time of day of the number of roots>Calculating t i Instantaneous root number sigma at moment, utilization t i Calculating t by instantaneous root sigma of moment i Position velocity data of time satellite in J2000 coordinate system +.>
By t 0 Number of time-of-day squaresCalculate the number of roots at time t +.>/>
Wherein: />
Using the number of squares at time tCalculating the instantaneous root number sigma (a, i, omega, zeta, eta, lambda) at the moment t
△σ s The expression of (2) is as follows:
if the track extrapolation semi-major axis check enable flag is enabled, checking the transient root number semi-major axis as follows:
|a×6378.137-28125|≤△a
wherein: Δa defaults to 1000km, Δa may be altered by data betting.
If the conditions are met, checking to be correct, setting an extrapolation effective mark of the measurement and control track as effective, and converting the instantaneous root number into position speed data; and (3) checking errors, setting an extrapolation valid mark of the measurement and control track as invalid, and keeping the last value of the extrapolated track data. If 5 consecutive checks are wrong, the fiducial point use flag is set to the old fiducial point and if there is only one set, the switch is made to the default extrapolation point.
If the track extrapolation semi-long axis check enable flag is disabled, an extrapolated track valid flag is set to valid.
Calculating coordinates of satellite at t moment in J2000 coordinate system by using instantaneous root number sigma at t momentAnd speed->The units are the guard length unit and the guard length unit/guard time unit respectively.
/>
Wherein,
p＝a
the calculation method of u is as follows:
ω＝a tan 2(-η,ξ)
M＝λ-ω
f＝M+2e sin M
u＝ω+f
wherein atan2 represents a two-dimensional arctangent function.
Coordinates are setAnd speed->Units are converted to meters and meters per second.
Track source astronomical track extrapolation calculation process
And filtering the simplified orbit model by utilizing the satellite star sensitive quaternion and the ground sensitive geocentric vector through UKF to obtain the current inertial system position and speed vector of the satellite.
Obtaining current measurements
Quaternary number Q of star-sensitive attitude determination inertial system bi Conversion to a gesture matrix R bi By R bi And a geodetic attitude-determining geodetic vector E b Calculating the measured value y of the geocentric vector of the inertial system at the current moment k The method comprises the following steps:
calculate Sigma points
P pair P x,k-1 The matrix is subjected to Cholesky decomposition to obtain P x,k-1 Square root G of the array:
wherein pij is P x,k-1 Elements i, j E [1,6 ]]。
Equation of time propagation
And calculating the predicted value of the Sigma point according to the system state equation.
Taking a two-body model represented by six numbers as a state equation, and a recursive algorithm is as follows:
step1: converting position and speed into six track numbers
Wherein->
Step2: calculating the six numbers of the next period:
wherein->
Ek is calculated by using an iterative method, and is iterated for 60 times at most, and each iteration is calculated as follows: ek (Ek) k+1 ＝M+e*sin(Ek k ). If E k+1 -E k ≤1e -6 Ending the iteration;
step3: conversion of six track numbers into position and speed
Wherein->/>
Wherein->
Calculating a predicted value of a state quantity
Calculating predictive value of covariance matrix
Calculating a predicted value of the observed quantity
The observed quantity is a satellite inertial system geocentric vector, and is obtained by using the satellite position inverse sign in the state quantity and normalizing.
And calculating a predicted value of the output quantity.
Measurement update equation
Calculating gain arrays
Updating covariance matrix
Updating state quantity
The orbit source data autonomous diagnostic strategy comprises: and adopting a policy of priority management for the track source, wherein the priority can be modified by filling. Data preprocessing operation is firstly carried out on the track sources respectively. And the orbit source 1 adopts a J2 perturbation model to extrapolate the root number according to the dynamics principle of the satellite to calculate the position and the velocity of the current inertial system. The orbit source 2 calculates the position and velocity of the current inertial frame from the ephemeris and time T. And the orbit source 3 filters the simplified orbit model through UKF according to the star-sensitive quaternion and the ground-sensitive geocentric vector, and calculates the position and the speed of the current inertial system.
In summary, the above embodiments describe in detail different configurations of the software design method for extrapolating the navigation satellite orbit, however, the present invention includes but is not limited to the configurations listed in the above embodiments, and any contents of transformation based on the configurations provided in the above embodiments fall within the scope of the present invention. One skilled in the art can recognize that the above embodiments are illustrative.
In the present specification, each embodiment is described in a progressive manner, and each embodiment is mainly described in a different point from other embodiments, and identical and similar parts between the embodiments are all enough to refer to each other. For the system disclosed in the embodiment, the description is relatively simple because of corresponding to the method disclosed in the embodiment, and the relevant points refer to the description of the method section.
The above description is only illustrative of the preferred embodiments of the present invention and is not intended to limit the scope of the present invention, and any alterations and modifications made by those skilled in the art based on the above disclosure shall fall within the scope of the appended claims.
Claims (4)
1. A method for designing navigation satellite orbit extrapolation software, comprising:
the orbit data preprocessing module receives the number of the observing and controlling injection orbit, the ephemeris of the navigation satellites and the astronomical navigation attitude data as orbit data to be calculated;
the track data backup and recovery module performs data backup on the number of the track to be observed and controlled;
the track source preprocessing self-checking module performs self-checking on track data to be calculated according to the check code and performs mutual checking on the track data to be calculated;
the orbit self-diagnosis module selects the number of the observing and controlling injection orbit according to an orbit self-diagnosis algorithm, and the navigation satellite ephemeris or astronomical navigation attitude data is used as orbit extrapolation calculation data;
the orbit extrapolation calculation module obtains real-time orbit data of the satellite by using the orbit extrapolation calculation data, and the real-time orbit data is used for closed-loop calculation of a satellite attitude control system;
the track source main selection switching module performs autonomous selection and switching of track data to be calculated according to the track source priority, the track switch state and the track data self-diagnosis result;
the system comprises a first orbit source, a second orbit source, a navigation satellite ephemeris, a third orbit source and an astronomical navigation attitude data, wherein the number of the measurement and control upper injection orbits is a first orbit source and comprises the number of the flat roots injected into the orbital ground of a measurement and control system, the navigation satellite ephemeris is a second orbit source and comprises the step of obtaining the position speed under WGS-84 from the load ephemeris, and the astronomical navigation attitude data is a third orbit source and comprises orbit data for astronomical orbit calculation according to satellite attitude sensitive information;
when the number of the orbit-determining ground surface injection is flat, carrying out orbit calculation according to the load ephemeris or carrying out orbit extrapolation calculation through satellite attitude sensitive information;
the first track source, the second track source and the third track source are subjected to priority management, and an optimal track is selected by a track source data mutual checking mechanism so as to realize uninterrupted switching among different track sources and ensure that track data input to gesture control are continuously and effectively;
the track data to be calculated is subjected to local data backup and remote data backup;
after receiving the orbit data of the measurement and control betting, the navigation satellite orbit extrapolation software respectively performs remote data backup in the EEPROM of the spaceborne computer and the bus data terminal under the spaceborne computer;
diagnosing and selecting by adopting a three-in-two strategy, and storing the number of the measurement and control injection tracks into main data, backup data and default data;
after the track number is injected into the ground, performing time stamp and data validity verification on the injected data, and updating corresponding backup position data according to the storage logic after the verification is successful;
the ephemeris data are generated for a single loading machine of a navigation satellite, the ephemeris and the loading time t are read from a loading task processor through a 1553 bus every cycle by a satellite-borne computer, and the J2000 position and the J2000 speed are calculated by using the Beidou loading;
calculating a reference time product second t1:
t 1 ＝WNgro*604800+t oe
calculating time deviation:
t k ＝t-t oe
wherein t is the current Beidou, t oe Reference time for ephemeris; considering the beginning or end of the week transform, if tk is greater than 302400, tk is subtracted by 604800; if tk is less than-302400, then tk is added to 604800;
calculating a straight-up point angle:
A 0 ＝27906100+ΔA
n＝n 0 +Δn, whereinμ＝3.986004418e 14
M k ＝M 0 +nt k ，
Calculating a close point angle:
M k ＝E k -esinE k
the iteration method is used for calculation, and the maximum number of iterations is 60, and each iteration is calculated as follows:
E k+1 ＝E k -(E k -e*sinE k -M k )/(1-e*cosE k )
if E k+1 -E k ≤1e -6 Ending the iteration;
calculation F k ：
F k ＝V k +ω, wherein,
calculating δu k 、δr k 、δi k ：
Calculation u k 、r k 、i k ：
Calculated geodetic position XYZ:
wherein,
calculated geodesic velocities Vx, vy, vz:
wherein omega e ＝7.292115E-5rad/s
Wherein,
ground-fixed system position and speed-J2000 position and speed;
r J ＝(G·N·P) T ·r W wherein r is J For the J2000 position r W Is a ground fixation system position;
wherein,speed J2000>Is the ground fastening speed->
ω e ＝[0,0,ω e ]Is the earth rotation rate vector;
the root extrapolation is carried out by adopting a J2 perturbation model according to the dynamics principle of the satellite; the extrapolation process comprises Kepler root conversion and utilization of t 0 Number of time-of-day squaresCalculating t i Time of day of the number of roots>By t i Time of day of the number of roots>Calculating t i Instantaneous root number sigma at moment, utilization t i Calculating t by instantaneous root sigma of moment i Position and velocity data of time satellite in J2000 coordinate system
By t 0 Number of time-of-day squaresCalculate the number of roots at time t +.>
Wherein: />
Using the number of squares at time tCalculating the instantaneous root sigma (a, i, omega, ζ, eta, lambda) at the moment t:
wherein Δσ s The expression of (2) is as follows:
if the track extrapolation semi-major axis check enable flag is enabled, checking the transient root number semi-major axis as follows:
|a×6378.137-28125|≤Δa
wherein: the default value of delta a is 1000km, and the delta a is changed through data uploading;
if the conditions are met, checking to be correct, setting an extrapolation effective mark of the measurement and control track as effective, and converting the instantaneous root number into position speed data; checking errors, setting an extrapolation valid mark of the measurement and control track as invalid, and keeping the last value of the extrapolated track data; if the error is checked for 5 times continuously, setting the reference point use mark as an old reference point, and if only one group exists, switching to a default extrapolation point;
if the track extrapolation semi-long axis verification enabling mark is forbidden, an extrapolation track valid mark is set to be valid;
calculating coordinates of satellite at t moment in J2000 coordinate system by using instantaneous root number sigma at t momentAnd speed->The units are a guard length unit and a guard length unit/guard time unit respectively;
wherein,
p＝a
the calculation method of u is as follows:
ω＝atan2(-η,ξ)
M＝λ-ω
f＝M+2esinM
u＝ω+f
wherein atan2 represents a two-dimensional arctangent function;
coordinates are setAnd speed->Units are converted to meters and meters per second.
2. The method for designing navigation satellite orbit extrapolation software as claimed in claim 1,
the working modes of the navigation satellite orbit extrapolation software comprise a ground control mode and an on-board autonomous mode;
under a ground control mode, selecting a first track source through measuring and controlling an uplink instruction;
in the autonomous on-board mode, the navigation satellite orbit extrapolation software performs autonomous selection and switching according to the orbit source priority, the orbit switching state and the orbit data self-diagnosis result.
3. The method for designing navigation satellite orbit extrapolation software as claimed in claim 1,
the satellite star sensitive quaternion and the ground sensitive geocentric vector are utilized, and the simplified orbit model is filtered through an unscented Kalman filter to obtain the current inertial system position and speed vector of the satellite;
acquiring a current measured value, and determining a quaternion Q of a attitude-determining inertial system of the star sensor bi Conversion to a gesture matrix R bi By R bi And a geodetic attitude-determining geodetic vector E b Calculating the measured value y of the geocentric vector of the inertial system at the current moment k ：
Sigma points were calculated:
p pair P x,k-1 The matrix is subjected to Cholesky decomposition to obtain P x,k-1 Square root G of the array:
wherein pij is P x,k-1 Elements i, j E [1,6 ]]；
Establishing a time propagation equation, and calculating a predicted value of the Sigma point according to a system state equation;
taking a two-body model represented by six numbers as a state equation, and calculating a predicted value of the state quantity:
calculating a predicted value of the covariance matrix:
calculating a predicted value of the observed quantity:
the observed quantity is a satellite inertial system geocentric vector, and is obtained by using the satellite position inverse number in the state quantity and normalizing;
calculating a predicted value of the output quantity:
measuring an update equation, and calculating a gain array:
updating the covariance matrix:
updating state quantity:
4. a method for designing software for extrapolating orbit of navigation satellite according to claim 3, wherein a recursive algorithm using a two-body model represented by six numbers as the state equation is as follows:
converting position and speed into six track numbers
Wherein->
Calculating the six numbers of the next period:
wherein the method comprises the steps of
Ek is calculated by using an iterative method, and is iterated for 60 times at most, and each iteration is calculated as follows: ek (Ek) k+1 ＝M+e*sin(Ek k ) The method comprises the steps of carrying out a first treatment on the surface of the If E k+1 -E k ≤1e -6 Ending the iteration;
conversion of six track numbers into position and speed
Wherein->
Wherein->
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2001033349A2 (en) * | 1999-11-03 | 2001-05-10 | Accenture Llp | Architectures for netcentric computing systems |
| CN103542854A (en) * | 2013-11-02 | 2014-01-29 | 中国人民解放军国防科学技术大学 | Autonomous orbit determination method based on satellite borne processor |
| CN109521448A (en) * | 2018-12-18 | 2019-03-26 | 清华大学 | Satellite-based navigation receiver positioning time service method and device based on orbital tracking prediction |
| CN110553653A (en) * | 2019-08-23 | 2019-12-10 | 上海航天控制技术研究所 | spacecraft orbit determination method based on multi-source data driving |
| CN110865902A (en) * | 2019-10-25 | 2020-03-06 | 上海航天控制技术研究所 | Batch processing method for storage and upfilling diagnosis of pipeline navigation reference track |
| CN112014869A (en) * | 2020-08-12 | 2020-12-01 | 中国科学院微小卫星创新研究院 | Astronomical navigation-based inter-satellite link autonomous navigation method and system |
-
2020
- 2020-12-30 CN CN202011621293.7A patent/CN112836339B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2001033349A2 (en) * | 1999-11-03 | 2001-05-10 | Accenture Llp | Architectures for netcentric computing systems |
| CN103542854A (en) * | 2013-11-02 | 2014-01-29 | 中国人民解放军国防科学技术大学 | Autonomous orbit determination method based on satellite borne processor |
| CN109521448A (en) * | 2018-12-18 | 2019-03-26 | 清华大学 | Satellite-based navigation receiver positioning time service method and device based on orbital tracking prediction |
| CN110553653A (en) * | 2019-08-23 | 2019-12-10 | 上海航天控制技术研究所 | spacecraft orbit determination method based on multi-source data driving |
| CN110865902A (en) * | 2019-10-25 | 2020-03-06 | 上海航天控制技术研究所 | Batch processing method for storage and upfilling diagnosis of pipeline navigation reference track |
| CN112014869A (en) * | 2020-08-12 | 2020-12-01 | 中国科学院微小卫星创新研究院 | Astronomical navigation-based inter-satellite link autonomous navigation method and system |
Non-Patent Citations (6)
| Title |
|---|
| "BRDC导航星历轨道计算算法梳理";远方的驼队;《链接：https://www.cnblogs.com/whugreyhound/p/8458826.html》;第1-3页 * |
| "北斗卫星整网集中式自主定轨算法研究";林夏等;《吉林大学学报(信息科学版)》;第428-432页 * |
| "卫星天文导航自主定轨精度及误差分析";季玮等;《电子设计工程》;第90-93页 * |
| "卫星轨道外推算法实现";刘成等;《电子设计工程》;第132-134页 * |
| "遥感卫星的轨道外推方法";杨海峰等;《全球定位系统》;第1-4页 * |
| "采用多敏感器相互校正的卫星入轨姿态计算方法";李笑月等;《国防科技大学学报》;第24-29页 * |
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