CN110132267B - Optical fiber inertial navigation system of air-space-ground integrated aircraft and optical fiber inertial navigation on-orbit alignment method - Google Patents

Optical fiber inertial navigation system of air-space-ground integrated aircraft and optical fiber inertial navigation on-orbit alignment method Download PDF

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CN110132267B
CN110132267B CN201910389490.1A CN201910389490A CN110132267B CN 110132267 B CN110132267 B CN 110132267B CN 201910389490 A CN201910389490 A CN 201910389490A CN 110132267 B CN110132267 B CN 110132267B
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inertial navigation
attitude
orbit
optical fiber
speed
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CN110132267A (en
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吴枫
邵添羿
谷丛
幸伟
刘美霞
林建华
董建腾
张铭涛
姜峰
萨日娜
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Shanghai Aerospace Control Technology Institute
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    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
    • G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
    • G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
    • G01C21/165—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation combined with non-inertial navigation instruments
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C25/00—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass
    • G01C25/005—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass initial alignment, calibration or starting-up of inertial devices

Abstract

The invention discloses an inertial navigation on-orbit alignment method, which is suitable for an aerospace-ground integrated aircraft and comprises the following steps: judging that the aerospace-ground integrated aircraft is in an on-orbit state according to the tabulated data of the inertial navigation of the aircraft and the GNSS speed information; constructing filtering quantity measurement by using the star sensor information, the GNSS information and the attitude, speed and position information resolved by the inertial navigation system, and finishing the binding of the initial values of the attitude, speed and position of the aircraft in orbit through a Kalman filter; and meanwhile, analyzing an on-orbit alignment fault mechanism of the air-space-ground integrated aircraft inertial navigation system, and giving a judgment threshold value. The method effectively solves the problem of determining the initial attitude, speed and position of the air-space-ground integrated aircraft optical fiber inertial navigation system during the orbit, provides a fault judgment mechanism, and has the advantages of high alignment precision and easy realization of engineering.

Description

Optical fiber inertial navigation system of air-space-ground integrated aircraft and optical fiber inertial navigation on-orbit alignment method
Technical Field
The invention relates to the field of aerospace, in particular to an in-orbit alignment method for an optical fiber inertial navigation system of an aerospace-ground integrated aircraft.
Background
The air-space-ground integrated aircraft has the functions of an aviation platform and a space-flight platform, can perform hypersonic flight in the atmosphere like a common airplane, can enter the space like a rocket and a satellite, and can flexibly perform orbital maneuver flight. The air-ground integrated aircraft has revolutionary progress such as reusability, low launching cost, maintainability, turnaround time, flexible maneuverability and the like, also has great military potential, and is a preferred weapon equipment platform for controlling space and competing for astrology in the future.
The whole flight task of the air-space-ground integrated aircraft spans two different fields of aviation and aerospace, when the air-space-ground integrated aircraft is in orbit, the speed of the aircraft is very high, the specific force sensitive to an accelerometer is basically zero, if the inertial navigation of the aircraft is restarted due to an accident, the inertial navigation cannot determine the initial attitude through a conventional initial self-alignment mode, and a new alignment strategy needs to be searched to complete the determination of the initial attitude, the speed and the position of the inertial navigation of the aircraft in the orbit process.
In published documents in recent years, more researches on the alignment of the conventional ground compass are carried out, but the researches on the alignment of the conventional ground compass and an aircraft in orbit are basically not carried out, and a corresponding technical scheme is not available for solving the problems that the navigation precision of the aircraft in the initial stage of the orbit is poor or the initial attitude, the speed and the position of the aircraft cannot be determined due to the inertial navigation restart.
Disclosure of Invention
In view of the defects of the prior art, the technical problems to be solved by the invention are as follows: aiming at the special space environment of the aircraft in orbit, an optical fiber inertial navigation system of an air-space-ground integrated aircraft and an optical fiber inertial navigation in-orbit alignment method are provided, and the problems that the navigation precision of the aircraft in the initial orbit stage is poor or the initial attitude, speed and position cannot be determined due to the restart of the aircraft inertial navigation are solved.
The invention provides an inertial navigation on-orbit alignment method, which is applied to an aerospace-ground integrated aircraft and comprises the following steps:
s1, performing interrupt sampling on the three gyros and the three summers by inertial navigation of the air-space-ground integrated aircraft, performing attitude updating, speed updating and position updating by using the acquired gyros and summers, resolving in real time to obtain attitude, speed and position information of the aircraft, and periodically receiving external GNSS data and star sensor data;
s2, judging whether the aircraft is in an on-orbit state or not according to the acquired addend data and the GNSS speed information by inertial navigation of the air-space-ground integrated aircraft, and ensuring that the GNSS data and the star sensor data are effective at the same time at the on-orbit alignment starting moment;
further, determining whether the aircraft is in an on-track condition includes determining that,
determination condition a 1: the speed of the aircraft is high, the acceleration sensed by the inertial navigation accelerometer is small, and the speed information v of the GNSS is>1000(m/s) and the inertial navigation specific force information f < 3 (m/s)2)。
S3, on the premise that the step S2 is effective, constructing Kalman measurement by using the received star sensor information and GNSS information and the inertial navigation attitude, speed and position information of the aircraft, and performing on-orbit alignment through a Kalman filter;
further, the above-mentioned measuring process includes,
firstly, selecting a north-east-ground coordinate system as a navigation reference system (n system), a front-right-lower coordinate system as a carrier coordinate system (b system), and an inertial coordinate system as an i system;
obtaining initial values of attitude, speed and position of Strapdown Inertial Navigation (SINS) at the end time of on-orbit alignment through Kalman filtering, and selecting an error state as an attitude misalignment angle phinVelocity error δ vinnsPosition error δ pinsGyro constant zero offset epsilonb，
Accelerometer constant bias +b。
Further, the kalman filtering process described above includes,
establishing 15-dimensional Kalman filter state variables
Figure GDA0002845970680000021
Constructing SINS/GNSS Kalman filtering quantity measurement by using the difference between the received GNSS data and the inertial navigation speed position, wherein the measurement equation of the measurement is
Figure GDA0002845970680000022
In the above formula
Figure GDA0002845970680000023
pinsThe velocity and position values of inertial navigation are obtained;
Figure GDA0002845970680000024
pgnssfor GNSS second pulse tPMeasuring velocity and position at a time; w is avAnd wpRespectively, velocity and position measurement noise.
Before the measurement equation is constructed, a coordinate system is unified, an attitude matrix output by the star sensor is relative to an i system, an attitude matrix output by the strapdown inertial navigation is relative to an n system, and CNS output is an inertial attitude Cb iConstructing a matching measurement using the SINS navigation result
Figure GDA0002845970680000025
The structure mode is as follows:
Figure GDA0002845970680000026
wherein n ispAnd ncRespectively representing the navigation coordinate systems determined by the calculated attitude and the calculated position, and setting the navigation coordinate system at the real position as nt. Then the formula (3) can be unfolded into
Figure GDA0002845970680000031
Will be provided with
Figure GDA0002845970680000032
Is decomposed into
Figure GDA0002845970680000033
And take into account
Figure GDA0002845970680000034
δ θ is a position error angle.
The position error angle delta theta is in the north-east-ground coordinate system
Figure GDA0002845970680000035
Further, let the star sensor measure the error as vcnsThen the inertial attitude of the output can be expressed as
Figure GDA0002845970680000036
By using
Figure GDA0002845970680000037
And
Figure GDA0002845970680000038
constructing a measurement matching matrix, and determining a measurement matching matrix,
Figure GDA0002845970680000039
measurement of extracted quantity from equation (8)
Figure GDA00028459706800000310
The measurement equation can be obtained as
Figure GDA00028459706800000311
Further, the above-mentioned condition for discriminating the alignment fault includes one or a combination of the following two,
a) in the in-orbit alignment process, aligning faults when the combined filtering correction times of inertial navigation and GNSS of the aircraft are less than 80% of in-orbit alignment time;
b) in the in-orbit alignment process, the alignment fault is detected when the inertial navigation of the aircraft and the satellite sensitive combined filtering correction times are less than 40% of in-orbit alignment time.
In addition, the invention provides an inertial navigation on-orbit alignment system of an aerospace-ground integrated aircraft, which comprises the following modules:
the calculation module is used for calculating and obtaining the attitude, speed and position information of the aircraft in real time;
the receiving module is used for periodically receiving external GNSS data and star sensor data;
the judging module is connected with the calculating module and the receiving module, judges whether the aircraft is in an on-orbit state or not according to the acquired adding data and the speed information, and simultaneously ensures that the GNSS data and the star sensor data are simultaneously effective at the on-orbit alignment starting moment;
the measuring module is connected with the calculating module and the receiving module, measures the matching quantity of the attitude construction attitude calculated by utilizing the information of the star sensor and the inertial navigation system, and measures the speed and position quantity according to the difference between the GNSS data and the inertial navigation speed and position;
the Kalman filter module is connected with the measurement module, and is used for carrying out Kalman filter on the speed position measurement quantity to obtain initial values of the inertial navigation attitude, the speed and the position at the time of finishing the on-orbit alignment when judging whether the aircraft is in an on-orbit state to take effect;
and the feedback module is used for providing an alignment fault judgment condition aiming at the on-orbit alignment mechanism, setting a corresponding threshold value and judging the alignment fault.
Compared with the prior art, the optical fiber inertial navigation system and the optical fiber inertial navigation on-orbit alignment method of the air-space-ground integrated aircraft have the advantages that,
(1) the conventional equipment-GNSS and star sensor of the air-space-ground integrated aircraft are fully utilized, the on-orbit initial value binding and error correction are carried out on the inertial navigation of the aircraft, and the alignment precision is high;
(2) the method is simple in implementation, and the Kalman filter quantity measurement structure is easy;
(3) on-orbit alignment is realized, and on-orbit alignment fault detection is realized, so that alignment faults can be timely and accurately judged, and the reliability of the system is improved;
in conclusion, the inertial navigation on-orbit alignment method provided by the invention utilizes the GNSS and the star sensor to assist the inertial navigation to perform alignment, can accurately judge alignment faults, and is simple, feasible, high in alignment precision and reliability, easy in engineering realization and high in engineering application value.
Drawings
FIG. 1 is a schematic flow chart of an inertial navigation on-orbit alignment method of an aerospace-ground integrated aircraft provided by the invention;
FIG. 2 is a 12-hour attitude dynamics trajectory diagram of an aerospace-geostationary integrated aircraft under a navigation coordinate system in engineering practice by adopting the method provided by the invention;
FIG. 3 is a graph of a 12-hour speed dynamics trajectory of an aerospace-geostationary integrated aircraft under n systems in engineering practice by using the method provided by the invention;
FIG. 4 is a diagram of a 12-hour position dynamics trajectory of an aerospace-geostationary integrated aircraft under n systems in engineering practice by adopting the method provided by the invention;
FIG. 5 is a graph of error curves of attitude and dynamic attitude of air-ground integrated aircraft inertial navigation in-orbit alignment in engineering practice by using the method provided by the invention;
FIG. 6 is a graph of error curves of velocity and dynamic velocity of inertial navigation in-orbit alignment of a aerospace integrated aircraft in engineering practice by using the method provided by the invention;
FIG. 7 is a graph of error curves of the position and dynamic position of the air-ground integrated aircraft inertial navigation in-orbit alignment in engineering practice by adopting the method provided by the invention;
FIG. 8 is a gyro constant drift curve diagram of the inertial navigation on-orbit alignment estimation of the aerospace integrated aircraft in engineering practice by using the method provided by the invention;
FIG. 9 is a graph of an addition constant offset curve of an air-ground integrated aircraft inertial navigation on-orbit alignment estimation in engineering practice by adopting the method provided by the invention.
Detailed Description
The optical fiber inertial navigation system and the optical fiber inertial navigation on-track alignment method of the air-space-ground integrated aircraft provided by the invention are further described in detail below with reference to the accompanying drawings and specific embodiments.
Referring to fig. 1, a flow chart of an inertial navigation on-orbit alignment method for an aerospace-ground integrated aircraft provided by the present invention is shown, which includes the following steps:
step 1, resolving and outputting attitude, speed and position information of an air-space-ground integrated aircraft inertial navigation according to a sampled gyroscope and acceleration information period; meanwhile, the GNSS and the star sensor periodically send GNSS data and star sensor data to the air-space-ground integrated aircraft inertial navigation.
Step 2, judging whether the aircraft is in an in-orbit state or not by the air-space-ground integrated aircraft inertial navigation according to the adding data and the GNSS speed information, and ensuring that the GNSS data and the star sensor data are valid at the same time at the in-orbit alignment starting moment;
step 3, when the aircraft meets the on-orbit alignment condition, comparing the received star sensor information and GNSS information with the inertial navigation attitude, speed and position information of the aircraft to construct Kalman measurement, and then updating through filtering to obtain initial attitude, speed and position values of inertial navigation at the alignment finishing moment;
and 4, in the in-orbit alignment process, the alignment fault is judged in time, and the system reliability is improved.
Two points need to be explained here.
Firstly, before real-time update data obtained by inertial navigation, a required reference coordinate system is firstly solved: n-navigation coordinate system, with X, Y, Z axes pointing in the north-east-ground direction, respectively, of the geographical location of the vehicle; b-inertial unit (IMU) coordinate system, with X, Y, Z axes pointing in the front-right-down direction of the inertial measurement unit, respectively; i-the inertial coordinate system.
And performing inertial navigation resolving including attitude updating, speed updating and position updating by using the inertial navigation initial information and the acquired gyroscope and counting data, so as to acquire attitude information, speed information and position information of three axes of inertial navigation in real time.
Secondly, the inertial navigation periodically receives data of the GNSS and the star sensor, and as the air-ground integrated aircraft traverses the aerospace and aviation fields and is provided with a plurality of sensors to guarantee flight, the strapdown inertial navigation system, the GNSS and the star sensor are the most common and most basic sensing equipment. The conventional GNSS transmission period is 1s, the star sensor transmission period is 200ms, and the aircraft inertial navigation periodically receives the GNSS and star sensor data to assist in alignment or navigation.
Further, the inertial navigation on-track condition determined in step 2 is generally set as follows in the engineering.
After the air-space-ground integrated aircraft is in orbit, the flight speed of the aircraft in the special space environment of the orbit is very high (about 7800m/s), and the specific force sensitive by the inertial navigation accelerometer is basically 0. According to the special condition, the on-orbit alignment condition is easily distinguished from the ground alignment condition, and the judgment condition for setting the aircraft on the orbit is as follows: | vb|>1000(m/s) and specific force measurement | fb|＜3m/s2。
Further, constructing the in-orbit alignment kalman filter in step 3 means constructing the in-orbit alignment kalman filter when the aircraft is in the in-orbit alignment state and the GNSS and the star sensor are valid. In the filter construction, the difference between GNSS data and inertial navigation speed is used for construction measurement, the inertial navigation attitude and the star sensor attitude are unified in a coordinate system, then the construction attitude measurement is carried out, Kalman filtering calculation is carried out, various errors of the aircraft inertial navigation are estimated and corrected, and attitude, speed and position initial values at the time of finishing the on-orbit alignment are obtained. The method specifically comprises the following steps.
Step 301, establishing a state equation
The selection error state is: attitude misalignment angle phinError in velocity
Figure GDA0002845970680000061
Position error δ pinsConstant zero bias epsilon of gyrobAccelerometer constant bias
Figure GDA0002845970680000062
Establishing 15-dimensional Kalman filter state variables
Figure GDA0002845970680000063
Establishing corresponding state equation
Figure GDA0002845970680000064
In the formula FIIs a 15 multiplied by 15 dimensional state transition matrix established according to a classic strapdown inertial navigation error equation,
Figure GDA0002845970680000071
Figure GDA0002845970680000072
the system noise corresponding to the gyro and accelerometer data.
Step 302, establish SINS/GNSS measurement equation
Constructing SINS/GNSS Kalman filtering quantity measurement by using the difference between the received GNSS data and the inertial navigation speed position, wherein the corresponding measurement equation is
Figure GDA0002845970680000073
In the formula
Figure GDA0002845970680000074
pinsThe velocity and position values of inertial navigation are obtained;
Figure GDA0002845970680000075
pgnssmeasuring speed and position quantities of the GNSS; w is avAnd wpRespectively, velocity and position measurement noise.
Step 303, establishing SINS/CNS measurement equation
When measuring the structure quantity of a strapdown inertial navigation and star sensor system (CNS), the attitude matrix output by a star sensor is relative to an i system, and the attitude matrix output by an SINS is relative to an n system, and a unified coordinate system is needed firstly. Considering CNS output as inertial attitude
Figure GDA0002845970680000076
Construction of matching measurements using SINS navigation results
Figure GDA0002845970680000077
Is constructed in a manner that
Figure GDA0002845970680000078
In the formula npAnd ncRespectively representing the navigation coordinate systems determined by the calculated attitude and the calculated position, and setting the navigation coordinate system at the real position as ntThen formula (3) can be developed into
Figure GDA0002845970680000079
Will be provided with
Figure GDA00028459706800000710
Is decomposed into
Figure GDA00028459706800000711
And take into account
Figure GDA00028459706800000712
Delta theta is a position error angle with a north-east-ground coordinate system
Figure GDA00028459706800000713
Substituting into formula (5), ignoring high-order small term arrangement to obtain
Figure GDA00028459706800000714
Let the star sensor measure the error as vcnsThen the inertial attitude of the output can be expressed as
Figure GDA00028459706800000715
By using
Figure GDA00028459706800000716
And
Figure GDA00028459706800000717
constructing a measurement matching matrix, and determining a measurement matching matrix,
Figure GDA0002845970680000081
measurement of extracted quantity from equation (8)
Figure GDA0002845970680000082
The measurement equation can be obtained as
Figure GDA0002845970680000083
Further, in step 4, while in-orbit alignment, in order to improve system reliability and guarantee flight safety, an alignment fault discrimination mechanism is provided and a fault threshold is set.
The in-orbit alignment algorithm of the aircraft inertial navigation is carried out by the aid of a GNSS and a star sensor, and in order to guarantee the alignment accuracy of the inertial navigation, the filtering times of the GNSS, the star sensor and an SINS in an alignment time period are set, wherein the in-orbit alignment algorithm comprises the following two steps:
a) in the in-orbit alignment process, when the combined filtering correction times of the inertial navigation and GNSS of the aircraft is less than 80% of the in-orbit alignment time, aligning faults;
b) in the in-orbit alignment process, the alignment fault is detected when the correction times of the inertial navigation and satellite sensitivity combined filtering of the aircraft are less than 40% of the in-orbit alignment time.
When the aircraft is in-orbit aligned in inertial navigation, the selected state quantity comprises a gyro constant zero offset epsilonbAccelerometer constant bias
Figure GDA0002845970680000084
In order to further improve the alignment precision and the system reliability and tighten the alignment condition, the gyro constant zero offset estimated by a Kalman filter is set, and the condition that the added constant offset is greater than a set fault threshold value indicates an on-orbit alignment fault: filtered estimated epsilonb>1 °/h or
Figure GDA0002845970680000085
The fault is aligned.
The method of the invention is applied to engineering practice, and a semi-physical simulation system is adopted to verify and compare the invention.
The test system adopts a fiber optic strapdown inertial measurement unit, the precision of a fiber optic gyroscope is 0.01 degrees/h, the precision of an accelerometer is 40 mu g, the inertial navigation sampling period is 5ms, and the resolving period is 10 ms. The measurement error of the star sensor is 20', the measurement error of the GNSS position is 10m, and the measurement error of the GNSS speed is 0.1 m/s. The total on-track alignment time is 300 s.
Setting a dynamic track:
the orbit height h is set to 400 km, the earth radius Re6378160m, the radius of the orbit R is
R＝Re+h＝6778160m
Flight speed v and angular velocity omega of aircraftrCentripetal acceleration acAre respectively as
Figure GDA0002845970680000091
Figure GDA0002845970680000092
ac＝8.6758934(m/s2)
The rising point right ascension omega of the track is 20 degrees, and the inclination angle of the track
Figure GDA0002845970680000093
2-4 are the dynamics of the aerospace-geostationary aircraft for 12 hours under the navigation coordinate system, including attitude, velocity and position trajectories.
5-7 are error curves of attitude, speed and position of the aerospace-ground integrated aircraft inertial navigation in-orbit alignment and corresponding dynamic results. It can be seen that the algorithm has high alignment speed and high accuracy, wherein the attitude error is 0.1', the speed error is 0.05m/s, and the position error is less than 5 m.
And FIGS. 8 and 9 show a gyro constant zero offset and an addition constant offset for the air-ground integrated aircraft inertial navigation on-orbit alignment estimation. And according to the alignment fault judgment condition, the alignment is normal and has no fault under the condition.
In summary, the inertial navigation on-orbit alignment method disclosed in this embodiment has been verified by semi-physical simulation in an aircraft inertial navigation engineering prototype, and the method is feasible, the engineering technology is easy to implement, and the method has practicability. Aiming at the on-orbit special space environment, the invention adopts the GNSS and the satellite-sensitive auxiliary aircraft inertial navigation, and carries out the on-orbit alignment method through the Kalman filter, thereby overcoming the limitation that the conventional alignment algorithm can not be adopted when the measurement of the accelerometer in the space environment is zero, simultaneously accurately identifying the alignment fault, realizing the on-orbit accurate alignment of the aircraft inertial navigation, and being easy to carry out engineering.
While the present invention has been described in detail with reference to the preferred embodiments, it should be understood that the above description should not be taken as limiting the invention. Various modifications and alterations to this invention will become apparent to those skilled in the art upon reading the foregoing description. Accordingly, the scope of the invention should be determined from the following claims.

Claims (7)

1. An optical fiber inertial navigation on-orbit alignment method is suitable for an aerospace-ground integrated aircraft, and is characterized by comprising the following steps:
the method comprises the steps that the optical fiber inertial navigation of the air-space-ground integrated aircraft carries out interrupt sampling on three gyroscopes and three accelerometers, the attitude, speed and position information of the aircraft is obtained through real-time calculation, and meanwhile, external GNSS data and star sensor data are received periodically;
judging whether the aircraft is in an in-orbit state or not according to the collected accelerometer data and the GNSS speed information, and ensuring that the GNSS data and the star sensor data are valid at the same time at the in-orbit alignment starting moment;
constructing an attitude matching quantity measurement after unifying the star sensor information and an attitude resolved by an optical fiber inertial navigation system into a coordinate system, constructing a speed and position quantity measurement according to the difference between GNSS data and the speed and position of the optical fiber inertial navigation, and obtaining initial values of the attitude, speed and position of the optical fiber inertial navigation at the end moment of on-orbit alignment through a Kalman filter;
the coordinate system is unified before the measurement equation is constructed, the attitude matrix output by the star sensor is relative to the i system, the attitude matrix output by the optical fiber inertial navigation is relative to the n system, and the output by the star sensor is the inertial attitude
Figure FDA0002936907920000011
Matching measurement construction by using fiber inertial navigation result
Figure FDA0002936907920000012
The structure mode is as follows:
Figure FDA0002936907920000013
n in the formula (3)pAnd ncIndividual watchA navigation coordinate system determined by the calculated attitude and the calculated position, and a navigation coordinate system n at the real positiont(ii) a Then the formula (3) can be unfolded into
Figure FDA0002936907920000014
Will be provided with
Figure FDA0002936907920000015
Is decomposed into
Figure FDA0002936907920000016
And take into account
Figure FDA0002936907920000017
δ θ is a position error angle;
aiming at an on-orbit alignment mechanism, an alignment fault judgment condition and a corresponding threshold value are set, and an alignment fault is judged;
the alignment fault comprises one or a combination of the following two types,
a) in the in-orbit alignment process, when the combined filtering correction times of the optical fiber inertial navigation and the GNSS of the aircraft are less than 80% of the in-orbit alignment time, the fault is aligned;
b) in the in-orbit alignment process, when the correction times of the optical fiber inertial navigation of the aircraft and the satellite sensitive combined filtering are less than 40% of the in-orbit alignment time, the fault is aligned.
2. The method for in-orbit alignment of fiber optic inertial navigation according to claim 1, wherein the determining whether the aircraft is in the in-orbit state according to the collected accelerometer data and GNSS velocity information comprises,
judging a condition A1, if the speed of the aircraft is high and the acceleration sensed by the fiber inertial navigation accelerometer is small after the fiber inertial navigation of the aerospace-ground integrated aircraft is guided into the rail, the judging condition set on the rail is as follows: velocity information | v of GNSSb|>1000m/s and the fiber inertial navigation specific force information | fb|＜3m/s2。
3. The method of claim 1, wherein the measuring of the matching amount of the attitude is constructed after unifying the star sensor information and the attitude of the optical fiber inertial navigation system to form a coordinate system, and the measuring of the speed and the position is constructed according to the difference between the GNSS data and the speed and the position of the optical fiber inertial navigation system, and the initial values of the attitude, the speed and the position of the optical fiber inertial navigation system at the end of the alignment of the orbit are obtained through a Kalman filter, comprising,
selecting a north-east-ground coordinate system as a navigation reference system n system, a front-right-lower coordinate system as a carrier coordinate system b system and an inertia coordinate system as an i system;
through Kalman filtering, initial values of the attitude, the speed and the position of the optical fiber inertial navigation at the end moment of on-orbit alignment are obtained, and the selected error state is as follows: attitude misalignment angle phinError in velocity
Figure FDA0002936907920000021
Position error δ pinsConstant zero bias epsilon of gyrobAccelerometer constant bias
Figure FDA0002936907920000022
4. The method of in-orbit alignment for fiber optic inertial navigation of claim 3, wherein the Kalman filtering comprises,
establishing a 15-dimensional Kalman filter state variable as shown in the following formula (1):
Figure FDA0002936907920000023
constructing the measurement of the fiber inertial navigation/GNSS Kalman filtering quantity by using the difference between the received GNSS data and the fiber inertial navigation speed position, wherein the measurement equation of the measurement is as follows:
Figure FDA0002936907920000024
in the formula (2)
Figure FDA0002936907920000025
pinsRespectively representing the speed and position values of the optical fiber inertial navigation;
Figure FDA0002936907920000026
pgnssrespectively GNSS second pulse tPMeasuring velocity and position at a time; w is avAnd wpRespectively, velocity and position measurement noise.
5. The method of claim 4, wherein the position error angle δ θ is in a north-east-ground coordinate system
Figure FDA0002936907920000031
6. The method of claim 3, wherein the optical fiber inertial navigation in-orbit alignment method,
let the star sensor measure the error as vcnsThen the inertial attitude of the output can be expressed as
Figure FDA0002936907920000032
By using
Figure FDA0002936907920000033
And
Figure FDA0002936907920000034
constructing a measurement matching matrix, and determining a measurement matching matrix,
Figure FDA0002936907920000035
measurement of extracted quantity from equation (8)
Figure FDA0002936907920000036
The measurement equation can be obtained as
Figure FDA0002936907920000037
7. An optical fiber inertial navigation system of an aerospace-ground integrated aircraft suitable for the optical fiber inertial navigation on-orbit alignment method according to any one of claims 1 to 6, comprising the following modules:
the calculation module is used for calculating and obtaining the attitude, speed and position information of the aircraft in real time;
the receiving module is used for periodically receiving external GNSS data and star sensor data;
the judging module is connected with the calculating module and the receiving module, judges whether the aircraft is in an on-orbit state or not according to the collected accelerometer data and the GNSS speed information, and simultaneously ensures that the GNSS data and the star sensor data are simultaneously effective at the on-orbit alignment starting moment;
the measuring module is connected with the calculating module and the receiving module, measures the matching quantity of the attitude construction attitude by utilizing the information of the star sensor and the attitude construction attitude calculated by the optical fiber inertial navigation system, and measures the speed and position quantity according to the difference between the GNSS data and the speed and position of the optical fiber inertial navigation system;
the Kalman filter module is connected with the measurement module, constructs attitude matching quantity measurement after unifying the star sensor information and an attitude unified coordinate system solved by the optical fiber inertial navigation system when judging whether the aircraft is in an in-orbit state, constructs speed and position quantity measurement according to the difference between GNSS data and the speed and position of the optical fiber inertial navigation system, and obtains initial values of the attitude, the speed and the position of the optical fiber inertial navigation system at the end moment of orbit alignment through the Kalman filter;
and the feedback module is used for providing an alignment fault judgment condition aiming at the on-orbit alignment mechanism, setting a corresponding threshold value and judging the alignment fault.
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