CN106153051B - A kind of spacecraft shading device combined navigation methods - Google Patents

A kind of spacecraft shading device combined navigation methods Download PDF

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CN106153051B
CN106153051B CN201610493980.2A CN201610493980A CN106153051B CN 106153051 B CN106153051 B CN 106153051B CN 201610493980 A CN201610493980 A CN 201610493980A CN 106153051 B CN106153051 B CN 106153051B
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inertial
follows
location
correction amount
error correction
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CN106153051A (en
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王献忠
张丽敏
张肖
张国柱
程颢
汤敏兰
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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
    • G01C21/24—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 specially adapted for cosmonautical navigation
    • 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
    • G01C21/005—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 with correlation of navigation data from several sources, e.g. map or contour matching
    • 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

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  • Engineering & Computer Science (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • General Physics & Mathematics (AREA)
  • Astronomy & Astrophysics (AREA)
  • Navigation (AREA)

Abstract

The present invention relates to a kind of spacecraft shading device combined navigation methods, the star based on PI filtering is quick/the autonomous celestial combined navigation of horizon instrument/inertia, solve the problems, such as due to add table exist drift cannot the location/velocity for a long time based on inertial navigation output calculate orbit parameter.The position and speed of drift correction inertial navigation output is directed toward based on the quick the earth's core determined with horizon instrument of star; inertia-celestial combined navigation can inhibit inertial navigation accumulated error; present invention GNSS compatible on star is abnormal; and without under the ground in real time abnormal conditions such as upper note orbit parameter; utilize the quick orbit parameter corresponding with horizon instrument inertia-location/velocity calculating of celestial combined navigation output of star, it can be ensured that based on the posture of the quick determining ontology relative orbit system of star.Compared with prior art, it is simple, effective to be algorithm for advantages and beneficial effects, and is easy to Project Realization.

Description

A kind of spacecraft shading device combined navigation methods
Technical field
The present invention relates to a kind of spacecraft shading device combined navigation methods, in particular to it is a kind of based on PI filtering star it is quick/horizon instrument/ The autonomous astronomical integrated navigation method of inertia.
Background technique
Orbit parameter is needed based on the quick determining ontology relative orbit system posture of star, table is added to there is drift, inertial navigation integral causes Navigation accumulated error cannot calculate orbit parameter based on the location/velocity that inertial navigation exports for a long time.
Ground is generally basede on WGS-84 noninertial system and carries out inertial reference calculation, not applicable in high latitude area；It is in-orbit general Track reckoning is carried out based on the flat root of track infused on ground, cannot be used for a long time, and be not suitable for becoming rail situation.
It is directed toward the position and speed of drift correction inertial navigation output based on the quick the earth's core determined with horizon instrument of star, and is based on J2000 inertial system carries out inertia-celestial combined navigation and is not limited by latitude, and can inhibit inertial navigation accumulated error.
GNSS compatible is abnormal on star, and infuses under the abnormal conditions such as orbit parameter that ground is not upper in real time, quick using star Orbit parameter corresponding with horizon instrument inertia-location/velocity calculating of celestial combined navigation output, it is ensured that originally based on the quick determination of star The posture of body relative orbit system.
Summary of the invention
In order to solve the problems existing in the prior art, the object of the present invention is to provide a kind of spacecraft shading device combined navigation methods, real Star now based on PI filtering is quick/the autonomous celestial combined navigation of horizon instrument/inertia, estimated based on the quick direction deviation with horizon instrument the earth's core of star Location/velocity calibration corrections to be counted, and orbit parameter is determined based on inertia-celestial combined navigation, algorithm is simple, effective, and It is easy to Project Realization.
In order to achieve the above object, the technical solution of the present invention is to provide a kind of spacecraft shading device combined navigation methods, in which:
Based on star is quick and horizon instrument the earth's core is directed toward deviation and calculates inertial navigation location error, estimated location/speed is filtered based on PI Calibration corrections；
Inertial reference calculation is carried out based on J2000 inertial system, and carries out inertia-astronomy group based on location/velocity calibration corrections Close navigation；
Orbit parameter is calculated based on inertia-celestial combined navigation output location/velocity, and based on the quick determining celestial body phase of star To the posture of track system.
A kind of star based on PI filtering that the present invention uses is quick/the autonomous astronomical integrated navigation method of horizon instrument/inertia, and it is existing There is technology to compare, advantages and beneficial effects are:
The present invention, which has derived, is based on quick star and horizon instrument the earth's core direction deviation and PI filtering estimated location/velocity error amendment Inertia-celestial combined navigation of amount, algorithm is simple, effective, and is easy to Project Realization.
The present invention is based on J2000 inertial systems to carry out inertia-celestial combined navigation, acquires the position of opposite J2000 inertial system And speed, orbit parameter needed for being converted into the quick attitude algorithm of star seek the posture of the quick relative orbit system of star.
GNSS compatible is abnormal on star, and infuses under the abnormal conditions such as orbit parameter that ground is not upper in real time, present invention benefit With the quick orbit parameter corresponding with horizon instrument inertia-location/velocity calculating of celestial combined navigation output of star, it can be ensured that be based on The posture of the quick determining ontology relative orbit system of star.
Detailed description of the invention
Fig. 1 is the schematic diagram of spacecraft shading device combined navigation methods of the present invention.
Specific embodiment
As shown in Figure 1, the present invention provides a kind of spacecraft shading device combined navigation methods, realize the star filtered based on PI it is quick/Horizon The autonomous celestial combined navigation of instrument/inertia is directed toward estimation of deviation location/velocity calibration corrections with horizon instrument the earth's core based on star is quick, And orbit parameter is determined based on inertia-celestial combined navigation, it specifically includes:
Based on star is quick and horizon instrument the earth's core is directed toward deviation and calculates inertial navigation location error, estimated location/speed is filtered based on PI Calibration corrections；
Inertial reference calculation is carried out based on J2000 inertial system, and carries out inertia-astronomy group based on location/velocity calibration corrections Close navigation；
Orbit parameter is calculated based on inertia-celestial combined navigation output location/velocity, and based on the quick determining celestial body phase of star To the posture of track system.
The calculating process of spacecraft shading device combined navigation methods of the present invention includes:
S1, inertial reference calculation is carried out based on J2000 inertial system
If position vector of the satellite under J2000 inertia is r=[x y z]T, then weight of the satellite under J2000 inertial system Three axis component of power acceleration is as follows:
Wherein:For the distance of satellite to the earth's core, μ is Gravitational coefficient of the Earth, ReFor the earth half Diameter, J2For perturbation of earths gravitational field coefficient.
If plus the pose transformation matrix of table coordinate system to J2000 inertial system is Aia, it is non-used to acquire satellite under J2000 inertial system Property acceleration aa,iIt is as follows:
aa,i=Aia·aa,a
Wherein: aa,aFor the acceleration added under table coordinate system for adding table to measure.
Enable ag,i=[agx,i agy,i agz,i]T, acquire satellite accelerations a under J2000 inertial systemiIt is as follows:
ai=ag,i+aa,i
Inertial reference calculation is carried out in J2000 inertial system, obtains position ri, speed viIt is as follows:
vi=∫ ai·dt
ri=∫ vi·dt。
S2, q is asked by the location/velocity that inertial navigation exportsoi
If the position r under J2000 inertial systemi=[x y z]T, speedThus acquire position/ Speed scalar r/v and orbital plane normal vectorIt is as follows:
It enablesIt is as follows to acquire right ascension of ascending node Ω, orbit inclination angle i and latitude argument u:
I=cos-1(σ3/σ)
Wherein: k=σ1/ σ, h=- σ2/ σ,Quadrant is carried out in Ω, i and u solution to sentence Not.
J2000 inertial system is acquired to track system attitude quaternion q by Ω, i and uoiIt is as follows:
Wherein:
S3, quick based on star and horizon instrument asks the earth's core to be directed toward the angle of deviation
If the star quick consideration exposure time difference and data acquisition delay amendment, and after deducting installation matrix, it is opposite to obtain celestial body J2000 inertial system attitude quaternion is qbi, in conjunction with track system with respect to J2000 inertial system attitude quaternion qoi, it is opposite to acquire ontology Track system attitude quaternion qboIt is as follows:
Enable qbo=[q0 q1 q2 q3]T, it is as follows with the transition matrix that quaternary number form formula indicates:
It is as follows to turn the transition matrix that sequence indicates with Eulerian angles 3-1-2:
Wherein:For roll attitude angle, θ is pitch attitude angle, and ψ is yaw-position angle.
It is as follows to acquire celestial body relative orbit system posture:
The earth's core direction angle of deviation approximate can resolve as follows:
D θ=θ-θH
Wherein,Roll angle, θ for horizon instrument outputHFor the pitch angle of horizon instrument output.
S4, it is based on the earth's core error in pointing angular estimation location error
It is directed toward by relative orbit system the earth's core and rolls the angle of deviationPitch deviation angle d θ and inertial reference calculation output the earth's core away from R acquires under track system X to as follows with Y-direction location error:
Dx=rd θ
Consider the coupled relation of dz and dx, it is as follows using ratio estimate that dz is based on dx:
Dz=kpxz·dx
Wherein: kpxzProportion estimation.
Enable error delta r in position under track systemo=[dx dy dz]T, it is transformed into error delta r in position under J2000 inertial systemiSuch as Under:
Δri=Aoi T·Δro
Wherein: AoiIt is inertial system to track system transition matrix, by qoiIt acquires.
S5, estimated location/velocity error correction amount is filtered based on PI
Enabling based on kth step location error under the quick J2000 inertial system with horizon instrument estimation of star is Δ ri,kIf to Δ ri,kLimit Location/velocity error is after widthKth walks inertial reference calculation location error correction amount dri,kIt is as follows:
Wherein:
kp,rProportionality coefficient, k are estimated for location error correction amountp,rIt, can be with independent estimations location error for 3 × 3 diagonal matrixs Correction amount.
Enable dri,k=[dxi,k dyi,k dzi,k]TWithThree axis independent estimations positions It is as follows to set calibration corrections form:
Wherein:
kp,x、kp,y、kp,zProportionality coefficient is estimated for three shaft position calibration corrections.
In view of location error also reflects velocity error, it is based on location error correction amount drikEstimating speed calibration corrections. If to dri,kLocation error correction amount after clipping isKth walks inertial reference calculation velocity error correction amount dvi,kIt is as follows:
Wherein:
kp,vProportionality coefficient, k are estimated for velocity error correction amountp,vIt, can be with independent estimations velocity error for 3 × 3 diagonal matrixs Correction amount.
Enable dvi,k=[dvxi,k dvyi,k dvzi,k]TWithThree axis independent estimations Velocity error correction amount form is as follows:
Wherein:
kp,dx、kp,dy、kp,dzProportionality coefficient is estimated for three axle speed calibration corrections.
S6, it is based on the integrated navigation of location/velocity calibration corrections
Location/velocity calibration corrections are stepped up in inertial navigation integral process, it can be ensured that the stationarity of error correction. Location/velocity calibration corrections are deducted in J2000 inertial system, integral algorithm progress inertial reference calculation is as follows using simplifying:
vi,k=vi,k-1+[ai,k-1+(ai,k-ai,k-1)/2]·T-dvi,k-1
ri,k=ri,k-1+[vi,k-1+(vi,k-vi,k-1)/2]·T-dri,k-1
Wherein:
ai,k-1For -1 step acceleration of kth；
ai,kAcceleration is walked for kth；
dvi,k-1For the velocity error correction amount of -1 step of kth estimation；
vi,k-1The speed of velocity error correction amount is deducted for -1 step of kth；
vi,kThe speed for deducting velocity error correction amount is walked for kth；
dri,k-1For the location error correction amount of -1 step of kth estimation；
ri,k-1The position of location error correction amount is deducted for -1 step of kth；
ri,kThe position for deducting location error correction amount is walked for kth；
T is the navigation period.
S7, the posture that celestial body relative orbit system is determined based on inertia-celestial combined navigation:
Corresponding orbit parameter is calculated based on inertia-celestial combined navigation output location/velocity, is based on orbit parameter The Attitude Algorithm of celestial body relative orbit system is calculated, is referred to the relevant algorithm of the quick calculating attitude angle of star described previously.Such as with The numerical value of speed and position after deducting calibration corrections in S6, replaces the position and speed in S2 under J2000 inertial system, in turn The attitude quaternion of celestial body relative orbit system can be calculated by the process of S2 and S3.
In conclusion the present invention is based on PI filtering estimation plus the drift of table acceleration, algorithm simply, effectively, and is easy to engineering It realizes；It is not limited, can be used with whole day area by latitude based on J2000 inertial system inertial reference calculation；Based on PI filtering star it is quick with ground Level inertia-celestial combined navigation can use for a long time, and become still applicable when rail.
It is discussed in detail although the contents of the present invention have passed through above preferred embodiment, but it should be appreciated that above-mentioned Description is not considered as limitation of the present invention.After those skilled in the art have read above content, for of the invention A variety of modifications and substitutions all will be apparent.Therefore, protection scope of the present invention should be limited to the appended claims.

Claims (7)

1. a kind of spacecraft shading device combined navigation methods, which is characterized in that include following procedure:
Based on star is quick and horizon instrument the earth's core is directed toward deviation and calculates inertial navigation location error, estimated location/velocity error is filtered based on PI Correction amount；
Inertial reference calculation is carried out based on J2000 inertial system, and carries out inertia-astronomy combination based on location/velocity calibration corrections and leads Boat；
Orbit parameter is calculated based on inertia-celestial combined navigation output location/velocity, and based on the quick determining celestial body of star with respect to rail The posture of road system；
Wherein, it is quick to star according to exposure the time difference and data acquisition delay be modified, and deduct installation matrix after, obtain celestial body phase It is q to J2000 inertial system attitude quaternionbi, in conjunction with J2000 inertial system to track system attitude quaternion qoi, it is opposite to acquire ontology Track system attitude quaternion qboIt is as follows:
Enable qbo=[q0 q1 q2 q3]T, it is as follows with the transition matrix that quaternary number form formula indicates:
It is as follows to turn the transition matrix that sequence indicates with Eulerian angles 3-1-2:
Wherein:For roll attitude angle, θ is pitch attitude angle, and ψ is yaw-position angle；
It is as follows to acquire celestial body relative orbit system posture:
The earth's core is directed toward angle of deviation approximation and is resolved, and obtains relative orbit system the earth's core and is directed toward the rolling angle of deviationPitch deviation angle d θ It is as follows:
D θ=θ-θH
Wherein,Roll angle, θ for horizon instrument outputHFor the pitch angle of horizon instrument output.
2. spacecraft shading device combined navigation methods as described in claim 1, which is characterized in that wherein carried out based on J2000 inertial system The process S1 of inertial reference calculation includes:
If position r of the satellite under J2000 inertiai=[x y z]T, then acceleration of gravity three of the satellite under J2000 inertial system Axis component is as follows:
Wherein: position scalarWith it is corresponding at a distance from satellite to the earth's core, μ is Gravitational coefficient of the Earth, ReFor earth radius, J2For perturbation of earths gravitational field coefficient；
If plus the pose transformation matrix of table coordinate system to J2000 inertial system is Aia, acquire satellite under J2000 inertial system it is non-inertial plus Speed aa,iIt is as follows:
aa,i=Aia·aa,a
Wherein: aa,aFor the acceleration added under table coordinate system for adding table to measure；
Enable ag,i=[agx,i agy,i agz,i]T, acquire satellite accelerations a under J2000 inertial systemiIt is as follows:
ai=ag,i+aa,i
Inertial reference calculation is carried out in J2000 inertial system, obtains position ri, speed viIt is as follows:
vi=∫ ai·dt
ri=∫ vi·dt。
3. spacecraft shading device combined navigation methods as claimed in claim 2, which is characterized in that further include the position exported by inertial navigation Set/speed asks J2000 inertial system to track system attitude quaternion qoiProcess S2:
If the position r under J2000 inertial systemi=[x y z]T, speedAcquire position scalar r, speed Scalar v and orbital plane normal vectorIt is as follows:
It enablesAcquire right ascension of ascending node Ω, orbit inclination angle i and latitude argument u:
I=cos-1(σ3/σ)
Wherein: k=σ1/ σ, h=- σ2/ σ,
J2000 inertial system is acquired to track system attitude quaternion qoiIt is as follows:
Wherein:
4. spacecraft shading device combined navigation methods as claimed in claim 3, which is characterized in that wherein estimated based on the earth's core error in pointing angle The process S4 for counting location error includes:
It is directed toward by relative orbit system the earth's core and rolls the angle of deviationPitch deviation angle d θ and the satellite of inertial reference calculation output are to the earth's core Distance, i.e. position scalar r acquires under track system X to as follows with Y-direction location error:
Dx=rd θ
It is as follows using ratio estimate to Z-direction location error dz based on dx: dz=k using the coupled relation of dz and dxpxz·dx；
Wherein: kpxzProportion estimation；
Enable error delta r in position under track systemo=[dx dy dz]T, it is transformed into error delta r in position under J2000 inertial systemiIt is as follows:
Δri=Aoi T·Δro
Wherein: AoiFor inertial system to track system transition matrix, by J2000 inertial system to track system attitude quaternion qoiIt acquires.
5. spacecraft shading device combined navigation methods as claimed in claim 4, which is characterized in that wherein based on PI filtering estimated location/ The process S5 of velocity error correction amount includes:
Enabling based on kth step location error under the quick J2000 inertial system with horizon instrument estimation of star is Δ ri,kIf to Δ ri,kAfter clipping Location/velocity error isKth walks inertial reference calculation location error correction amount dri,kIt is as follows:
Wherein: kp,rProportionality coefficient is estimated for location error correction amount；
Enable dri,k=[dxi,k dyi,k dzi,k]TWithIt misses three axis independent estimations positions Poor correction amount form is as follows:
Wherein: kp,x、kp,y、kp,zProportionality coefficient is estimated for three shaft position calibration corrections.
6. spacecraft shading device combined navigation methods as claimed in claim 5, which is characterized in that wherein based on PI filtering estimated location/ The process S5 of velocity error correction amount is further included based on location error correction amount dri,kThe mistake of estimating speed calibration corrections Journey:
If to dri,kLocation error correction amount after clipping isKth walks inertial reference calculation velocity error correction amount dvi,kIt is as follows:
Wherein: kp,vProportionality coefficient is estimated for velocity error correction amount；
Enable dvi,k=[dvxi,k dvyi,k dvzi,k]TWithThree axis independent estimations speed Calibration corrections form is as follows:
Wherein: kp,dx、kp,dy、kp,dzProportionality coefficient is estimated for three axle speed calibration corrections.
7. spacecraft shading device combined navigation methods as claimed in claim 6, which is characterized in that wherein repaired based on location/velocity error The process S6 of positive quantity integrated navigation includes:
Location/velocity calibration corrections are deducted in J2000 inertial system, integral algorithm progress inertial reference calculation is as follows using simplifying:
vi,k=vi,k-1+[ai,k-1+(ai,k-ai,k-1)/2]·T-dvi,k-1
ri,k=ri,k-1+[vi,k-1+(vi,k-vi,k-1)/2]·T-dri,k-1
Wherein:
ai,k-1For -1 step acceleration of kth；
ai,kAcceleration is walked for kth；
dvi,k-1For the velocity error correction amount of -1 step of kth estimation；
vi,k-1The speed of velocity error correction amount is deducted for -1 step of kth；
vi,kThe speed for deducting velocity error correction amount is walked for kth；
dri,k-1For the location error correction amount of -1 step of kth estimation；
ri,k-1The position of location error correction amount is deducted for -1 step of kth；
ri,kThe position for deducting location error correction amount is walked for kth；
T is the navigation period.
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CN106767844B (en) * 2017-01-05 2019-05-28 北京航天自动控制研究所 A method of improving earth sensor body geocentric vector precision
CN107179084B (en) * 2017-06-27 2020-08-11 上海航天控制技术研究所 Pseudo-range and tabulation combined navigation and drift estimation method for GNSS compatible machine
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CN111027137B (en) * 2019-12-05 2023-07-14 中国人民解放军63620部队 High-precision dynamic construction method for spacecraft dynamics model based on telemetry data

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