CN105277195B - A kind of opposite installation error in-orbit identification method between star sensor unit - Google Patents
A kind of opposite installation error in-orbit identification method between star sensor unit Download PDFInfo
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- CN105277195B CN105277195B CN201510741062.2A CN201510741062A CN105277195B CN 105277195 B CN105277195 B CN 105277195B CN 201510741062 A CN201510741062 A CN 201510741062A CN 105277195 B CN105277195 B CN 105277195B
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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
- G01C21/02—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by astronomical means
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- 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
Abstract
The present invention provides a kind of opposite installation error in-orbit identification method between star sensor unit, including：Star sensor measures and the mathematical statistical model of installation error is established；Star sensor is established with respect to fix error angle observational equation；Fix error angle optimal estimation algorithm designs, the design of fix error angle model nonlinear approximating method.The present invention is simple and effective, does not need to calculate the posture and attitude angular velocity of spacecraft.Star sensor can be estimated with respect to fix error angle in the case of dynamical variable is unknown.In actual task implementation procedure, the present invention is than needing the real-time fix error angle method of estimation more practicability and effectiveness for obtaining spacecraft attitude, relative to traditional method of estimation, the Data Detection of the invention caused to the insensitivity of long-term loss of data on star and estimation are easier.
Description
Technical field
The present invention relates to satellite high-precision postures to determine on the star in field installation error estimation technique between sensor, specifically relates to
And a kind of opposite fix error angle in-orbit identification method between star sensor unit.
Background technology
More star sensors can be generally housed, to cope with attitude maneuver, orbital position difference leads to the sun on platform body
Light enters primary sensor field range, and then starts image-forming condition and the suitable star sensor unit of illumination condition, persistently
The platform stance information higher to precision.In addition, more star sensor redundancies use, the reliability of system can be also improved.
Due to during Spacecraft Launch impact vibration, in orbit by according to environment difference etc., causing different stars sensitive
There are long period with respect to fix error angle between device.In-orbit identification is needed to go out opposite fix error angle between different star sensors, is adopted
On the basis of mounting condition, the better primary sensor of illumination condition, the fix error angle of other star sensors is compensated.
It realizes the indifference transition in star sensor switching use, improves spacecraft attitude and determine system responding ability.
The in-orbit method of estimation of opposite fix error angle during existing star sensitivity, usually there are following defects：
(1) it needs to calculate the attitude angle of spacecraft or attitude angular velocity information, to the more demanding of acquisition of information；
(2) without applying star sensor noise characteristic itself in data estimation, precision is relatively low；
(3) more sensitive to long term data loss, the robustness of method of estimation is relatively low.
Invention content
For above-mentioned deficiency in the prior art, the object of the present invention is to provide opposite between a kind of star sensor unit
Installation error in-orbit identification method determines application demand, specific to star sensor unit based on satellite model high-precision attitude
Between proposed with respect to the estimation of fix error angle, available for estimation of the spacecraft star sensor with respect to fix error angle, and be not required to
Calculate the posture and attitude angular velocity of spacecraft.By establishing the quick opposite installation error observational equation of star, and using greatly seemingly
Right method of estimation obtains fix error angle estimated value, is a kind of opposite installation error of in-orbit star sensor based on mathematical statistics
Method of estimation.
To achieve the above object, the present invention is achieved by the following technical solutions.
Opposite installation error in-orbit identification method, includes the following steps between a kind of star sensor unit：
Step S1 establishes the mathematical statistical model of star sensor measurement and installation error；
Step S2 establishes opposite fix error angle observational equation between star sensor；
Step S3 designs opposite fix error angle optimal estimation algorithm；
Step S4 designs fix error angle model nonlinear fitting algorithm.
Preferably, the step S1, specifically includes following sub-step：
Step S101, it is as follows that derivation installation error defines method：
In formula,Matrix initial value, M are installed for ground i-th of star sensor obtained by calibratingiPacify for i-th of star sensor
Fill error matrix, θiFor i-th of star sensor installation error matrix corner vector, I is unit matrix, O ([θi]2) represent high-order
It is infinitely small；I is positive integer.
Step S102, the related definition that star sensor measures are as follows：
It is i-th of star sensor in tkThe not corrected ontology attitude matrix that moment obtains, AkIt is true for spacecraft
Attitude matrix, relationship therebetween are as follows：
It is i-th of star sensor in moment tkAttitude matrixEstimated value,Survey for i-th of star sensor
Measure error matrix, ()TRepresent transposition oeprator.Further
ξI, kIt is i-th star sensor in moment tkMeasurement noise vector, and
|ξI, k| ＜ 1
It is with very big possibility
E { ξI, k}=0
E{ξI, kξJ, k}=δijδkk′PI, k
Wherein, E { } represents mean operation, PI, kTo measure i-th of sensor in tkThe mean square deviation at moment, δijIt is quick for i star
Scaling factor between sensor and j-th of star sensor, δkk′For time tkWith time tk′Proportionality factor, δijWith δkk′Numerical value close
System is such as formula：
Preferably, in the step S2, star sensor is established in moment tkIt is as follows with respect to fix error angle observational equation：
Zk=Hkψ+ΔZk
Wherein, ZkFor observed quantity, HkFor observing matrix, ψ is quantity of state, Δ ZkFor observation noise,Hk
=I, Δ Zk~N (0, Pk) T, i.e. Δ ZkIt is 0 to obey mean value, variance PkGaussian Profile；
ZkMiddle component zIj, kObservational equation it is as follows, wherein i=1, j=2 ..., nk：
Wherein, θiAnd θjThe respectively fix error angle of i-th of star sensor and j-th of star sensor, (θj-θi) it is jth
A star sensor is with respect to the fix error angle of i-th of star sensor, quantity of state ψ as to be estimated.ξI, kAnd ξJ, kRespectively i-th
A star sensor and j-th of star sensor are in tkThe measurement noise vector at moment, Δ zIj, kFor observed quantity zIj, kObservation noise.
Available, the observation noise Δ z by formula (1)Ij, kFor star sensor i measurement noise vectors ξI, kIt is sweared with star sensor j measurement noises
Measure ξJ, kDifference.Observed quantity zIj, kAcquisition methods it is as follows：
[zIj, k]lmFor matrix [zIj, k] in l rows, m row element, matrix [zIj, k] computational methods it is as follows
It is i-th of star sensor in moment tkObtained ontology attitude matrixEstimated value,For j-th of star
Sensor is in moment tkObtained ontology attitude matrixThe transposed matrix of estimated value.
Preferably, in the step S3, quantity of state ψ is estimated using maximum likelihood optimal estimation method.Greatly seemingly
Right equation is as follows
(Ψ) is minimized ψ, obtains following equation
Wherein
Quantity of state estimated value Ψ*(prior-free) it is as follows
∑ is adduction oeprator,For the equation matrix of observation noise, ()-1Represent inverse matrix operation symbol.
Preferably, it is sensitive to star in the entire orbital period using nonlinear least square fitting method in the step S4
The fundamental equation of device time-varying fix error angle is estimated.Shown in the fix error angle of Fourier formalism such as formula (4), u is satellite
Ascending node argument.
F (u)=x0+x1sin(u)+x2cos(u)+x3sin(u)+x4cos(u)+...xnsin(u)+xn+1cos(u) (4)
Wherein, n is natural number.
Using nonlinear least square fitting method, the coefficient x for meeting below equation is found,
Min is is minimized oeprator, and then can obtain the Fourier coefficient (x in fundamental equation (4)0, x1, x2, x3,
x4...) estimated value.
Compared with prior art, the present invention has the advantages that：
(1) present invention is simple and effective, does not need to calculate the posture and attitude angular velocity of spacecraft；
(2) present invention can opposite fix error angle be estimated between star sensor in the case of dynamical variable is unknown
Meter；
(3) in actual task implementation procedure, the method ratio that the present invention provides needs to obtain in real time by Kalman filter
The method of estimation more practicability and effectiveness of spacecraft attitude is taken, it is emphasized that relative to traditional method of estimation, the present invention is right
The insensitivity of long-term loss of data so that the Data Detection on star is simpler with editing.
Description of the drawings
Upon reading the detailed description of non-limiting embodiments with reference to the following drawings, other feature of the invention,
Objects and advantages will become more apparent upon：
Fig. 1 is opposite installation error in-orbit identification method flow diagram between star sensor unit of the present invention.
Specific embodiment
It elaborates below to the implementation of the present invention：The present embodiment carries out in fact lower based on the technical solution of the present invention
It applies, gives detailed embodiment and specific operating process.It should be pointed out that those of ordinary skill in the art are come
It says, without departing from the inventive concept of the premise, various modifications and improvements can be made, these belong to the protection of the present invention
Range.
This implementation provides a kind of opposite installation error in-orbit identification method between star sensor unit, exists with respect to installation error
Rail discrimination method is as follows：
1st, data acquisition in the quick opposite fix error angle observational equation of star
Wherein, θiAnd θjThe respectively fix error angle of star sensor i and star sensor j, ξI, kAnd ξJ, kRespectively star is sensitive
Device i and star sensor j are in tkThe measurement noise vector at moment, Δ zIj, kFor observed quantity zIj, kObservation noise.Measure zIj, kIt obtains
Method is as follows：
[zIj, k]lmFor matrix [zIj, k] in l rows, m row element, matrix [zIj, k] computational methods see formula (3)
I is unit matrix,It is sensor i in moment tkObtained ontology attitude matrixEstimated value, i.e., full appearance
The measurement attitude matrix of state sensor iWith its initial installation matrix SoiThe numerical value of multiplication.
It is sensor j in moment tkObtained ontology attitude matrixThe transposed matrix of estimated value, i.e., full posture are quick
The measurement attitude matrix of sensor jWith its initial installation matrix SojMatrix after multiplication takes transposition.
Following observational equation can be built according to above-mentioned formula
Zk=Hkψ+ΔZk
Wherein Zk=zIj, k, Hk=I, Δ Zk~N (0, Pk)T, PkComputational methods such as formula under.
Pk=D (Δ zij)=D (ξi-ξj)
=D (ξi)+D(ξj)-2Cov(ξi, ξj).
=PI, k+PJ, k
Wherein PI, kIt is star sensor i in tkThe mean square deviation at moment, PJ, kIt is star sensor j in tkThe mean square deviation at moment, D ()
For variance oeprator, Cov () is covariance oeprator.
2nd, two star sensors opposite installation error angular estimation in certain period
Maximum likelihood equations is as follows to be estimated to quantity of state ψ using Maximum Likelihood Estimation
(Ψ) is minimized ψ, obtains following equation
Wherein
Quantity of state estimated value is as follows
3rd, the fitting of whole rail time-varying fix error angle mathematical model
The substantially square of star sensor time-varying fix error angle in the entire orbital period is represented in the form of the Fourier space
Journey, as shown in formula (4),
F (u)=x0+x1sin(u)+x2cos(u)+x3sin(2u)+x4cos(2u)+... (4)
Then using nonlinear least square fitting method, the coefficient x for meeting below equation is found,
It can obtain the Fourier coefficient (x in fundamental equation (4)0, x1, x2, x3, x4...) estimated value.
Opposite installation error in-orbit identification method between star sensor unit provided in this embodiment, including：Star sensor
It measures and the mathematical statistical model of installation error is established；Star sensor is established with respect to fix error angle observational equation；Installation error
Optimal estimation algorithm in angle designs, the design of fix error angle model nonlinear approximating method.It solves since different star sensors are in-orbit
Mounting arrangement changes, and causes to obtain the skimble-scamble problem of platform attitude angle measurement data by different star sensors, ensures different
Star sensor switches in use, the indifference transition of attitude information.The present embodiment is simple and effective, does not need to calculate the posture of spacecraft
And attitude angular velocity.Star sensor can be estimated with respect to fix error angle in the case of dynamical variable is unknown.In reality
During the tasks carrying of border, the present embodiment is more practical than needing the real-time fix error angle method of estimation for obtaining spacecraft attitude
Effectively, relative to traditional method of estimation, the present embodiment causes the insensitivity of long-term loss of data the data on star to examine
It surveys easier with estimation.
Specific embodiments of the present invention are described above.It is to be appreciated that the invention is not limited in above-mentioned
Particular implementation, those skilled in the art can make various deformations or amendments within the scope of the claims, this not shadow
Ring the substantive content of the present invention.
Claims (5)
1. a kind of opposite installation error in-orbit identification method between star sensor unit, which is characterized in that include the following steps：
Step S1 establishes the mathematical statistical model of star sensor measurement and installation error；
Step S2 establishes the quick opposite fix error angle observational equation of star；
Step S3 designs opposite fix error angle optimal estimation algorithm；
Step S4 designs fix error angle model nonlinear fitting algorithm；
The step S1, specifically includes following sub-step：
Step S101, it is as follows that installation error defines method：
In formula,It demarcates to obtain i-th of star sensor installation matrix initial value, M for groundiFor i-th of star sensor installation error
Matrix, θiFor i-th of star sensor installation error matrix corner vector, I is unit matrix, O ([θi]2) represent higher-order shear deformation；
I is positive integer；
Step S102, the related definition that star sensor measures are as follows：
It is i-th of star sensor in tkThe not corrected ontology attitude matrix that moment obtains, AkFor the true posture of spacecraft
Matrix, the relationship of the two are as follows：
It is i-th of star sensor in moment tkAttitude matrixEstimated value,It is i-th of star sensor in moment tk's
Measurement error matrix, ()TRepresent transposition oeprator, ξI, kIt is i-th star sensor in moment tkMeasurement noise vector, can
To obtain following relational expression：
2. opposite installation error in-orbit identification method between star sensor unit according to claim 1, which is characterized in that institute
It states in step S2, establishes star sensor in moment tkIt is as follows with respect to fix error angle observational equation：
Zk=Hkψ+ΔZk
Wherein, ZkFor observed quantity, HkFor observing matrix, ψ is quantity of state, Δ ZkFor observation noise,
Hk=I, Δ Zk~N (0, Pk)T, i.e. Δ ZkIt is 0 to obey mean value, variance PkGaussian Profile；
ZkMiddle component zIj, kObservational equation it is as follows, wherein i=1, j=2 ..., nk：
Wherein, θiAnd θjThe respectively fix error angle of i-th of star sensor and j-th of star sensor, (θj-θi) it is j-th of star
Sensor is with respect to the fix error angle of i-th of star sensor, quantity of state ψ as to be estimated；ξI, kAnd ξJ, kRespectively i-th of star
Sensor and j-th of star sensor are in tkThe measurement noise vector at moment, Δ zIj, kFor observed quantity zIj, kObservation noise；It measures
Measure zIj, kAcquisition methods it is as follows：
[zIj, k]lmFor matrix [zIj, k] in l rows, m row element, matrix [zIj, k] computational methods such as formula (3)
It is i-th of star sensor in moment tkObtained ontology attitude matrixEstimated value,Exist for j-th of sensor
Moment tkObtained ontology attitude matrixThe transposed matrix of estimated value.
3. opposite installation error in-orbit identification method between star sensor unit according to claim 1, which is characterized in that institute
It states in step S3, quantity of state ψ is estimated using optimal estimation algorithm.
4. opposite installation error in-orbit identification method between star sensor unit according to claim 3, which is characterized in that institute
It states in step S3, the least square estimation method, Maximum Likelihood Estimation can be used in optimal estimation algorithm.
5. opposite installation error in-orbit identification method between star sensor unit according to claim 1, which is characterized in that institute
It states in step S4, approximating method uses nonlinear least-square approximating method, to the fix error angle of the form of Fourier space
Mathematical model F (u) is estimated, obtains coefficient xo, x1, x2, x3, x4... estimated value, u be satellite ascending node argument；Installation misses
Declinate mathematical model F (u) expression formulas are：
F (u)=x0+x1sin(u)+x2cos(u)+x3sin(u)+x4cos(u)+...xnsin(u)+xn+1cos(u) (4)
Wherein, n is natural number.
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| CN106742083B (en) * | 2016-11-09 | 2019-01-08 | 上海卫星工程研究所 | A kind of free attachment device of face default value unloading based on in-orbit flexible release |
| CN106940196A (en) * | 2017-03-30 | 2017-07-11 | 上海航天控制技术研究所 | A kind of star sensor installs thermal distortion correction method |
| CN107228683B (en) * | 2017-06-27 | 2020-04-10 | 上海航天控制技术研究所 | Slow-variation error real-time on-orbit correction method among multiple star sensors |
| CN108195403B (en) * | 2017-12-28 | 2020-05-22 | 中国人民解放军国防科技大学 | Method and device for constructing star sensor on-orbit attitude measurement data comprehensive error model |
| CN108827320B (en) * | 2018-06-08 | 2021-08-17 | 西安电子科技大学 | Star sensor system convenient for on-orbit replacement and on-orbit replacement method |
| CN113091753B (en) * | 2021-03-02 | 2022-08-12 | 上海卫星工程研究所 | Satellite attitude guidance method and system for satellite sensitive view field protection |
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