CN105160125B - A kind of simulating analysis of star sensor quaternary number - Google Patents
A kind of simulating analysis of star sensor quaternary number Download PDFInfo
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Abstract
A kind of simulating analysis of star sensor quaternary number, (1) establishes satellite using emulation tool, and the preliminary orbit radical of satellite is arranged；(2) orbital position of the orbital tracking and satellite of satellite in the limiting time period under J2000 inertial coodinate systems is obtained；(3) roll angle at moonscope moment is calculated；(4) drift angle of satellite in the limiting time period is calculated；(5) attitude matrix of the star sensor measuring coordinate system relative to J2000 inertial coodinate systems in the limiting time period is calculated；(6) star sensor quaternary number is calculated.The present invention carries out simulation analysis independent of excessive hypothesis, the consideration a variety of in-orbit task gesture modes of satellite, using the method for numerical computations to star sensor quaternary number, solves the high accuracy analysis validation problem of star sensor attitude measure function and performance.In addition, the method for the present invention is alternatively arranged as a kind of interpretation method of star sensor attitude measure data, solve the problems, such as that the attitude measurement data of real-time dynamic change can not accurate interpretation.
Description
Technical field
The present invention relates to it is a kind of using simulation analysis obtain satellite star sensor quaternary number method, belong to the attitude of satellite with
Orbits controlling field.
Background technology
In recent years, with the rapid development of satellite technology, the requirement to its positioning accuracy is also higher and higher, is defended as guarantee
The technical research of the attitude measurement sensor of star lofty stance precision and lofty stance stability is also more and more urgent.Star sensor is to defend
Important measuring part in star attitude control system, and current widely applied optical attitude sensor；Its ether is aerial
Reference source of the fixed star as attitude measurement, direction of the output sensor optical axis in inertial reference system.Star sensor has posture
It determines precision height, without movable member, high reliability, is applicable to various track applications.
Currently, use of the Optical remote satellite to star sensor attitude measure data, is applied not only to determine posture over the ground, more
Important role is inserted into camera image using star sensor quaternary number as image auxiliary data, under be transmitted to user for really
Camera inertial attitude is determined, to which high-precision determines ground target.Attitude data of the terrestrial user to star sensor and gyro to measure
Carry out Kalman filter, it is ensured that possess degree of precision low frequency star sensor data and high frequency gyro data fusion after short
There is higher precision in time interval, the attitude data after fusion will be greatly improved into determining for image for framing in this way
Position precision.
With the raising that Optical remote satellite imaging resolution and image quality require, it is original to carry out accurate star sensor
Measurement data analysis verification has been increasingly becoming the overall necessary work of remote sensing satellite.Star sensor is that satellite is very crucial
The image quality for being directly related to remote sensing satellite, image position accuracy, camera imaging space are directed toward by component, function and performance
The stars such as precision integrated index realization.
Invention content
The technology of the present invention solves the problems, such as：Overcome the deficiencies of the prior art and provide that a kind of computational accuracy is higher, does not depend on
Assume in excessive, consider a variety of in-orbit task gesture modes of satellite, using numerical computations method to satellite normally over the ground posture,
The method that the star sensor quaternary number of roll attitude maneuver model carries out simulation analysis.
Technical solution of the invention is：A kind of simulating analysis of star sensor quaternary number, steps are as follows：
(1) satellite is established using emulation tool, the preliminary orbit radical of satellite is set；
(2) the preliminary orbit radical for the satellite and setting established according to step (1) obtains satellite in the limiting time period
Orbital position R under J2000 inertial coodinate systems of orbital tracking and satellitesat(t)；
(3) the satellite orbital position R obtained by step (2)sat(t) the ground target point the earth longitude and latitude observed with needs
Spend [Lond,Latd], calculate the roll angle at moonscope moment；
(4) it is defended in the limiting time period obtained by the moonscope moment roll angle of step (3) calculating, step (2)
Star orbital tracking calculates the drift angle of satellite in the limiting time period；
(5) it is defended in the limiting time period that moonscope moment roll angle, the step (4) calculated according to step (3) calculates
Orbit elements of satellite in the limiting time period obtained by star drift angle, step (2) calculates star sensor in the limiting time period and surveys
Measure attitude matrix of the coordinate system relative to J2000 inertial coodinate systems；
(6) star sensor measuring coordinate system sits relative to J2000 inertia in the limiting time period calculated according to step (5)
The attitude matrix of system is marked, star sensor quaternary number of the star sensor measuring coordinate system relative to J2000 inertial coodinate systems is calculated.
Steps are as follows for step (3) specific implementation：
(3.1) according to the earth longitude and latitude [Lon of ground target pointd,Latd], calculate ground target in the limiting time period
Position R of the point under J2000 inertial coodinate systemsif(t)；
(3.2) according to Rif(t) with orbital position R of the satellite under J2000 inertial coodinate systemssat(t), satellite is obtained to exist
The vector R of ground target point is directed toward under J2000 inertial coodinate systemsf(t)；
(3.3) by vector Rf(t) satellite orbit coordinate system is transformed by J2000 inertial coodinate systems, obtains satellite orbit seat
Vector R under mark systemo(t)；
(3.4) according to vector Ro(t) roll angle at moonscope moment is obtained.
The attitude matrix that satellite body coordinate system is transformed by satellite orbit coordinate system used in the step (5), root
It rotates and determines according to the Euler around satellite body reference axis, the corresponding Eulerian angles of attitude matrix are related with rotation order, rotate order
It is identical as sequence is turned used in satellite control system.
Order is moved by ZXY shaft rotations to obtain, then satellite body coordinate system is transformed by satellite orbit coordinate system when Eulerian angles
Attitude matrix Abo,(Z-X-Y)It is as follows：
Wherein,θ is satellite roll angle, pitch angle；Ψ is the satellite drift angle that step (4) calculates.
Order is moved by XZY shaft rotations to obtain, then satellite body coordinate system is transformed by satellite orbit coordinate system when Eulerian angles
Attitude matrix Abo,(X-Z-Y)It is as follows：
Wherein,θ is satellite roll angle, pitch angle；Ψ is the satellite drift angle that step (4) calculates.
Compared with the prior art, the present invention has the advantages of：
(1) star sensor quaternary number simulating analysis proposed by the present invention, computational accuracy is higher, independent of excessive vacation
If, fully consider a variety of in-orbit task gesture modes of satellite, it is quickly and easily normally right to satellite using the method for numerical computations
The star sensor quaternary number progress simulation analysis of ground posture, roll attitude maneuver model, efficiently solves the survey of star sensor posture
Measure the high accuracy analysis validation problem of function and performance.It is not only able in design of satellites and factory testing stage to star sensor
Function and performance are verified, additionally it is possible to which the source after satellier injection for verifying star epigraph influence factor is conducive to satellite
Totally from the quantitative in-orbit image quality of assurance satellite of the angle of system, supplemented for the ground test verification of satellite imagery link
One important means；
(2) attitude matrix of the satellite body coordinate system of the present invention relative to orbital coordinate system, at present practices well
It assumes that satellite body coordinate system and orbital coordinate system overlap and (assumes that the corresponding three axis Eulerian angles of satellite of attitude matrix are all
Zero)；Or set yaw angle to 0 (not considering that drift angle influences), and drift angle causes along rail and wears the picture shifting of rail direction, causes
Image is fuzzy, reduces image quality, is an important factor for influencing linear array CCD camera push-scanning image performance；Or by three axis Europe of satellite
Draw angle (can be obtained by satellite telemetering data, the method is limited to acquisition time and the acquiring way of data) as known quantity,
And normally only consider that 3-1-2 turns sequence (normally posture over the ground).Optical remote satellite is since viewing field of camera angle is smaller, in-orbit frequent need
Attitude maneuver is carried out, is observing visual field with the wide range for wearing rail direction along rail to obtain, it is therefore necessary to consider that satellite is a variety of
In-orbit task gesture mode；Optical remote satellite attitude control accuracy is more demanding, thus simulation analysis calculate when need compared with
For accurate attitude of satellite data, it is assumed that the way that satellite body coordinate system overlapped with orbital coordinate system or do not considered drift angle will
Simulation calculation precision can be substantially reduced；
(3) the method for the present invention can be used as a kind of interpretation method of star sensor attitude measure data.Currently used data
Interpretation method is interpretation of transfiniting, and by setting the variation range of data, bound comparison is carried out to collected data.And star is quick
With the attitude of satellite and orbital position real-time dynamic change occurs for sensor attitude data, and it is accurately fixed to be carried out according only to data area
Justice and interpretation.Be usually taken in actual test application judge the trend curve of quaternary number whether smoothly, the method without sharp cutting edge of a knife or a sword, and defend
When carrying out Large Angle Attitude Maneuver, a degree of transition, this qualitative interpretation will necessarily occur star for star sensor quaternary number
Method can not carry out interpretation, cause to fail to judge, probability of miscarriage of justice it is larger.The method of the present invention solves the star sensor of real-time dynamic change
Attitude measurement data can not accurate interpretation the problem of, can find in time, orientation problem, increase data interpretation work it is accurate
Property and validity, improve satellite failure early warning diagnosis capability.
Description of the drawings
Fig. 1 is the work flow diagram of the method for the present invention；
Fig. 2 is the quaternion algebra for the normal posture over the ground that the method for the present invention calculates according to bias contribution schematic diagram；
Fig. 3 is the quaternion algebra for the roll attitude maneuver model that the method for the present invention calculates according to bias contribution schematic diagram.
Specific implementation mode
The coordinate system used is needed to include herein：J2000 inertial coodinate systems, orbital coordinate system, satellite body coordinate system, star
Sensor measuring coordinate system, WGS-84 coordinate systems.The above coordinate system is defined separately below.
J2000 inertial coodinate systems
J2000 inertial coodinate systems OiXiYiZi, this coordinate system is the coordinate system of an inertial space, this coordinate system is with the earth's core
For origin Oi, XiAxis forward direction is directed toward UTC Universal Time Coordinated 1 day 12 January in 2000:The average first point of Aries direction of the earth measured when 00,
ZiAxis forward direction is directed toward the earth in UTC Universal Time Coordinated 1 day 12 January in 2000:The average axis of rotation the North measured when 00, YiAxis and Xi
Axis, ZiAxis is vertical, XiAxis, YiAxis, ZiAxis forms right-handed coordinate system.
Orbital coordinate system
Orbital coordinate system OoXoYoZo, origin OoThe centroid position when satellite is in-orbit, ZoAxis is directed toward the earth's core, X by barycenteroAxis exists
In orbit plane and ZoAxis is vertical and is directed toward satellite velocities direction, YoAxis and XoAxis, ZoAxis constitutes right hand rectangular coordinate system and and rail
The normal parallel of road plane；This coordinate system is rotation in space.
Satellite body coordinate system
Satellite body coordinate system ObXbYbZb, origin ObCenter positioned at satellite-rocket docking face, XbIt overlaps, is directed toward with the satellite longitudinal axis
Satellite y direction, in the same direction with heading under satellite flight state, ZbAxis is directed toward the earth's core, Y under satellite flight statebAxis
With XbAxis, ZbAxis constitutes right-handed coordinate system, and (the satellite longitudinal axis is defined as on celestial body, crosses satellite-rocket docking face center, is detached perpendicular to the satellite and the rocket
The axis that stellar interior is positive direction is directed toward in face).
Star sensor measuring coordinate system
Star sensor measuring coordinate system OsXsYsZs, origin OsPositioned at the center of star sensor ccd array, ZsAxis is along optical axis side
To XsAxis is perpendicular to optical axis and consistent with the direction of CCD rows scanning in CCD fronts, YsAxis and XsAxis, ZsIt is straight that axis constitutes the right hand
Angular coordinate system.
WGS-84 coordinate systems
WGS-84 coordinate systems OfXfYfZf, origin OfFor earth centroid, the Z of the earth's core rectangular coordinate system in spacefAxis is directed toward BIH
Direction agreement earth pole (CTP) that (international time) 1984.0 defines, XfZero meridian plane and CTP of axis direction BIH1984.0 is red
The intersection point in road, YfAxis and ZfAxis, XfAxis is vertically formed right-handed coordinate system.
The present invention is described in further detail below in conjunction with the accompanying drawings, as shown in Figure 1, the step of this simulating analysis
It is as follows：
(1) satellite is established using emulation tool, the preliminary orbit radical of satellite is set.
This step uses STK as emulation tool.STK softwares are opened, satellite is created, the preliminary orbit root of satellite is set
Number, including epoch time, semi-major axis, eccentricity, orbit inclination angle, right ascension of ascending node, argument of perigee, true anomaly, selection
Deduction model of the HPOP models as satellite orbit.
(2) the preliminary orbit radical for the satellite and setting established according to step (1) uses the REPORT work(of STK softwares
Can, using Δ t as emulation cycle (Δ t=1 seconds), obtain orbital tracking (semi-major axis a, the eccentricity of satellite in the limiting time period
E, right ascension of ascending node Ω, orbit inclination angle i, argument of perigee ω, true anomaly f) and rail of the satellite under J2000 inertial coodinate systems
Road position Rsat(t)：
Wherein, t indicates that UTC time, subscript " U " represent satellite.
Limiting time period choosing method herein：Using STK softwares, by the earth longitude and latitude [Lon of ground target pointd,
Latd] in corresponding position earth station's (earth station's height, which can simplify, is set as 0) is established, it is obtained using the ACCESS functions of STK softwares
Access time section of the satellite to the earth station, Visual simulation analysis is taken to need to increased or decrease the length of the time interval.
(3) the satellite orbital position R obtained by step (2)sat(t) the ground target point the earth longitude and latitude observed with needs
Spend [Lond,Latd], calculate the roll angle at moonscope moment.
The orbital position R of satellite under known J2000 inertial coodinate systemssat(t), the earth longitude and latitude of ground target point
[Lond,Latd], calculate the roll angle at moonscope moment.First according to the earth longitude and latitude of ground target point, when calculating restriction
Between position R of the ground target point under J2000 inertial coodinate systems in the periodif(t), then according to Rif(t) and satellite orbital position
Rsat(t), the vector R that satellite is directed toward ground target point is obtainedf(t), it then by the vector median filters to satellite orbit coordinate system, obtains
The roll angle at moonscope moment.It is as follows：
A. by ground target point the earth longitude and latitude [Lond,Latd] it is converted into the earth's core longitude and latitude [Lonc,Latc], it calculates public
Formula is：
Lonc=Lond Latc=tan-1[(1-f′)2tan Latd]
Wherein, f '=1/298.257223563 indicates the compression of the Earth.
B. target point the earth's core is calculated away from R：
Wherein, Re=6378137 meters.
C. appearance of the J2000 inertial coodinate systems relative to WGS-84 coordinate systems in the limiting time period is calculated according to UTC time
State matrix Aif(t), computational methods are published in National Defense Industry Press《Spacecraft orbit is theoretical》Have in (Liu Linzhu, 2000)
Detailed description.
D. position R of the ground target point under J2000 inertial coodinate systems in the limiting time period is calculatedif(t)：
Wherein, subscript " S " represents ground target point；Ay(α)、Az(α) indicates the basis element change square rotated around y, z-axis respectively
Battle array：
E. the vector R of satellite direction ground target point under J2000 inertial coodinate systems in the limiting time period is calculatedf(t)：
Wherein, subscript " U " represents satellite.
F. by vector Rf(t) satellite orbit coordinate system is transformed by J2000 inertial coodinate systems, obtains satellite orbit coordinate system
Under vector Ro(t)：
Wherein, AoiIndicate attitude matrix of the orbital coordinate system relative to J2000 inertial coodinate systems, it can be according to orbit parameter meter
It calculates：Right ascension of ascending node Ω, orbit inclination angle i, argument of perigee ω, true anomaly f, then orbital coordinate system is relative to J2000 inertia
The attitude matrix A of coordinate systemoiIt can write：
Wherein, u is satellite argument, there is u=ω+f.
G. roll angle roll (t), the pitch angle pitch (t) of satellite in the limiting time period are calculated：
H. t at the time of pitch angle pitch (t) is 0 in the limiting time periodObservationCorresponding roll angle roll (tObservation), i.e.,
For the roll angle at moonscope moment.Satellite is limited to the observation of target point the attitude maneuver range of satellite, only works as satellite
The attitude angle at moment is observed within the scope of the attitude maneuver of satellite, satellite could correctly execute attitude maneuver and observed object.It defends
In the in-orbit motion process of star, it is 0 to be changed by positive maximum value to the observation pitch angle of fixed target, then becomes negative maximum from 0
Value, wherein must have satellite pitch angle be 0 at the time of point.In moment point tObservation, satellite roll angle must satisfy condition：
|roll(tObservation)|≤rollmax
Wherein, rollmaxIndicate the maximum roll angle determined by attitude of satellite maneuvering range.
(4) orbit elements of satellite obtained by the moonscope moment roll angle of step (3) calculating, step (2) calculates
The drift angle of satellite in the limiting time period.
A. normally posture (sets roll angle to satellite over the groundPitching angle theta=0 °), drift angle ΨpIt can write：
Wherein, ωeExpression rotational-angular velocity of the earth (Unit be °/s)；ωnIndicate track angle speed
Degree；I indicates orbit inclination angle；A indicates semi-major axis；E indicates eccentricity；ω indicates argument of perigee；F indicates true anomaly；U is indicated
Satellite argument has u=ω+f；μ indicates Gravitational coefficient of the Earth (μ=398610)；P indicates semi-focal chord of satellite orbit；R indicates satellite the earth's core
Away from.
B. satellite roll attitude maneuver model, if roll angle is(the roll attitude maneuvering range for being limited to satellite), partially
Flow angle ΨpIt can write：
Wherein, β indicates geocentric angle；R indicate target point the earth's core away from；υrIndicate the radial component of satellite absolute velocity；Remaining
Symbol definition is same as above.
(5) it is defended in the limiting time period that moonscope moment roll angle, the step (4) calculated according to step (3) calculates
Orbit elements of satellite in the limiting time period obtained by star drift angle, step (2) calculates star sensor in the limiting time period and surveys
Measure attitude matrix of the coordinate system relative to J2000 inertial coodinate systems.
Attitude matrix A of the star sensor measuring coordinate system relative to J2000 inertial coodinate systemssiCalculation formula is as follows：
Asi=Asb*Abo*Aoi
a.AsbIndicate attitude matrix of the star sensor measuring coordinate system relative to satellite body coordinate system, i.e. star sensor
Matrix is installed.It is three axis and three axis of satellite body coordinate system to provide star sensor measuring coordinate by satellite configuration layout designs personnel
Angle, to after angle remainder string i.e. obtain the installation matrix of star sensor.
b.AboIndicate attitude matrix of the satellite body coordinate system relative to orbital coordinate system, it can be by satellite body reference axis of having mercy on
Euler rotation provide.Under normal circumstances, the corresponding Eulerian angles of attitude matrix are related with rotation order, turn sequence used in calculating herein
Identical as sequence is turned used in satellite control system (sequence should be turned used in satellite control system：Normally posture is generally turned using 3-1-2 over the ground
Sequence, roll attitude maneuver model generally turn sequence using 1-3-2).
If three axis Eulerian angles of satellite are roll angle(roll (the t calculated by step (3)Observation)), pitching angle theta (setting θ
=0), the yaw angle Ψ (Ψ calculated by step (4)p).If Eulerian angles move order by ZXY shaft rotations and obtain (i.e. 3-1-2 turns sequence),
Attitude matrix A of the satellite body coordinate system relative to orbital coordinate systembo,(Z-X-Y)It can write：
If Eulerian angles are moved order by XZY shaft rotations and obtained (i.e. 1-3-2 turns sequence), satellite body coordinate system is sat relative to track
Mark the attitude matrix A of systembo,(X-Z-Y)It can write：
c.AoiDefine f definition of same step (3).
(6) star sensor measuring coordinate system sits relative to J2000 inertia in the limiting time period calculated according to step (5)
Mark the attitude matrix A of systemsi, calculate star sensor quaternary number of the star sensor measuring coordinate system relative to J2000 inertial coodinate systems
[q0,q1,q2,q3](q0For scalar).
If attitude matrix A of the star sensor measuring coordinate system relative to J2000 inertial coodinate systemssiForm is as follows：
Star sensor quaternary number is found out using following formula：
Embodiment
The method of the present invention is verified using certain sun synchronization circular orbit satellite in-orbit data.The satellite was in 2013 4
On the moon 26 12：13：04.505 successful launch, the satellite and the rocket detach after satellite flight 764.451s.The measurement track at satellier injection moment
Parameter is as shown in table 1.
1 satellier injection moment of table orbit parameter
| Parameter name | Parameters at injection (J2000 inertial coodinate system wink roots) |
| Moment epoch | 2013-4-2612:25:(49.4050 Beijing time) |
| Semi-major axis (m) | 7025821.6501 |
| Eccentricity | 0.0011386731 |
| Orbit inclination angle (°) | 98.0436723 |
| Right ascension of ascending node (°) | 191.7772358 |
| Argument of perigee (°) | 192.9939337 |
| Mean anomaly (°) | 335.6950378 |
The attitude of satellite maneuvering range is designed as rotating direction ± 35 °, no pitching maneuverability.By the method for the present invention
Each step, you can high-precision star sensor quaternion algebra evidence, concrete outcome such as Fig. 2, Fig. 3 institute are obtained by simulation analysis
Show.
The normal selection of posture over the ground satellite afternoon 21 on June 17th, 2013:29:34--21:40:13 data when passing by, it is distant
4525422 seconds when surveying star -- 4526061 seconds (when referring to accumulative star after the satellite and the rocket detach when telemetering star herein).Roll attitude maneuver model
Choose the satellite morning 10 on June 17th, 2013:29:18--10:40:06 data when passing by, 4485806 seconds when telemetering star --
4486454 seconds (when referring to accumulative star after the satellite and the rocket detach when telemetering star herein), satellite is 14.3275 ° in the motor-driven angle of rotating direction.
It can be seen that the star sensor quaternary number and satellite telemetering data calculated using the method for the present invention from Fig. 2, Fig. 3
Deviation is respectively less than 0.0003, and computational accuracy is higher.
The content that description in the present invention is not described in detail belongs to the known technology of those skilled in the art.
Claims (5)
1. a kind of simulating analysis of star sensor quaternary number, it is characterised in that steps are as follows：
(1) satellite is established using emulation tool, the preliminary orbit radical of satellite is set；
(2) the preliminary orbit radical for the satellite and setting established according to step (1) obtains the rail of satellite in the limiting time period
The orbital position R of road radical and satellite under J2000 inertial coodinate systemssat(t)；
(3) the satellite orbital position R obtained by step (2)sat(t) the ground target point the earth longitude and latitude observed with needs
[Lond,Latd], calculate the roll angle at moonscope moment；
(4) satellite rail in the limiting time period obtained by the moonscope moment roll angle of step (3) calculating, step (2)
Road radical calculates the drift angle of satellite in the limiting time period；
(5) satellite is inclined in the limiting time period that moonscope moment roll angle, the step (4) calculated according to step (3) calculates
Orbit elements of satellite in the limiting time period obtained by angle, step (2) is flowed, star sensor in the limiting time period is calculated and measures seat
Attitude matrix A of the mark system relative to J2000 inertial coodinate systemssi：
Asi=Asb*Abo*Aoi
AsbIndicate attitude matrix of the star sensor measuring coordinate system relative to satellite body coordinate system；AboIndicate satellite body coordinate
It is the attitude matrix relative to orbital coordinate system；AoiIndicate attitude matrix of the orbital coordinate system relative to J2000 inertial coodinate systems；
(6) according to star sensor measuring coordinate system in the limiting time period of step (5) calculating relative to J2000 inertial coodinate systems
Attitude matrix, calculate star sensor quaternary number of the star sensor measuring coordinate system relative to J2000 inertial coodinate systems.
2. a kind of simulating analysis of star sensor quaternary number according to claim 1, it is characterised in that：The step
(3) steps are as follows for specific implementation：
(3.1) according to the earth longitude and latitude [Lon of ground target pointd,Latd], it calculates ground target point in the limiting time period and exists
Position R under J2000 inertial coodinate systemsif(t)；
(3.2) according to Rif(t) with orbital position R of the satellite under J2000 inertial coodinate systemssat(t), it is used in J2000 to obtain satellite
Property coordinate system under be directed toward ground target point vector Rf(t)；
(3.3) by vector Rf(t) satellite orbit coordinate system is transformed by J2000 inertial coodinate systems, obtained under satellite orbit coordinate system
Vector Ro(t)；
(3.4) according to vector Ro(t) roll angle at moonscope moment is obtained.
3. a kind of simulating analysis of star sensor quaternary number according to claim 1, it is characterised in that：The step
(5) that is used in is transformed into the attitude matrix of satellite body coordinate system by satellite orbit coordinate system, according to around satellite body coordinate
The Euler of axis, which rotates, to be determined, the corresponding Eulerian angles of attitude matrix are related with rotation order, rotation order and satellite control system institute
It is identical with sequence is turned.
4. a kind of simulating analysis of star sensor quaternary number according to claim 1, it is characterised in that：Work as Eulerian angles
Order is moved by ZXY shaft rotations to obtain, then the attitude matrix A of satellite body coordinate system is transformed by satellite orbit coordinate systembo,(Z-X-Y)
It is as follows：
Wherein,θ is satellite roll angle, pitch angle；Ψ is the satellite drift angle that step (4) calculates.
5. a kind of simulating analysis of star sensor quaternary number according to claim 1, it is characterised in that：Work as Eulerian angles
Order is moved by XZY shaft rotations to obtain, then the attitude matrix A of satellite body coordinate system is transformed by satellite orbit coordinate systembo,(X-Z-Y)
It is as follows：
Wherein,θ is satellite roll angle, pitch angle；Ψ is the satellite drift angle that step (4) calculates.
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