CN105160125A - Simulation analysis method for star sensor quaternion - Google Patents

Abstract

A simulation analysis method for a star sensor quaternion comprises: (1) establishing a satellite by using a simulation tool, and setting an initial orbital element of the satellite; (2) obtaining the orbital element of the satellite within a limited time cycle and an orbital position of the satellite in a J2,000 inertial coordinate system; (3) calculating a rolling angle of the satellite at the observation moment; (4) calculating a drift angle of the satellite within the limited time cycle; (5) calculating an attitude matrix of a star sensor measurement coordinate system for the J2,000 inertial coordinate system within the limited time cycle; and (6) calculating the star sensor quaternion. The simulation analysis method does not depend on too many assumptions, considers various in-orbit task attitude modes of the satellite, performs simulation analysis on the star sensor quaternion with a numerical value calculation method, and solves the problem in high-precision analysis and verification of star sensor attitude measurement function and performance. In addition, the method also can serve as an interpretation method for star sensor attitude measurement data, thereby solving the problem that the attitude measurement data dynamically changed in real time cannot be accurately interpreted.

Description

A kind of simulating analysis of star sensor hypercomplex number

Technical field

The present invention relates to a kind of method adopting simulation analysis to obtain Satellite sensor hypercomplex number, belong to the attitude of satellite and orbits controlling field.

Background technology

In recent years, along with developing rapidly of satellite technology, also more and more higher to the requirement of its positioning precision, the technical research as the attitude measurement sensor ensureing satellite lofty stance precision and lofty stance degree of stability is also more and more urgent.Star sensor is the important measurement component in satellite attitude control system, is also the optical attitude sensor of current widespread use; The aerial fixed star of its ether, as the reference source of attitude measurement, exports the sensing of sensor optical axis in inertial reference system.Star sensor have attitude determination accuracy high, without movable member, high reliability, be applicable to various track application.

At present, Optical remote satellite is to the use of star sensor attitude measure data, not only for determining attitude over the ground, prior effect is inserted in camera image using star sensor hypercomplex number as image auxiliary data, under pass to user for determining camera inertial attitude, thus high precision Area Objects definitely.The attitude data of terrestrial user to star sensor and gyro to measure carries out Kalman filter, can ensure to have after the low frequency star sensor data of degree of precision and high frequency gyro data merge in short time interval, have higher precision, like this attitude data after merging being used for framing will improve the positioning precision of image greatly.

Along with the raising that Optical remote satellite imaging resolution and image quality require, carry out accurate star sensor raw measurement data analysis verification, become necessity work that remote sensing satellite is overall gradually.Star sensor is the very crucial parts of satellite, and its function and performance will be directly connected to the realization of the ground integrated index of star such as image quality, image position accuracy, camera imaging spatial direction precision of remote sensing satellite.

Summary of the invention

Technology of the present invention is dealt with problems and is: overcome the deficiencies in the prior art, there is provided a kind of computational accuracy higher, do not rely on too much hypothesis, consider satellite multiple task in-orbit gesture mode, use the star sensor hypercomplex number of the method for numerical evaluation to satellite normal attitude, roll attitude maneuver model over the ground to carry out the method for simulation analysis.

Technical solution of the present invention is: a kind of simulating analysis of star sensor hypercomplex number, and step is as follows:

(1) use emulation tool to set up satellite, the preliminary orbit radical of satellite is set;

(2) satellite set up according to step (1) and the preliminary orbit radical of setting, obtain orbital tracking and the orbital position R of satellite under J2000 inertial coordinates system of satellite in the limiting time cycle _sat(t);

(3) according to the satellite orbital position R of step (2) gained _satthe terrain object point the earth longitude and latitude [Lon of (t) and needs observation _d, Lat _d], calculate the roll angle in moonscope moment;

(4) the moonscope moment roll angle calculated according to step (3), orbit elements of satellite in the limiting time cycle of step (2) gained, calculate the drift angle of satellite in the limiting time cycle;

(5) in the limiting time cycle that the moonscope moment roll angle calculated according to step (3), step (4) calculate satellite drift angle, step (2) gained the limiting time cycle in orbit elements of satellite, to calculate in the limiting time cycle star sensor surving coordinate system relative to the attitude matrix of J2000 inertial coordinates system;

(6) in the limiting time cycle calculated according to step (5), star sensor surving coordinate system is relative to the attitude matrix of J2000 inertial coordinates system, calculates the star sensor hypercomplex number of star sensor surving coordinate system relative to J2000 inertial coordinates system.

Described step (3) specific implementation step is as follows:

(3.1) the earth longitude and latitude [Lon of base area appearance punctuate _d, Lat _d], calculate the position R of terrain object point under J2000 inertial coordinates system in the limiting time cycle _if(t);

(3.2) according to R _if(t) and the orbital position R of satellite under J2000 inertial coordinates system _satt (), obtains the vector R that satellite points to terrain object point under J2000 inertial coordinates system _f(t);

(3.3) by vector R _ft () is transformed into satellite orbit coordinate system by J2000 inertial coordinates system, obtain the vector R under satellite orbit coordinate system _o(t);

(3.4) according to vector R _ot () obtains the roll angle in moonscope moment.

Use in described step (5) by the attitude matrix of satellite orbit ordinate transform to satellite body coordinate system, rotate according to the Euler around satellite body coordinate axis and determine, Eulerian angle corresponding to attitude matrix are relevant with rotation order, and rotating order and satellite control system, used to turn sequence identical.

Rotate order when Eulerian angle by ZXY axle to draw, then by the attitude matrix A of satellite orbit ordinate transform to satellite body coordinate system _{bo, (Z-X-Y)}as follows:

Wherein, θ is satellite roll angle, the angle of pitch; Ψ is the satellite drift angle that step (4) calculates.

Rotate order when Eulerian angle by XZY axle to draw, then by the attitude matrix A of satellite orbit ordinate transform to satellite body coordinate system _{bo, (X-Z-Y)}as follows:

Wherein, θ is satellite roll angle, the angle of pitch; Ψ is the satellite drift angle that step (4) calculates.

The present invention's advantage is compared with prior art:

(1) the star sensor hypercomplex number simulating analysis of the present invention's proposition, computational accuracy is higher, do not rely on too much hypothesis, take into full account satellite multiple task in-orbit gesture mode, use the method for numerical evaluation to carry out simulation analysis to the star sensor hypercomplex number of satellite normal attitude, roll attitude maneuver model over the ground quickly and easily, efficiently solve the high accuracy analysis validation problem of star sensor attitude measure function and performance.Not only can verify star sensor function and performance in design of satellites and factory testing stage, can also for verifying the source of star epigraph influence factor after satellier injection, be beneficial to satellite totally from the assurance satellite image quality in-orbit that the angle of system is quantitative, for the ground test checking of satellite imagery link supplements an important means;

(2) satellite body coordinate system of the present invention is relative to the attitude matrix of orbital coordinate system, and current practices well is that hypothesis satellite body coordinate system overlaps (namely supposing that the satellite three axle Eulerian angle that attitude matrix is corresponding are all zero) with orbital coordinate system; Or crab angle is set to 0 (namely not considering that drift angle affects), and drift angle causes along rail and wears rail direction picture and move, and causes image blurring, reducing image quality, is the key factor of effect string array CCD camera push-scanning image performance; Or using satellite three axle Eulerian angle as known quantity (obtain by satellite telemetering data, the method is limited to acquisition time and the acquiring way of data), and general only consider that 3-1-2 turns sequence (normally over the ground attitude).Optical remote satellite is less due to viewing field of camera angle, often needs in-orbit to carry out attitude maneuver, to obtain along rail and the visual field of observation in a big way of wearing rail direction, therefore must consider satellite multiple task in-orbit gesture mode; Optical remote satellite attitude control accuracy requires higher, therefore need the comparatively accurate attitude of satellite data when simulation analysis calculates, suppose that satellite body coordinate system overlaps with orbital coordinate system or do not consider that the way of drift angle will reduce simulation calculation precision greatly;

(3) the inventive method can be used as a kind of interpretation method of star sensor attitude measure data.Data interpretation method conventional is at present interpretation of transfiniting, and by the variation range of setting data, carries out bound compare the data collected.And star sensor attitude data, with the attitude of satellite and orbital position, real-time dynamic change occurs, only accurate definition and interpretation cannot be carried out according to data area.Usually take in actual Test Application to judge the trend curve of hypercomplex number whether level and smooth, without the method for sharp cutting edge of a knife or a sword, and satellite is when carrying out Large Angle Attitude Maneuver, transition to a certain degree will inevitably be there is in star sensor hypercomplex number, this qualitative interpretation method cannot carry out interpretation, cause fail to judge, probability of miscarriage of justice is larger.The inventive method solves the star sensor attitude measure data of real-time dynamic change cannot the problem of accurate interpretation, can Timeliness coverage, orientation problem, adds accuracy and the validity of data interpretation work, improves satellite failure early warning diagnosis capability.

Accompanying drawing explanation

Fig. 1 is the workflow diagram of the inventive method;

Fig. 2 is that the quaternion algebra of the normal attitude over the ground that the inventive method calculates is according to bias contribution schematic diagram;

Fig. 3 is that the quaternion algebra of the roll attitude maneuver model that the inventive method calculates is according to bias contribution schematic diagram.

Embodiment

The coordinate system used is needed to comprise herein: J2000 inertial coordinates system, orbital coordinate system, satellite body coordinate system, star sensor surving coordinate system, WGS-84 coordinate system.Define above coordinate system respectively below.

J2000 inertial coordinates system

J2000 inertial coordinates system O _ix _iy _iz _i, this coordinate is the coordinate system of an inertial space, and this coordinate system take the earth's core as initial point O _i, X _ithe direction in the average first point of Aries of the earth that axle forward measures when pointing to UTC Universal Time Coordinated 12:00 on the 1st January in 2000, Z _iaxle forward points to the average axis of rotation the North that the earth measures when UTC Universal Time Coordinated 12:00 on the 1st January in 2000, Y _iaxle and X _iaxle, Z _iaxle is vertical, X _iaxle, Y _iaxle, Z _iaxle forms right-handed coordinate system.

Orbital coordinate system

Orbital coordinate system O _ox _oy _oz _o, initial point O _osatellite in-orbit time centroid position, Z _oaxle points to the earth's core by barycenter, X _oaxle is in orbit plane and Z _oaxle vertically and point to satellite velocities direction, Y _oaxle and X _oaxle, Z _oaxle form right hand rectangular coordinate system and with the normal parallel of orbit plane; This coordinate system is rotate in space.

Satellite body coordinate system

Satellite body coordinate system O _bx _by _bz _b, initial point O _bbe positioned at the center in satellite-rocket docking face, X _boverlap with the satellite longitudinal axis, point to satellite y direction, under satellite flight state with heading in the same way, Z _baxle points to the earth's core under satellite flight state, Y _baxle and X _baxle, Z _baxle forms right-handed coordinate system (the satellite longitudinal axis is defined as on celestial body, crosses center, satellite-rocket docking face, perpendicular to satellite and the rocket parting plane, points to the axis that stellar interior is positive dirction).

Star sensor surving coordinate system

Star sensor surving coordinate system O _sx _sy _sz _s, initial point O _sbe positioned at the center of star sensor ccd array, Z _saxle along optical axis direction, X _saxle is with the direction of CCD line scanning consistent perpendicular to optical axis in CCD front, Y _saxle and X _saxle, Z _saxle forms right hand rectangular coordinate system.

WGS-84 coordinate system

WGS-84 coordinate system O _fx _fy _fz _f, initial point O _ffor earth centroid, the Z of its earth's core rectangular coordinate system in space _faxle points to agreement earth pole (CTP) direction of BIH (international time) 1984.0 definition, X _faxle points to zero meridian ellipse of BIH1984.0 and the intersection point in CTP equator, Y _faxle and Z _faxle, X _faxle vertically forms right-handed coordinate system.

Be described in further detail the present invention below in conjunction with accompanying drawing, as shown in Figure 1, the step of this simulating analysis is as follows:

(1) use emulation tool to set up satellite, the preliminary orbit radical of satellite is set.

This step uses STK as emulation tool.Open STK software, newly-built satellite, the preliminary orbit radical of satellite is set, comprises epoch time, semi-major axis, excentricity, orbit inclination, right ascension of ascending node, argument of perigee, true anomaly, select HPOP model as the deduction model of satellite orbit.

(2) satellite set up according to step (1) and the preliminary orbit radical of setting, use the REPORT function of STK software, with Δ t for emulation cycle (Δ t=1 second), obtain orbital tracking (semi-major axis a, eccentric ratio e, right ascension of ascending node Ω, orbit inclination i, argument of perigee ω, true anomaly f) and the orbital position R of satellite under J2000 inertial coordinates system of satellite in the limiting time cycle _sat(t):

$R_{sat}\left(t\right)=\left[\begin{array}{c}{x}_U\left(t\right)\\ y_U\left(t\right)\\ z_U\left(t\right)\end{array}\right]$

Wherein, t represents the UTC time, and subscript " U " represents satellite.

Limiting time cycle choosing method herein: use STK software, by the earth longitude and latitude [Lon of terrain object point _d, Lat _d] set up land station (land station height can simplify be set to 0) at correspondence position, use the ACCESS function of STK software to obtain satellite interval to the access time of this land station, Visual simulation analysis needs the length increasing or reduce this time interval.

(3) according to the satellite orbital position R of step (2) gained _satthe terrain object point the earth longitude and latitude [Lon of (t) and needs observation _d, Lat _d], calculate the roll angle in moonscope moment.

The orbital position R of satellite under known J2000 inertial coordinates system _satthe earth longitude and latitude [the Lon of (t), terrain object point _d, Lat _d], calculate the roll angle in moonscope moment.First the earth longitude and latitude of base area appearance punctuate, calculates the position R of terrain object point under J2000 inertial coordinates system in the limiting time cycle _ift (), then according to R _if(t) and satellite orbital position R _satt (), obtains the vector R that satellite points to terrain object point _f(t), then by this vector median filters to satellite orbit coordinate system, obtain the roll angle in moonscope moment.Concrete steps are as follows:

A. by terrain object point the earth longitude and latitude [Lon _d, Lat _d] be converted into the earth's core longitude and latitude [Lon _c, Lat _c], computing formula is:

Lon _c＝Lon _dLat _c＝tan ^-1[(1-f′) ²tanLat _d]

Wherein, f '=1/298.257223563, represents compression of the earth.

B. impact point the earth's core is calculated apart from R:

$R=R_e\frac{1-f^{\′}}{\sqrt{1-f^{\′}(2-f^{\′}){\cos }^2{\mathrm{Lat}}_c}}$

Wherein, R _e=6378137 meters.

C. according to J2000 inertial coordinates system in the UTC Time Calculation limiting time cycle relative to the attitude matrix A of WGS-84 coordinate system _ift (), has a detailed description in " spacecraft orbit theoretical " (Liu Linzhu, 2000) that computing method are published in National Defense Industry Press.

D. the position R of terrain object point under J2000 inertial coordinates system in the limiting time cycle is calculated _if(t):

Wherein, subscript " S " represents terrain object point; A _y(α), A _z(α) the basis element change matrix rotated around y, z-axis is represented respectively:

$A_y\left(\mathrm{\α}\right)=\left[\begin{array}{ccc}cos\mathrm{\α}& 0& -sin\mathrm{\α}\\ 0& 1& 0\\ sin\mathrm{\α}& 0& \cos \mathrm{\α}\end{array}\right]$ $A_z\left(\mathrm{\α}\right)=\left[\begin{array}{ccc}cos\mathrm{\α}& \sin \mathrm{\α}& 0\\ -sin\mathrm{\α}& \cos \mathrm{\α}& 0\\ 0& 0& 1\end{array}\right]$

E. the vector R of satellite sensing terrain object point under J2000 inertial coordinates system in the limiting time cycle is calculated _f(t):

$R_f\left(t\right)=R_{if}\left(t\right)-R_{sat}\left(t\right)=\left[\begin{array}{c}{x}_S\left(t\right)\\ y_S\left(t\right)\\ z_S\left(t\right)\end{array}\right]-\left[\begin{array}{c}{x}_U\left(t\right)\\ y_U\left(t\right)\\ z_U\left(t\right)\end{array}\right]$

Wherein, subscript " U " represents satellite.

F. by vector R _ft () is transformed into satellite orbit coordinate system by J2000 inertial coordinates system, obtain the vector R under satellite orbit coordinate system _o(t):

$R_o\left(t\right)=A_{oi}{R}_f\left(t\right)=A_{oi}\{\left[\begin{array}{c}{x}_S\left(t\right)\\ y_S\left(t\right)\\ z_S\left(t\right)\end{array}\right]-\left[\begin{array}{c}{x}_U\left(t\right)\\ y_U\left(t\right)\\ z_U\left(t\right)\end{array}\right]\}=\left[\begin{array}{c}{x}_{US}^o\left(t\right)\\ y_{US}^o\left(t\right)\\ z_{US}^o\left(t\right)\end{array}\right]$

Wherein, A _oirepresent that orbital coordinate system is relative to the attitude matrix of J2000 inertial coordinates system, can calculate according to orbit parameter: right ascension of ascending node Ω, orbit inclination i, argument of perigee ω, true anomaly f, then orbital coordinate system is relative to the attitude matrix A of J2000 inertial coordinates system _oican write:

$A_{oi}=\left[\begin{array}{ccc}0& 1& 0\\ 0& 0& -1\\ -1& 0& 0\end{array}\right]\left[\begin{array}{ccc}\cos u& \sin u& 0\\ -\sin u& \cos u& 0\\ 0& 0& 1\end{array}\right]\left[\begin{array}{ccc}1& 0& 0\\ 0& \cos i& \sin i\\ 0& -\sin i& \cos i\end{array}\right]\left[\begin{array}{ccc}cos\mathrm{\Ω}& \sin \mathrm{\Ω}& 0\\ -sin\mathrm{\Ω}& \cos \mathrm{\Ω}& 0\\ 0& 0& 1\end{array}\right]$

Wherein, u is satellite argument, has u=ω+f.

G. roll angle roll (t), angle of pitch pitch (t) of satellite in the limiting time cycle is calculated:

$roll\left(t\right)={\tan }^{-1}\[-\frac{y_{US}^o\left(t\right)}{z_{US}^o\left(t\right)}\]$ $pitch\left(t\right)={\sin }^{-1}\[\frac{x_{US}^o\left(t\right)}{r_{US}\left(t\right)}\]$

Wherein, $r_{US}\left(t\right)=\sqrt{{\[x_{US}^o\left(t\right)\]}^2+{\[y_{US}^o\left(t\right)\]}^2+{\[z_{US}^o\left(t\right)\]}^2}.$

H. in the limiting time cycle, angle of pitch pitch (t) is the moment t of 0 _observationcorresponding roll angle roll (t _observation), be the roll angle in moonscope moment.The observation of satellite to impact point is limited to the attitude maneuver scope of satellite, only has when the attitude angle in moonscope moment is within the scope of the attitude maneuver of satellite, and satellite correctly could perform attitude maneuver and observed object.Satellite, in-orbit in motion process, is changed to 0 to the observation angle of pitch of fixed target by positive maximal value, then becomes negative maximal value from 0, wherein must have the satellite angle of pitch be 0 moment point.At this moment point t _observation, satellite roll angle must satisfy condition:

| roll (t _observation) |≤roll _max

Wherein, roll _maxrepresent the maximum roll angle determined by attitude of satellite maneuvering range.

(4) the moonscope moment roll angle calculated according to step (3), the orbit elements of satellite of step (2) gained, calculate the drift angle of satellite in the limiting time cycle.

A. the normal attitude over the ground of satellite (namely sets roll angle pitching angle theta=0 °), drift angle Ψ _pcan write:

${\mathrm{\ψ}}_p={\tan }^{-1}\left(\frac{{\mathrm{cosusini\ω}}_e}{{\mathrm{cosi\ω}}_e-{\mathrm{\ω}}_n}\right)$

${\mathrm{\ω}}_n=\frac{\sqrt{\mathrm{\μ}p}}{r^2}$ p＝a(1-e ²) $r=\frac{p}{1+e\cos f}$

Wherein, ω _eexpression rotational-angular velocity of the earth ( unit is °/s); ω _nrepresent orbit angular velocity; I represents orbit inclination; A represents semi-major axis; E represents excentricity; ω represents argument of perigee; F represents true anomaly; U represents satellite argument, has u=ω+f; μ represents Gravitational coefficient of the Earth (μ=398610); P represents semi-focal chord of satellite orbit; R represents satellite the earth's core distance.

B. satellite roll attitude maneuver model, if roll angle is (being limited to the roll attitude maneuvering range of satellite), drift angle Ψ _pcan write:

${\mathrm{\υ}}_r=\sqrt{\frac{\mathrm{\μ}}{p}}e\sin f$

Wherein, β represents geocentric angle; R represents impact point the earth's core distance; υ _rrepresent the radial component of satellite absolute velocity; All the other symbol definitions are the same.

(5) in the limiting time cycle that the moonscope moment roll angle calculated according to step (3), step (4) calculate satellite drift angle, step (2) gained the limiting time cycle in orbit elements of satellite, to calculate in the limiting time cycle star sensor surving coordinate system relative to the attitude matrix of J2000 inertial coordinates system.

Star sensor surving coordinate system is relative to the attitude matrix A of J2000 inertial coordinates system _sicomputing formula is as follows:

A _si＝A _sb*A _bo*A _oi

A.A _sbrepresent the attitude matrix of star sensor surving coordinate system relative to satellite body coordinate system, i.e. the installation matrix of star sensor.The angle of star sensor surving coordinate system three axle and satellite body coordinate system three axle is provided, to the installation matrix namely obtaining star sensor after angle remainder string by satellite configuration topological design personnel.

B.A _borepresent the attitude matrix of satellite body coordinate system relative to orbital coordinate system, can be rotated by the Euler of rich satellite body coordinate axis and provide.Generally, the Eulerian angle that attitude matrix is corresponding are relevant with rotation order, calculate the sequence that turns used herein and should turn sequence identical (satellite control system is used turns sequence: normal attitude over the ground generally adopts 3-1-2 to turn sequence, and roll attitude maneuver model generally adopts 1-3-2 to turn sequence) with satellite control system is used.

If three axle Eulerian angle of satellite are roll angle (roll (the t calculated by step (3) _observation)), pitching angle theta (setting θ=0), the crab angle Ψ (Ψ calculated by step (4) _p).If Eulerian angle are rotated order by ZXY axle and drawn (namely 3-1-2 turns sequence), then satellite body coordinate system is relative to the attitude matrix A of orbital coordinate system _{bo, (Z-X-Y)}can write:

If Eulerian angle are rotated order by XZY axle and drawn (namely 1-3-2 turns sequence), then satellite body coordinate system is relative to the attitude matrix A of orbital coordinate system _{bo, (X-Z-Y)}can write:

C.A _oidefine the definition of same step (3) f item.

(6) in the limiting time cycle calculated according to step (5), star sensor surving coordinate system is relative to the attitude matrix A of J2000 inertial coordinates system _si, calculate the star sensor hypercomplex number [q of star sensor surving coordinate system relative to J2000 inertial coordinates system ₀, q ₁, q ₂, q ₃] (q ₀for scalar).

If star sensor surving coordinate system is relative to the attitude matrix A of J2000 inertial coordinates system _siform is as follows:

$A_{si}=\left[\begin{array}{ccc}{A}_{11}& A_{12}& A_{13}\\ A_{21}& A_{22}& A_{23}\\ A_{31}& A_{32}& A_{33}\end{array}\right]$

Following formula can be utilized to obtain star sensor hypercomplex number:

$q_0=\frac{\sqrt{1+A_{11}+A_{22}+A_{33}}}{2}$

Embodiment

Use certain sun synchronization circular orbit satellite in-orbit data the inventive method is verified.This satellite was in 12:13:04.505 successful launch on April 26 in 2013, and after satellite flight 764.451s, the satellite and the rocket are separated.The measurement orbit parameter in satellier injection moment is as shown in table 1.

Table 1 satellier injection moment orbit parameter

Parameter name	Parameters at injection (J2000 inertial coordinates system wink root)
		Moment epoch	2013-4-2612:25:49.4050 (Beijing time)
Semi-major axis (m)	7025821.6501
		Excentricity	0.0011386731
Orbit inclination (°)	98.0436723
		Right ascension of ascending node (°)	191.7772358
Argument of perigee (°)	192.9939337
		Mean anomaly (°)	335.6950378

This attitude of satellite maneuvering range is designed to rotating direction ± 35 °, without pitching maneuverability.Through each step of the inventive method, namely obtain high-precision star sensor quaternion algebra certificate by simulation analysis, concrete outcome as shown in Figure 2 and Figure 3.

Data when normal attitude is over the ground chosen satellite June 17, afternoon, 21:29:34--21:40:13 was passed by 2013, during remote measurement star 4525422 seconds--4526061 seconds (refer to during remote measurement star herein the satellite and the rocket be separated after accumulative star time).Data when roll attitude maneuver model is chosen satellite June 17, the morning, 10:29:18--10:40:06 was passed by 2013, during remote measurement star 4485806 seconds--4486454 seconds (refer to during remote measurement star herein the satellite and the rocket be separated after accumulative star time), satellite is 14.3275 ° in the motor-driven angle of rotating direction.

As can be seen from Fig. 2, Fig. 3, the star sensor hypercomplex number using the inventive method to calculate and the deviate of satellite telemetering data are all less than 0.0003, and computational accuracy is higher.

The content be not described in detail in instructions of the present invention belongs to the known technology of those skilled in the art.

Claims (5)

1. a simulating analysis for star sensor hypercomplex number, is characterized in that step is as follows:

(1) use emulation tool to set up satellite, the preliminary orbit radical of satellite is set;

(2) satellite set up according to step (1) and the preliminary orbit radical of setting, obtain orbital tracking and the orbital position R of satellite under J2000 inertial coordinates system of satellite in the limiting time cycle _sat(t);

(3) according to the satellite orbital position R of step (2) gained _satthe terrain object point the earth longitude and latitude [Lon of (t) and needs observation _d, Lat _d], calculate the roll angle in moonscope moment;

(4) the moonscope moment roll angle calculated according to step (3), orbit elements of satellite in the limiting time cycle of step (2) gained, calculate the drift angle of satellite in the limiting time cycle;

(5) in the limiting time cycle that the moonscope moment roll angle calculated according to step (3), step (4) calculate satellite drift angle, step (2) gained the limiting time cycle in orbit elements of satellite, to calculate in the limiting time cycle star sensor surving coordinate system relative to the attitude matrix of J2000 inertial coordinates system;

(6) in the limiting time cycle calculated according to step (5), star sensor surving coordinate system is relative to the attitude matrix of J2000 inertial coordinates system, calculates the star sensor hypercomplex number of star sensor surving coordinate system relative to J2000 inertial coordinates system.

2. the simulating analysis of a kind of star sensor hypercomplex number according to claim 1, is characterized in that: described step (3) specific implementation step is as follows:

(3.1) the earth longitude and latitude [Lon of base area appearance punctuate _d, Lat _d], calculate the position R of terrain object point under J2000 inertial coordinates system in the limiting time cycle _if(t);

(3.2) according to R _if(t) and the orbital position R of satellite under J2000 inertial coordinates system _satt (), obtains the vector R that satellite points to terrain object point under J2000 inertial coordinates system _f(t);

(3.3) by vector R _ft () is transformed into satellite orbit coordinate system by J2000 inertial coordinates system, obtain the vector R under satellite orbit coordinate system _o(t);

(3.4) according to vector R _ot () obtains the roll angle in moonscope moment.

3. the simulating analysis of a kind of star sensor hypercomplex number according to claim 1, it is characterized in that: use in described step (5) by the attitude matrix of satellite orbit ordinate transform to satellite body coordinate system, rotate according to the Euler around satellite body coordinate axis and determine, Eulerian angle corresponding to attitude matrix are relevant with rotation order, and rotating order and satellite control system, used to turn sequence identical.

4. the simulating analysis of a kind of star sensor hypercomplex number according to claim 1, is characterized in that: rotate order when Eulerian angle by ZXY axle and draw, then by the attitude matrix A of satellite orbit ordinate transform to satellite body coordinate system _{bo, (Z-X-Y)}as follows:

Wherein, θ is satellite roll angle, the angle of pitch; Ψ is the satellite drift angle that step (4) calculates.

5. the simulating analysis of a kind of star sensor hypercomplex number according to claim 1, is characterized in that: rotate order when Eulerian angle by XZY axle and draw, then by the attitude matrix A of satellite orbit ordinate transform to satellite body coordinate system _{bo, (X-Z-Y)}as follows:

Wherein, θ is satellite roll angle, the angle of pitch; Ψ is the satellite drift angle that step (4) calculates.