CN103604428A - Star sensor positioning method based on high-precision horizon reference - Google Patents

Abstract

The invention discloses a star sensor positioning method based on high-precision horizon reference. The star sensor positioning method comprises the following steps: firstly, acquiring an output Ci<s> of a CCD (charge coupled device) start sensor; combining the star sensor and a strapdown inertial navigation system, correcting an attitude of the strapdown inertial navigation system and compensating a mounting error of the star sensor to obtain high-precision horizon reference information; acquiring the high-precision horizon reference information supplied by the combined navigation system, namely acquiring a rolling angle and a pitching angle of a moving carrier to obtain an attitude conversion matrix Cb<N> from a carrier system to a quasi geographical coordinate system. Compared with the prior art, the star sensor positioning method has the advantages that the inertial navigation system is combined with the star sensor, and the attitude error of the inertial navigation system is corrected through filtration, so that the horizon reference information, on which the positioning of the star sensor depends, can be effectively improved; meanwhile, due to the determination of various error sources, the positioning precision of the star sensor is greatly improved.

Description

Star sensor localization method based on high level of accuracy benchmark

Technical field

The present invention relates to a kind of localization method of star sensor, relate in particular to a kind of star sensor localization method based on high level of accuracy benchmark.

Background technology

Star sensor is usingd indestructible natural celestial body as its navigation beacon, and the image that star sensor is photographed successively carries out importance in star map recognition, the extraction of celestial body barycenter, star pattern matching and attitude algorithm.Star sensor, without any prior imformation, just can independently be exported the attitude information in star sensor relative inertness space, its have accumulation in time of error, independent, passive, be not subject to the advantages such as human factor restriction and interference.The location of star sensor depends critically upon horizontal reference, so the done with high accuracy of horizontal reference is the important prerequisite that ensures its positioning precision.

Yet in application, conventionally utilize inertial navigation system that carrier horizontal reference information is provided, the shortcoming that inertial navigation system error accumulates in time declines star sensor positioning precision.

Summary of the invention

Object of the present invention provides a kind of star sensor localization method based on high level of accuracy benchmark with regard to being in order to address the above problem.

The present invention is achieved through the following technical solutions above-mentioned purpose:

A star sensor localization method based on high level of accuracy benchmark, comprises the following steps:

(1) gather the output of CCD star sensor

it is the attitude transition matrix of star sensor coordinate system relative inertness system;

(2) by the combination of star sensor and strapdown inertial navitation system (SINS), revise the attitude of strapdown inertial navitation system (SINS) and compensate the alignment error of star sensor, obtain the horizontal reference information of degree of precision;

(3) gather the high level of accuracy reference information that combinations thereof system provides, collect roll angle and the pitch angle of motion carrier, obtain the attitude transition matrix that carrier is tied to accurate geographic coordinate system

(4) according to Given information, solve i system with respect to the transition matrix between terrestrial coordinate system e system

(5) pass through step (1) to the given information of step (4), resolve and obtain accurate location matrix

calculate carrier positions information.

The output that the present invention gathers CCD star sensor is that coordinate system and the inertial coordinates system of CCD star sensor is the attitude information between i system

i system and boats and ships carrier coordinate system are the transition matrix between b system

be expressed as:

$C_i^b=C_s^b{C}_i^s$

Wherein:

for CCD star sensor coordinate system is the transition matrix between s system and b system,

when installing, star sensor determines, $C_s^b=\left[\begin{array}{ccc}1& -\mathrm{\&delta;}{A}_z& \mathrm{\&delta;}{A}_y\\ \mathrm{\&delta;}{A}_z& 1& -\mathrm{\&delta;}{A}_x\\ -\mathrm{\&delta;}{A}_y& \mathrm{\&delta;}{\mathrm{<figure-callout\; id="1"\; label="A\; x"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">A</figure-callout>}}_x& 1\end{array}\right],$ δ A wherein _x, δ A _y, δ A _zfor star sensor is along carrier coordinate system x, y, the alignment error angle in three directions of z axle; Celestial coordinate system is rotated to image space coordinate system and overlapped, and three rotations by following order represent:

$x_i{y}_i{z}_i\stackrel{z_i}{\&RightArrow;}{x}_1{y}_1{\mathrm{<figure-callout\; id="1"\; label="z"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">z</figure-callout>}}_1\stackrel{x_1}{\&RightArrow;}{x}_2{y}_2{\mathrm{<figure-callout\; id="2"\; label="z"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">z</figure-callout>}}_2\stackrel{y_2}{\&RightArrow;}{x}_s{y}_s{z}_s$

Corner is respectively

rotation matrix A is expressed as:

The coordinate of star sensor optical axis under celestial coordinate system is (α ₀, δ ₀), definition θ=90 °+δ ₀,

φ＝k ₀。α wherein ₀, δ ₀the right ascension and the declination that represent respectively optical axis, k ₀represent star sensor imaging surface Y _bthe angle of axle and pole axis and star sensor main shaft formed plane;

Star sensor output be expressed as:

$C_i^s=\left[\begin{array}{ccc}\sin {\mathrm{\&alpha;}}_0\cos k_0-\cos {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\sin k_0& -\cos {\mathrm{\&alpha;}}_0\cos k_0-\sin {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\sin {\mathrm{<figure-callout\; id="0"\; label="k"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">k</figure-callout>}}_0& \cos {\mathrm{\&delta;}}_0\sin k_0\\ -\sin {\mathrm{\&alpha;}}_0\sin k_0-\cos {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\mathrm{<figure-callout\; id="0"\; label="cos\; k"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">cos</figure-callout>}{k}_{}{}_0& \cos {\mathrm{\&alpha;}}_0\sin k_0-\sin {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\mathrm{<figure-callout\; id="0"\; label="cos\; k"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">cos</figure-callout>}{k}_{}{}_0& \cos {\mathrm{\&delta;}}_0\cos k_0\\ -\cos {\mathrm{\&alpha;}}_0\cos {\mathrm{\&delta;}}_0& -\sin {\mathrm{\&alpha;}}_0\cos {\mathrm{\&delta;}}_0& -\sin {\mathrm{\&delta;}}_0\end{array}\right]$

The attitude transition matrix of carrier coordinate system relative inertness coordinate system

The present invention is by star sensor and strapdown inertial navitation system (SINS) combination, revise the attitude of strapdown inertial navitation system (SINS), and compensate the alignment error of star sensor, obtain the attitude information of degree of precision, in Kalman filtering system, using INS errors as integrated navigation system state, specifically comprise SINS attitude error angle φ _e, φ _n, φ _u, velocity error δ V _e, δ V _n, site error δ λ, gyro Random Constant Drift ε _bx, ε _by, ε _bz, accelerometer constant error

the alignment error δ A of star sensor _x, δ A _y, δ A _z, so the state vector X of integrated navigation system is:

State equation is described as:

X(t)＝F(t)X(t)+G(t)W(t)

Wherein: F (t) is system state matrix, G (t) is system noise driving battle array, and W (t) is system white noise, here W (t)=[w _gx, w _gy, w _gz, w _ax, w _ay, w _az] ^t, w wherein _gx, w _gy, w _gzfor gyro white noise, w _ax, w _ay, w _azfor accelerometer white noise, inertial navigation system and star sensor all can be exported attitude of carrier angle information, the corresponding measurement information subtracting each other as integrated navigation system of attitude of carrier angle information of therefore inertial navigation system being exported with star sensor,

Z＝[φ _I-φ _S，θ _I-θ _S，γ _I-γ _S]

＝[δφ _I-δφ _S，δθ _I-δθ _S，δγ _I-δγ _S]

Bonding state vector X, can be listed as the measurement equation of writing integrated navigation system and be: Z=HX+V

Wherein: H is measurement matrix,

for star sensor is measured white noise, roll angle and the pitch angle output of calculating after filtering acquisition degree of precision are respectively θ, γ.

The present invention gathers the high level of accuracy reference information that combined system provides, and collects roll angle and the pitch angle of motion carrier, obtains the attitude transition matrix that carrier is tied to accurate geographic coordinate system twice rotation by following order represents, be respectively-γ of corner ,-θ;

Here by x _ny _nz _nbe referred to as N system; ?

be expressed as:

$C_b^N=C_1^{\mathrm{<figure-callout\; id="1"\; label="N\; C\; b"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">N</figure-callout>}}{C}_b^{}{}_1^{}$

$\begin{array}{c}=\left[\begin{array}{ccc}1& 0& 0\\ 0& \cos \mathrm{\&theta;}& -\sin \mathrm{\&theta;}\\ 0& \sin \mathrm{\&theta;}& \cos \mathrm{\&theta;}\end{array}\right]\left[\begin{array}{ccc}\cos \mathrm{\&gamma;}& 0& \sin \mathrm{\&gamma;}\\ 0& 1& 0\\ -\sin \mathrm{\&gamma;}& 0& \cos \mathrm{\&gamma;}\end{array}\right]\\ =\left[\begin{array}{ccc}\cos \mathrm{\&gamma;}& 0& \sin \mathrm{\&gamma;}\\ \sin \mathrm{\&theta;}\sin \mathrm{\&gamma;}& \cos \mathrm{\&theta;}& -\sin \mathrm{\&theta;}\cos \mathrm{\&gamma;}\\ -\cos \mathrm{\&theta;}\sin \mathrm{\&gamma;}& \sin \mathrm{\&theta;}& \cos \mathrm{\&theta;}\cos \mathrm{\&gamma;}\end{array}\right]\end{array}.$

Further, according to Given information, solve i system with respect to the transition matrix of terrestrial coordinate system e system

expression formula is as follows:

$C_e^i=\left[\begin{array}{ccc}\cos (A_j+w_{\mathrm{ie}}\&CenterDot;t)& -\sin (A_j+w_{\mathrm{ie}}\&CenterDot;t)& 0\\ \sin (A_j+w_{\mathrm{ie}}\&CenterDot;t)& \cos (A_j+w_{\mathrm{ie}}\&CenterDot;t)& 0\\ 0& 0& 1\end{array}\right]$

A _jinitial position and the angle between the first point of Aries, w _iefor rotational-angular velocity of the earth, t is the concrete time that universal time system provides, and is Given information.

Further, by resolving, obtain accurate location matrix

and then calculate carrier positions information, detailed process is as follows:

Known location matrix

expression formula is as follows:

basis

with

matrix last column correspondent equal calculates carrier positions

Beneficial effect of the present invention is:

The present invention is a kind of star sensor localization method based on high level of accuracy benchmark, compared with prior art, the present invention is by combining inertial navigation system and star sensor, by the attitude error of filtering and calibration inertial navigation system, effectively improve the horizontal reference information that star sensor location relies on, all kinds of error sources are determined simultaneously, have greatly improved the positioning precision of star sensor.

Accompanying drawing explanation

Fig. 1 utilizes star sensor installation error that Matlab emulation obtains to location precision figure;

Fig. 2 is the method positioning error curve map that utilizes Matlab emulation to obtain;

Fig. 3 is the steps flow chart block diagram of invention.

Embodiment

Below in conjunction with accompanying drawing, the invention will be further described:

As shown in Figure 1 to Figure 3: a kind of star sensor localization method based on high level of accuracy benchmark, comprises the following steps:

(1) gather the output of CCD star sensor

it is the attitude transition matrix of star sensor coordinate system relative inertness system;

(2) by the combination of star sensor and strapdown inertial navitation system (SINS), revise the attitude of strapdown inertial navitation system (SINS) and compensate the alignment error of star sensor, obtain the horizontal reference information of degree of precision;

(3) gather the high level of accuracy reference information that combinations thereof system provides, collect roll angle and the pitch angle of motion carrier, obtain the attitude transition matrix that carrier is tied to accurate geographic coordinate system

(4) according to Given information, solve i system with respect to the transition matrix between terrestrial coordinate system e system

(5) pass through step (1) to the given information of step (4), resolve and obtain accurate location matrix

calculate carrier positions information.

The output that the present invention gathers CCD star sensor is that coordinate system and the inertial coordinates system of CCD star sensor is the attitude information between i system

i system and boats and ships carrier coordinate system are the transition matrix between b system be expressed as:

$C_i^b=C_s^b{C}_i^s$

Wherein:

for CCD star sensor coordinate system is the transition matrix between s system and b system,

when installing, star sensor determines, $C_s^b=\left[\begin{array}{ccc}1& -\mathrm{\&delta;}{A}_z& \mathrm{\&delta;}{A}_y\\ \mathrm{\&delta;}{A}_z& 1& -\mathrm{\&delta;}{A}_x\\ -\mathrm{\&delta;}{A}_y& \mathrm{\&delta;}{\mathrm{<figure-callout\; id="1"\; label="A\; x"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">A</figure-callout>}}_x& 1\end{array}\right],$ δ A wherein _x, δ A _y, δ A _zfor star sensor is along carrier coordinate system x, y, the alignment error angle in three directions of z axle; Celestial coordinate system is rotated to image space coordinate system and overlapped, and three rotations by following order represent:

$x_i{y}_i{z}_i\stackrel{z_i}{\&RightArrow;}{x}_1{y}_1{\mathrm{<figure-callout\; id="1"\; label="z"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">z</figure-callout>}}_1\stackrel{x_1}{\&RightArrow;}{x}_2{y}_2{\mathrm{<figure-callout\; id="2"\; label="z"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">z</figure-callout>}}_2\stackrel{y_2}{\&RightArrow;}{x}_s{y}_s{z}_s$

Corner is respectively

rotation matrix A is expressed as:

The coordinate of star sensor optical axis under celestial coordinate system is (α ₀, δ ₀), definition θ=90 °+δ ₀,

φ＝k ₀。α wherein ₀, δ ₀the right ascension and the declination that represent respectively optical axis, k ₀represent star sensor imaging surface Y _bthe angle of axle and pole axis and star sensor main shaft formed plane.

Star sensor output

be expressed as:

$C_i^s=\left[\begin{array}{ccc}\sin {\mathrm{\&alpha;}}_0\cos k_0-\cos {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\sin k_0& -\cos {\mathrm{\&alpha;}}_0\cos k_0-\sin {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\mathrm{<figure-callout\; id="0"\; label="sin\; k"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">sin</figure-callout>}{k}_{}{}_0& \cos {\mathrm{\&delta;}}_0\sin k_0\\ -\sin {\mathrm{\&alpha;}}_0\sin k_0-\cos {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\mathrm{<figure-callout\; id="0"\; label="cos\; k"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">cos</figure-callout>}{k}_{}{}_0& \cos {\mathrm{\&alpha;}}_0\sin k_0-\sin {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\mathrm{<figure-callout\; id="0"\; label="cos\; k"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">cos</figure-callout>}{k}_{}{}_0& \cos {\mathrm{\&delta;}}_0\cos k_0\\ -\cos {\mathrm{\&alpha;}}_0\cos {\mathrm{\&delta;}}_0& -\sin {\mathrm{\&alpha;}}_0\cos {\mathrm{\&delta;}}_0& -\sin {\mathrm{\&delta;}}_0\end{array}\right]$

The attitude transition matrix of carrier coordinate system relative inertness coordinate system

The present invention, by star sensor and strapdown inertial navitation system (SINS) combination, revises the attitude of strapdown inertial navitation system (SINS), and compensates the alignment error of star sensor, obtains the attitude information of degree of precision.In Kalman filtering system, using INS errors as integrated navigation system state, specifically comprise SINS attitude error angle φ _e, φ _n, φ _u, velocity error δ V _e, δ V _n, site error δ λ,

gyro Random Constant Drift ε _bx, ε _by, ε _bz, accelerometer constant error

the alignment error δ A of star sensor _x, δ A _y, δ A _z, so the state vector X of integrated navigation system is:

State equation is described as:

X(t)＝F(t)X(t)+G(t)W(t)

Wherein: F (t) is system state matrix, G (t) is system noise driving battle array, and W (t) is system white noise, here W (t)=[w _gx, w _gy, w _gz, w _ax, w _ay, w _az] ^t, w wherein _gx, w _gy, w _gzfor gyro white noise, w _ax, w _ay, w _azfor accelerometer white noise, inertial navigation system and star sensor all can be exported attitude of carrier angle information, the corresponding measurement information subtracting each other as integrated navigation system of attitude of carrier angle information of therefore inertial navigation system being exported with star sensor,

Z＝[φ _I-φ _S，θ _I-θ _S，γ _I-γ _S]

＝[δφ _I-δφ _S，δθ _I-δθ _S，δγ _I-δγ _S]

Bonding state vector X, can be listed as the measurement equation of writing integrated navigation system and be: Z=HX+V

Wherein: H is measurement matrix,

for star sensor is measured white noise.Roll angle and the pitch angle output of calculating after filtering acquisition degree of precision are respectively θ, γ.

The present invention gathers the high level of accuracy reference information that combined system provides, and collects roll angle and the pitch angle of motion carrier, obtains the attitude transition matrix that carrier is tied to accurate geographic coordinate system twice rotation by following order represents, be respectively-γ of corner ,-θ;

Here by x _ny _nz _nbe referred to as N system; ?

be expressed as:

$\begin{array}{c}{C}_b^N=C_1^{\mathrm{<figure-callout\; id="1"\; label="N\; C\; b"\; filenames="HSA0000097868350000011.png"\; state="\{\{state\}\}">N</figure-callout>}}{C}_b^{}{}_1^{}\\ =\left[\begin{array}{ccc}1& 0& 0\\ 0& \cos \mathrm{\&theta;}& -\sin \mathrm{\&theta;}\\ 0& \sin \mathrm{\&theta;}& \cos \mathrm{\&theta;}\end{array}\right]\left[\begin{array}{ccc}\cos \mathrm{\&gamma;}& 0& \sin \mathrm{\&gamma;}\\ 0& 1& 0\\ -\sin \mathrm{\&gamma;}& 0& \cos \mathrm{\&gamma;}\end{array}\right]\\ =\left[\begin{array}{ccc}\cos \mathrm{\&gamma;}& 0& \sin \mathrm{\&gamma;}\\ \sin \mathrm{\&theta;}\sin \mathrm{\&gamma;}& \cos \mathrm{\&theta;}& -\sin \mathrm{\&theta;}\cos \mathrm{\&gamma;}\\ -\cos \mathrm{\&theta;}\sin \mathrm{\&gamma;}& \sin \mathrm{\&theta;}& \cos \mathrm{\&theta;}\cos \mathrm{\&gamma;}\end{array}\right]\end{array}$

Further, according to Given information, solve i system with respect to the transition matrix of terrestrial coordinate system e system

expression formula is as follows:

$C_e^i=\left[\begin{array}{ccc}\cos (A_j+w_{\mathrm{ie}}\&CenterDot;t)& -\sin (A_j+w_{\mathrm{ie}}\&CenterDot;t)& 0\\ \sin (A_j+w_{\mathrm{ie}}\&CenterDot;t)& \cos (A_j+w_{\mathrm{ie}}\&CenterDot;t)& 0\\ 0& 0& 1\end{array}\right]$

A _jinitial position and the angle between the first point of Aries, w _iefor rotational-angular velocity of the earth, t is the concrete time that universal time system provides, and is Given information.

Further, by resolving, obtain accurate location matrix

and then calculate carrier positions information, detailed process is as follows:

Known location matrix

expression formula is as follows:

basis

with

matrix last column correspondent equal calculates carrier positions

The present invention first gathers the attitude transition matrix of the star sensor coordinate system relative inertness system of star sensor output, then by the method for filtering, the attitude of inertial navigation system output is proofreaied and correct, and compensate the alignment error of star sensor, the high level of accuracy benchmark obtaining is offered to star sensor and position.In filtering, select the error of inertial navigation system as integrated navigation system state, observed quantity using the difference of the attitude angle of inertial navigation system and star sensor as combined system, by state equation X (t)=F (t) X (t)+G (t) W (t) and observation equation Z=HX+V, carry out filtering, and then proofread and correct the horizontal attitude information of inertial navigation system.

As a second aspect of the present invention, star sensor positioning error is only that roll angle and the pitch angle of motion carrier is relevant with horizontal reference information, by carrier coordinate system being done to twice reverse rotation, make the Z axis of motion carrier and the sky of geographic coordinate system to overlapping, obtain the attitude transition matrix of accurate geographic coordinate system relative inertness coordinate system, and then by resolving the positional information that obtains motion carrier.Whole irrelevant with course information in resolving process, only need the precision that improves transverse and longitudinal cradle angle can effectively improve star sensor positioning precision, need not proofread and correct processing to course information.

Matlab emulation

The very high and error of the measuring accuracy of star sensor is accumulation in time, but star sensor is difficult to require to be installed on carrier according to accurate orientation conventionally, and alignment error will have a strong impact on the navigation accuracy of star sensor, therefore need to compensate the alignment error of star sensor, so need the impact of emulation star sensor installation error on positioning precision, and the positioning error after filtering compensation alignment error and level of corrections benchmark is passed through in emulation.

Under following simulated conditions, the alignment error of star sensor is carried out to emulation experiment to location impact:

Carrier initial position: 45.7796 ° of north latitude, 126.6705 ° of east longitudes;

The true attitude angle of carrier: Ψ=0 °, θ=5 °, γ=5 °; Wherein: Ψ, θ, γ represents respectively course angle, pitch angle and roll angle; The horizontal reference of supposing star sensor acquisition is error free.

Equatorial radius: R _e=6378393.0m; By the available earth surface acceleration of gravity of universal gravitation: g ₀=9.78049; Rotational-angular velocity of the earth (radian per second): 7.2921158e-5; Constant: π=3.1415926; Star sensor is along carrier coordinate system x, y, and the alignment error angle in three directions of z axle is respectively 5,3,3; Simulation time: t=1 hour; Sample frequency: Hn=0.1; Emulation obtains star sensor installation error on the impact of positioning precision as shown in Figure 1: longitude error is 7.4426 jiaos minutes; Latitude error is 0.5356 jiao minute.

Under following simulated conditions, utilize the method compensation star sensor installation error and proofread and correct the horizontal reference information that inertial navigation is exported, utilize the method to resolve position, this description is carried out to emulation experiment:

Strapdown inertial navitation system (SINS) remains static; Carrier initial position: 45.7796 ° of north latitude, 126.6705 ° of east longitudes; The true attitude angle of carrier: Ψ=0 °, θ=5 °, γ=5 °; Wherein: Ψ, θ, γ represents respectively course angle, pitch angle and roll angle; Equatorial radius: R _e=6378393.0m; By the available earth surface acceleration of gravity of universal gravitation: g ₀=9.78049; Rotational-angular velocity of the earth (radian per second): 7.2921158e-5; The X of star sensor, Y, the alignment error of Z axis is respectively 5 ', 3 ', 3 '; Constant: π=3.1415926; Simulation time: t=24 hour; Sample frequency: Hn=0.1; Utilize the described method of invention to obtain positioning error as shown in Figure 2: after star sensor installation error being compensated and proofreaied and correct the horizontal reference of inertial navigation output, the star sensor location longitude error of 24 hours is about 1 jiao minute, latitude error is approximately 0.5 jiao minute, and passing is in time becoming periodic swinging.

Claims (6)

1. the star sensor localization method based on high level of accuracy benchmark, is characterized in that, comprises the following steps:

(1) gather the output of CCD star sensor

it is the attitude transition matrix of star sensor coordinate system relative inertness system;

(2) by the combination of star sensor and strapdown inertial navitation system (SINS), revise the attitude of strapdown inertial navitation system (SINS) and compensate the alignment error of star sensor, obtain the horizontal reference information of degree of precision;

(3) gather the high level of accuracy reference information that combinations thereof system provides, collect roll angle and the pitch angle of motion carrier, obtain the attitude transition matrix that carrier is tied to accurate geographic coordinate system

(4) according to Given information, solve i system with respect to the transition matrix between terrestrial coordinate system e system

(5) pass through step (1) to the given information of step (4), resolve and obtain accurate location matrix

calculate carrier positions information.

2. the star sensor localization method based on high level of accuracy benchmark according to claim 1, is characterized in that: the output that gathers CCD star sensor is that coordinate system and the inertial coordinates system of CCD star sensor is the attitude information between i system

i system and boats and ships carrier coordinate system are the transition matrix between b system

be expressed as:

$C_i^b=C_s^b{C}_i^s$

Wherein:

for CCD star sensor coordinate system is the transition matrix between s system and b system,

when installing, star sensor determines, $C_s^b=\left[\begin{array}{ccc}1& -\mathrm{\&delta;}{A}_z& \mathrm{\&delta;}{A}_y\\ \mathrm{\&delta;}{A}_z& 1& -\mathrm{\&delta;}{A}_x\\ -\mathrm{\&delta;}{A}_y& \mathrm{\&delta;}{A}_x& 1\end{array}\right],$ δ A wherein _x, δ A _y, δ A _zfor star sensor is along carrier coordinate system x, y, the alignment error angle in three directions of z axle; Celestial coordinate system is rotated to image space coordinate system and overlapped, and three rotations by following order represent:

$x_i{y}_i{z}_i\stackrel{z_i}{\&RightArrow;}{x}_1{y}_1{z}_1\stackrel{x_1}{\&RightArrow;}{x}_2{y}_2{z}_2\stackrel{y_2}{\&RightArrow;}{x}_s{y}_s{z}_s$

Corner is respectively

rotation matrix A is expressed as:

The coordinate of star sensor optical axis under celestial coordinate system is (α ₀, δ ₀), definition θ=90 °+δ ₀,

φ＝k ₀。α wherein ₀, δ ₀the right ascension and the declination that represent respectively optical axis, k ₀represent star sensor imaging surface Y _bthe angle of axle and pole axis and star sensor main shaft formed plane;

Star sensor output be expressed as:

$C_i^s=\left[\begin{array}{ccc}\sin {\mathrm{\&alpha;}}_0\cos k_0-\cos {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\sin k_0& -\cos {\mathrm{\&alpha;}}_0\cos k_0-\sin {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\sin k_0& \cos {\mathrm{\&delta;}}_0\sin k_0\\ -\sin {\mathrm{\&alpha;}}_0\sin k_0-\cos {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\cos k_0& \cos {\mathrm{\&alpha;}}_0\sin k_0-\sin {\mathrm{\&alpha;}}_0\sin {\mathrm{\&delta;}}_0\cos k_0& \cos {\mathrm{\&delta;}}_0\cos k_0\\ -\cos {\mathrm{\&alpha;}}_0\cos {\mathrm{\&delta;}}_0& -\sin {\mathrm{\&alpha;}}_0\cos {\mathrm{\&delta;}}_0& -\sin {\mathrm{\&delta;}}_0\end{array}\right]$

The attitude transition matrix of carrier coordinate system relative inertness coordinate system

3. the star sensor localization method based on high level of accuracy benchmark according to claim 2, it is characterized in that: by star sensor and strapdown inertial navitation system (SINS) combination, revise the attitude of strapdown inertial navitation system (SINS), and compensate the alignment error of star sensor, obtain the attitude information of degree of precision, in Kalman filtering system, using INS errors as integrated navigation system state, specifically comprise SINS attitude error angle φ _e, φ _n, φ _u, velocity error δ V _e, δ V _n, site error δ λ, gyro Random Constant Drift ε _bx, ε _by, ε _bz, accelerometer constant error

the alignment error δ A of star sensor _x, δ A _y, δ A _z, so the state vector X of integrated navigation system is:

State equation is described as:

X(t)＝F(t)X(t)+G(t)W(t)

Wherein: F (t) is system state matrix, G (t) is system noise driving battle array, and W (t) is system white noise, here W (t)=[w _gx, w _gy, w _gz, w _ay, w _az] ^t, w wherein _gx, w _gy, w _gzfor gyro white noise, w _ax, w _ay, w _azfor accelerometer white noise, inertial navigation system and star sensor all can be exported attitude of carrier angle information, the corresponding measurement information subtracting each other as integrated navigation system of attitude of carrier angle information of therefore inertial navigation system being exported with star sensor,

Z＝[φ _I-φ _S，θ _I-θ _S，γ _I-γ _S]

＝[δφ _I-δφ _S，δφ _I-δφ _S，δγ _I-δγ _S]

Bonding state vector X, can be listed as the measurement equation of writing integrated navigation system and be: Z=HX+V

Wherein: H is measurement matrix, for star sensor is measured white noise, roll angle and the pitch angle output of calculating after filtering acquisition degree of precision are respectively θ, γ.

4. the star sensor localization method based on high level of accuracy benchmark according to claim 3, it is characterized in that: gather the high level of accuracy reference information that combined system provides, collect roll angle and the pitch angle of motion carrier, obtain the attitude transition matrix that carrier is tied to accurate geographic coordinate system

, by twice rotation of following order, represent be respectively-γ of corner ,-θ;

Here by x _ny _nz _nbe referred to as N system; ?

be expressed as:

$\begin{array}{c}{C}_b^N=C_1^N{C}_b^1\\ =\left[\begin{array}{ccc}1& 0& 0\\ 0& \cos \mathrm{\&theta;}& -\sin \mathrm{\&theta;}\\ 0& \sin \mathrm{\&theta;}& \cos \mathrm{\&theta;}\end{array}\right]\left[\begin{array}{ccc}\cos \mathrm{\&gamma;}& 0& \sin \mathrm{\&gamma;}\\ 0& 1& 0\\ -\sin \mathrm{\&gamma;}& 0& \cos \mathrm{\&gamma;}\end{array}\right]\\ =\left[\begin{array}{ccc}\cos \mathrm{\&gamma;}& 0& \sin \mathrm{\&gamma;}\\ \sin \mathrm{\&theta;}\sin \mathrm{\&gamma;}& \cos \mathrm{\&theta;}& -\sin \mathrm{\&theta;}\cos \mathrm{\&gamma;}\\ -\cos \mathrm{\&theta;}\sin \mathrm{\&gamma;}& \sin \mathrm{\&theta;}& \cos \mathrm{\&theta;}\cos \mathrm{\&gamma;}\end{array}\right]\end{array}.$

5. the star sensor localization method based on high level of accuracy benchmark according to claim 4, is characterized in that: according to Given information, solve i system with respect to the transition matrix of terrestrial coordinate system e system

expression formula is as follows:

$C_e^i=\left[\begin{array}{ccc}\cos (A_j+w_{\mathrm{ie}}\&CenterDot;t)& -\sin (A_j+w_{\mathrm{ie}}\&CenterDot;t)& 0\\ \sin (A_j+w_{\mathrm{ie}}\&CenterDot;t)& \cos (A_j+w_{\mathrm{ie}}\&CenterDot;t)& 0\\ 0& 0& 1\end{array}\right]$

A _jinitial position and the angle between the first point of Aries, w _iefor rotational-angular velocity of the earth, t is the concrete time that universal time system provides, and is Given information.

6. the star sensor localization method based on high level of accuracy benchmark according to claim 5, is characterized in that: by resolving, obtain accurate location matrix and then calculate carrier positions information, detailed process is as follows:

Known location matrix

expression formula is as follows:

basis

with

matrix last column correspondent equal calculates carrier positions

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