CN100585333C - The Benyue segment base is in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor - Google Patents

Abstract

The Benyue segment base is in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor, comprise the following steps: that the image that (1) is taken ultraviolet sensors cuts apart, judge whether have the earth or the moon in the captured image according to cutting apart back target geometrical feature, if the no earth or the moon, then do not possess the condition that realizes that autonomous Orbit is estimated, wait for the image that take next time; Otherwise, carry out the identification of the earth, the moon, judge that the earth and the moon whether in same image, if the two not in same image, then changes step (2), otherwise change step (3); (2) according to the inertia attitude data of star sensor output, half angle, moon vector or the earth vector of ultraviolet sensors output carries out track and estimates; (3) only carry out track and estimate, or carry out track according to the output valve of the measured value of ultraviolet sensors and star sensor and estimate according to the measured value of ultraviolet sensors.

Description

The Benyue segment base is in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor

Technical field

The present invention relates to the autonomous Orbit algorithm for estimating of a kind of this month section, belong to optical imagery attitude sensor application.

Background technology

Along with the quick progress of imaging detection device and processor technology, the spacecraft attitude sensor is developed to imaging type by the unit scan formula gradually, and ultraviolet sensors is exactly a kind of big view field imaging formula attitude sensor that is different from traditional horizon instrument.

Ultraviolet sensors possesses the ability that the Benyue section is obtained the earth, lunar map picture, the earth/lunar map that utilization is obtained goes out the earth's core/moon heart vector as information calculations and finishes distance estimations with respect to the earth's core/moon heart, can realize two function a and nominal track here relatively, estimate that orbit error b cooperates the quick output inertia of star attitude to realize the independent navigation test.

Although abroad the Smart1 such as Europe also proposes to utilize high precision visible light camera and inertia attitude to carry out at the rail navigation experiment in the Benyue section, but the specific implementation algorithm does not relate to, and how to realize being that those skilled in the art thirsts for the technical barrier that solves always.

The aerospace journal, 2006 the 3rd phases, " based on the lunar orbiter independent combined navigation of federal UKF algorithm " introduced the Earth-moon transfer orbit utilize infrared earth sensor output the earth's core vector, utilize a ultraviolet sensors output month heart vector, star sensor calculate the inertia attitude by three groups of information combination after filtering realizes independent navigation.Compare its difference with the technology of the present invention and not enough to be that in the different embodiments of application target infrared earth sensor belongs to the scan type precision lower and only be suitable for near-earth ball track, so algorithm is unsuitable for lunar trajectory in the literary composition; The track method of estimation that does not relate to ultraviolet sensors photographic images correlation process method in the literary composition and use ultraviolet sensors separately in addition.

The space science journal, 2003 the 2nd phases, " the astronomical autonomous orbit determination method for determining posture research of a kind of satellite " has introduced the earth satellite independent navigation algorithm based on ultraviolet sensors, sun sensor, echo altimeter.Compare its difference with the bright technology of we and be that with deficiency the application target track that method is not suitable for near-earth satellite in the identical text estimates, do not relate to the Flame Image Process of ultraviolet sensors, introduce the track method of estimation of using ultraviolet sensors separately and cooperating star sensor.

Aerospace Control, 2004 the 22nd the 3rd phases of volume, in " based on the spacecraft independent navigation of ultraviolet sensors " the autonomous navigation of satellite method based on ultraviolet sensors is studied, provided the filtering algorithm that autonomous Orbit is determined.Method is suitable for near-earth orbit and does not relate to moon target in the literary composition, ultraviolet image relevant treatment and determination methods are not arranged in the literary composition, the earth, the moon that also do not have in the literary composition in addition to use ultraviolet sensors to obtain separately carry out the introduction that track is estimated, are a kind of filtering method rather than simple optical means by kinetics equation the most hereinafter.

U.S. Patent number US5319969, title " Method for determining 3-axis spacecraftattitude ".Introduce a kind of three-axis attitude of ultraviolet spectral coverage attitude sensor that utilizes and determined method, wherein do not related to the track estimation problem.

Summary of the invention

Technology of the present invention is dealt with problems and is: overcome the deficiencies in the prior art, the autonomous Orbit method of estimation of a kind of Benyue segment base in ultraviolet sensors, star sensor is provided, this method can only be utilized ultraviolet sensors, and perhaps ultraviolet sensors cooperates star sensor to carry out the autonomous Orbit estimation.

Technical solution of the present invention is: the Benyue segment base comprises the following steps: in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor

(1) image that ultraviolet sensors is taken is cut apart, and judges whether have the earth or the moon in the captured image according to cutting apart back target geometrical feature, if the no earth or the moon then do not possess the condition that realizes that autonomous Orbit is estimated, waits for the image that take next time; Otherwise, carry out the identification of the earth, the moon, judge the whether judgement in same image of the earth and the moon, if the two not in same image, then changes step (2), otherwise change step (3);

(2) according to the inertia attitude data of star sensor output, half angle, moon vector or the earth vector of ultraviolet sensors output carries out track and estimates;

(3) only carry out track and estimate, or carry out track according to the output valve of the measured value of ultraviolet sensors and star sensor and estimate according to the measured value of ultraviolet sensors.

After track is estimated, also carrying out orbit error estimates, described orbit error estimates to adopt track value and ground predicted value from host computer relatively to obtain orbit error, or takes the photograph in the image position according to the earth under the predicted orbit, moon image space and actual institute and compare and carry out the orbit error estimation.

Judge whether have the earth or the moon in the captured image, determination methods is in the described step (1):

Whether the first step forms image patch in the image after judgement is cut apart according to gray threshold, if there is not image patch, the no earth or the moon stop follow-up work, otherwise the image patch size are analyzed in the key diagram picture, calculate image patch interior pixel number N um, judge image patch interior pixel number N um and number threshold value T _NSize, if Num＞=T _N, image patch may mark for the earth or the moon, if Num＜T _N, image patch is not the earth or moon target;

Second step, the image patch number M that marks in the statistics first step, if

M=0 does not have the earth or the moon in the image, stop follow-up work;

M=1 has the earth or the moon in the image, but is not to exist simultaneously;

M=2 exists the earth, the moon simultaneously in the image;

M＞2, image memory stops follow-up work in big noise;

Wherein, gray threshold selects ground to inject or the priori default value, generally greater than 2 times maximum background gray levels;

Number threshold value T _N, decide by nominal track, crescent or half moon.

During M=1, the recognition methods of the earth or the moon is as follows in described second step:

The first step is calculated the earth/moon nominal half angle, and computing formula is:

${\∂}_E=a\sin (R_E/L_E)$ ${\∂}_M=a\sin (R_M/L_M)$

Wherein, L _E, L _MBe respectively the distance of satellite and the earth/moon;

R _E, R _MBe respectively earth radius, month radius of a ball;

In second step, image patch is justified Fitting Analysis obtain the image patch half angle Calculate

With the nominal half angle error in the first step:

$\mathrm{\Δ}{\∂}_E=\left|{\∂-\∂}_E\right|$ $\mathrm{\Δ}{\∂}_M=\left|{\∂-\∂}_M\right|$

When

This image patch is the moon;

This image patch is the earth;

Other does not satisfy design conditions, stops follow-up work;

T in the formula _ΔBe angle threshold, size is determined by the nominal orbit error.

During M=2, the recognition methods of the earth, the moon is as follows in described second step:

The first step is found the solution the average gray value G of two target image patches ₁, G ₂, when

G ₂＞G ₁, mark 2 image patches are the earth, mark 1 image patch is the moon;

G ₂＜G ₁, mark 1 image patch is the earth, mark 2 image patches are the moon.

Described step (2) middle orbit is estimated to exist only to occur the moon in the image and two kinds of situations of the earth only occur, wherein:

When only the moon occurring in the image, implementation procedure is:

The first step, calculate the satellite vector under month heart inertial system:

V _SC＝C _IB×V _M

In second step, calculate a satellite and a month heart distance L _M:

$L_M=R_M/{\sin \∂}_M$

The 3rd step, the orbital position under month heart inertial system:

[X _MI Y _MI Z _MI]＝L _M·V _SC

Wherein, R _MBe the moon radius of a ball;

C _IBBe star sensor output inertia attitude battle array;

V _M,

Be respectively moon vector, month half angle under the ultraviolet sensors output body series;

Only occur terrestrial time in image, implementation procedure is:

The first step is calculated the satellite vector under the Earth central inertial system, and computing formula is:

V _SCE＝C _IB×V _E

In second step, calculate satellite and geocentric distance L _E:

$L_E=R_E/\sin {\∂}_E$

The 3rd step, the orbital position under the Earth central inertial system:

[X _EI Y _EI Z _EI]＝L _E·V _SCE

Wherein, R _EBe earth radius;

C _IBBe star sensor output inertia attitude battle array;

V _E,

Be respectively earth vector under the ultraviolet sensors output body series, half angle.

Only carrying out track and estimate that implementation procedure is in the described step (3) according to the measured value of ultraviolet sensors:

The first step is according to the moon heart vector V of ultraviolet sensors measurement _M, month half angle

The phases of the moon, the earth's core vector V _E, half angle

The ground phase, known numeric value R _M, R _ECalculate the earth's core apart from L _E, month heart is apart from L _MAnd by the phases of the moon, the solar vector V that obtains mutually _S

In second step, calculate the poor of ground month vector, V _EM=V _E-V _M

The 3rd step is according to the solar vector V that calculates _S, the difference V of month vector _EMFind the solution satellite body and be tied to inertial system attitude transition matrix C _IB

In the 4th step, utilize the transition matrix C that calculates _IB, calculate the moon heart inertial system and Earth central inertial system under orbital position [X _MIY _MIZ _MI] and [X _EIY _EIZ _EI].

Orbital position [X under described month heart inertial system and the Earth central inertial system _MIY _MIZ _MI] and [X _EIY _EIZ _EI] computing formula is:

[X _MI Y _MI Z _MI]＝L _M·V _SC

[X _EI Y _EI Z _EI]＝L _E·V _SCE

Wherein, L _M-satellite and a month heart distance, $L_M=R_M/\sin {\∂}_M;$

L _E-satellite and geocentric distance, $L_E=R_E/\sin {\∂}_E;$

V _SCSatellite vector under the-month heart inertial system, V _SC=C _IB* V _M

V _SCESatellite vector under the-Earth central inertial system, V _SCE=C _IB* V _E

R _MBe the moon radius of a ball;

R _EBe earth radius;

C _IBBe star sensor output inertia attitude battle array;

V _M,

Be respectively moon vector, month half angle under the ultraviolet sensors output body series;

V _E,

Be respectively earth vector under the ultraviolet sensors output body series, half angle.

The output valve according to the measured value of ultraviolet sensors and star sensor in the described step (3) is carried out track and is estimated that implementation procedure is:

The first step, the attitude transition matrix C that utilizes star sensor to export _IBCalculate the orbital position [X under month heart inertial system and the Earth central inertial system respectively _MIY _MIZ _MI] and [X _EIY _EIZ _EI];

Second step was that the lower railway position is transformed under moon heart inertial system with Earth central inertial, and formula is:

[X′ _MI Y′ _MI Z′ _MI]＝[X _EI Y _EI Z _EI]+[X _D Y _D Z _D]

Wherein, [X _DY _DZ _D] be tied to moon translational movement of heart inertial system for Earth central inertial;

The 3rd step, calculate the track estimated value, computing formula is:

[X _MI Y _MI Z _MI]＝([X′ _MI Y′ _MI Z′ _MI]+[X _MI Y _MI Z _MI])/2。

The present invention compared with prior art beneficial effect is:

(1) the present invention is cut apart post analysis to the image of ultraviolet sensors output, and carry out earth and moon and discern, can only utilize ultraviolet sensors to carry out track after the identification estimates, also can utilize ultraviolet sensors to cooperate star sensor to finish track estimates, compare with existing track method of estimation, the present invention makes full use of the characteristics of ultraviolet sensors, realizes that autonomous Orbit estimates, has precision height, advantage that independence is strong.

(2) the present invention adopts Flame Image Process to cooperate the method for prior imformation to carry out the earth, moon appearance judgement and Target Recognition, suppose in advance in the image to compare for the earth or the moon carry out the mode that track estimates with prior art, the present invention more can react reality, solve practical problems.

(3) to have adopted the image relative method be comparison prediction image space and actual captured figure position to estimation of error of the present invention, utilize the picture position to replace the track difference generally to adopt filtering method to obtain error with prior art and compare, the inventive method is simple, be suitable for using on the engineering.

Description of drawings

Fig. 1 is the inventive method process flow diagram;

Fig. 2 takes emulating image for the present invention to the moon.

Embodiment

It is disclosed a kind of three-axis attitude sensor that utilizes the ultraviolet spectral coverage in the US5837894 title " Wide Field of View Sensor with diffractive OpticalCorrector " that the ultraviolet sensors that relates among the present invention can adopt the patent No. of Honeywell company application.Can also adopt disclosed a kind of three-axis attitude sensor that utilizes the ultraviolet spectral coverage in U.S. Pat 5319969 titles " Method for determining 3-axis spacecraft attitude ".

As shown in Figure 1, be the inventive method process flow diagram, concrete steps are as follows:

(1) ultraviolet image is read in internal memory, introduce the inertia attitude data of star sensor output simultaneously.Utilize target partitioning algorithm in the Flame Image Process that ultraviolet image is carried out target and cut apart, partitioning algorithm can be selected thresholding method.Specific algorithm is seen article " the threshold method summary of image segmentation " systems engineering and electronic technology, 2002Vol.24 No.6.The imaging of class full circle, semicircle will appear in the lunar trajectory medium ultraviolet earth, the moon, judge whether have the earth or the moon in the captured image according to cutting apart the back target.Do not have the earth or the moon in the ultraviolet image, do not possess the condition that realizes that autonomous Orbit is estimated so, wait for image transmitted next time.Carry out the judgement and the identification of the earth/moon according to nominal track or predicted orbit according to brightness of image, imaging image patch size.Possible spatial spread target only may be the earth, the moon in the captured image of ultraviolet sensors, and the earth and the moon can appear in the picture simultaneously under certain track and attitude permission.Then do not forward (2) if the moon/earth does not coexist in the picture, otherwise forward (3) to.

Judge with recognition methods as follows:

Whether the first step forms image patch in the image after judgement is cut apart according to gray threshold, if there is not image patch, the no earth or the moon stop follow-up work, otherwise the image patch size are analyzed in the key diagram picture, calculate image patch interior pixel number N um, judge image patch interior pixel number N um and number threshold value T _NSize, if Num＞=T _N, image patch may mark for the earth or the moon, if Num＜T _N, image patch is not the earth or moon target;

Second step, the image patch number M that marks in the statistics first step,

If M=0 does not have the earth or the moon in the image, stop follow-up work;

M=1 has the earth or the moon in the image, but is not to exist simultaneously;

M=2 exists the earth, the moon simultaneously in the image;

M＞2, image memory stops follow-up work in big noise;

Wherein, gray threshold selects ground to inject or the priori default value, generally greater than 2 times maximum background gray levels;

Number threshold value T _N, decide by nominal track, crescent or half moon.

Above in second step during M=1, the recognition methods of the earth or the moon is as follows:

At first, calculate the earth/moon nominal half angle, computing formula is:

${\∂}_E=a\sin (R_E/L_E)$ ${\∂}_M=a\sin (R_M/L_M)$

Wherein, L _E, L _MBe respectively the distance of satellite and the earth/moon;

R _E, R _MBe respectively earth radius, month radius of a ball;

Then, image patch is justified Fitting Analysis and obtain the image patch half angle

Calculate

With the nominal half angle error in the first step:

${\mathrm{\Δ}\∂}_E=|\∂-{\∂}_E|$ ${\mathrm{\Δ}\∂}_M=|\∂-{\∂}_M|$

When

This image patch is the moon;

This image patch is the earth;

Other does not satisfy design conditions, stops follow-up work;

T in the formula _ΔBe angle threshold, size is determined by the nominal orbit error.

During M=2, the recognition methods of the earth, the moon is as follows in described second step:

Find the solution the average gray value G of two target image patches ₁, G ₂,

Work as G ₂＞G ₁, mark 2 image patches are the earth, mark 1 image patch is the moon;

Work as G ₂＜G ₁, mark 1 image patch is the earth, mark 2 image patches are the moon.

(2) have only the moon or target of the earth in the image this moment, star sensor output inertia attitude battle array C _IB, the moon vector V under the ultraviolet sensors output body series _MOr earth vector V _EWith moon half angle

A. looking like with lunar map is example, so satellite vector under moon heart inertial system:

V _SC＝C _IB×V _M (1)

A satellite and a month heart distance L _M, establishing moon radius of a ball is R _M:

$L_M=R_M/\sin {\∂}_M---\left(2\right)$

Orbital position under moon heart inertial system so:

[X _MI Y _MI Z _MI]＝L _M·V _SC (3)

B. be example with the earth image, calculate the satellite vector under the Earth central inertial system, computing formula is:

V _SCE＝C _IB×V _E (4)

Satellite and geocentric distance L _E:

$L_E=R_E/\sin {\∂}_E---\left(5\right)$

Orbital position under the Earth central inertial system:

[X _EI Y _EI Z _EI]＝L _E·V _SCE (6)

Wherein, R _EBe earth radius;

C _IBBe star sensor output inertia attitude battle array;

V _E,

Be respectively earth vector under the ultraviolet sensors output body series, half angle.

(3) in the captured picture of ultraviolet sensors the earth, can be directly when the moon coexists rely on single ultraviolet sensors carry out track to estimate.The measured value of this moment has a moon heart vector V _M, month half-angle

The phases of the moon, the earth's core vector V _E, half-angle The ground phase, known numeric value is R _M, R _EMeasured value can also obtain the earth's core apart from L above utilizing _E, month heart is apart from L _MAnd solar vector V _S, utilize ultraviolet sensors to calculate solar vector V under the satellite body system _SConcrete formula is as follows:

$V_S=\frac{\sin {\mathrm{\θ}}_{\mathrm{sm}}}{\sin \∂}{V}_{\mathrm{AC},B}-(\sin {\mathrm{\θ}}_{\mathrm{sm}}\mathrm{ctg}\∂+\cos {\mathrm{\θ}}_{\mathrm{sm}})M_B$

Wherein, V _{Ac, B}For satellite body is a point vector in the bright arc in the hypograph;

Be the moon/earth apparent radius; θ _SmBe moon heart vector/the earth's core vector and solar vector angle; M _BFor satellite body is a heart vector/the earth's core vector next month.Recited above

θ _Sm, M _BFor getting earth value or moon value simultaneously.

Calculate the poor of ground month vector:

V _EM＝V _E-V _M (7)

According to the two vector V of inertia _S, V _EMCan find the solution the transition matrix C of inertia to body attitude battle array _IB, two vectors are decided the appearance computing formula and are seen " attitude of satellite dynamics and control ", Yuhang Publishing House, in September, 2003.

Can calculate orbital position under moon heart inertial system and the Earth central inertial system respectively according to formula (1)～(6) in the step (2) below.

When also there is the moon in the existing earth in the image, carry out track according to the output valve of the measured value of ultraviolet sensors and star sensor and estimate, calculate the moon heart inertial system and Earth central inertial system under orbital position [X _MIY _MIZ _MI] and [X _EIY _EIZ _EI];

[X _MIY _MIZ _MI] and [X _EIY _EIZ _EI] calculating can adopt computing formula in the step (2).

Set two coordinate systems and only have translational movement [X _DY _DZ _D], so unified have to selenocentric coordinate system:

[X′ _MI Y′ _MI Z′ _MI]＝[X _EI Y _EI Z _EI]+[X _D Y _D Z _D] (8)

According to two groups of data mean estimates:

[X _MI Y _MI Z _MI]＝([X′ _MI Y′ _MI Z′ _MI]+[X _MI Y _MI Z _MI])/2 (9)

For the result who calculates is estimated and subsequent analysis, also needing to carry out orbit error after track is estimated estimates, orbit error is estimated to adopt two kinds of methods, a kind of employing calculated value and predicted value relatively obtain orbit error, another kind compares according to the earth, moon image space and actual captured figure position under the predicted orbit and to obtain error.

Embodiment:

Input ultraviolet sensors view data is seen Fig. 2, star sensor output attitude quaternion:

q＝[q ₁?q ₂?q ₃?q ₄]＝[-0.1722692?0.7476803?-0.5625995?0.3078623]

Obtain the transition matrix R that inertia is tied to body series from hypercomplex number:

$R=\left[\begin{array}{ccc}-0.751& -0.604& -0.266\\ 0.088& 0.308& -0.947\\ 0.654& -0.735& -0.177\end{array}\right]$

Judge that image obtains image two image patches are arranged, selecting the number of pixels threshold value is 100, then has only an image patch to satisfy size requirements, can judge according to nominal orbital position in addition to be the moon, obtains a moon heart vector after treatment by image and describes M at body series _b:

M _b＝[0.2247?-0.27?0.936]

Obtain apparent radius:

${\∂}_M={2.64}^0$

Distance with the moon heart:

$L_M=\frac{R_M}{\sin {\∂}_M}=\frac{1738}{0.0461}=37701\mathrm{km}$

A month heart vector is described in month heart inertial system:

M _I＝R×M _b＝[-0.2552?-0.949?0.179]

The orbital position of satellite in month heart inertial system:

SC(X，Y，Z)＝-L _M·M _I＝[9621?35778?6748]。

Inventive concept and method can be applied in other space sensitive device track estimation, as long as the method for utilizing thought of the present invention to realize all falls into protection scope of the present invention, the unspecified part of the present invention belongs to general knowledge as well known to those skilled in the art.

Claims (10)

1, the Benyue segment base is characterized in that comprising the following steps: in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor

(1) image that ultraviolet sensors is taken is cut apart, and judges whether have the earth or the moon in the captured image according to cutting apart back target geometrical feature, if the no earth or the moon then do not possess the condition that realizes that autonomous Orbit is estimated, waits for the image that take next time; Otherwise, carry out the identification of the earth, the moon, judge the whether judgement in same image of the earth and the moon, if the two not in same image, then changes step (2), otherwise change step (3);

Describedly judge whether have the earth or the moon in the captured image, determination methods is:

Whether the first step forms image patch in the image after judgement is cut apart according to gray threshold, if there is not image patch, the no earth or the moon stop follow-up work, otherwise the image patch size are analyzed in the key diagram picture, calculate image patch interior pixel number N um, judge image patch interior pixel number N um and number threshold value T _NSize, if Num＞=T _N, image patch may mark for the earth or the moon, if Num＜T _N, image patch is not the earth or moon target;

Second step, the image patch number M that marks in the statistics first step, if

M=0 does not have the earth or the moon in the image, stop follow-up work;

M=1 has the earth or the moon in the image, but is not to exist simultaneously;

M=2 exists the earth, the moon simultaneously in the image;

M＞2, image memory stops follow-up work in big noise;

Wherein, gray threshold selects ground to inject or the priori default value, generally greater than 2 times maximum background gray levels;

Number threshold value T _N, decide by nominal track, crescent or half moon;

During M=1, the recognition methods of the earth or the moon is as follows in described second step:

At first, calculate earth nominal half angle Moon nominal half angle Computing formula is:

${\∂}_E=a\sin (R_E/L_E)$ ${\∂}_M=a\sin (R_M/L_M)$

Wherein, L _EBe the distance in satellite and the earth's core, L _MBe the distance of satellite with month heart; R _EBe earth radius, R _MBe the moon radius of a ball;

Then, image patch is justified Fitting Analysis and obtain the image patch half angle

Calculate

Error with above-mentioned nominal half angle:

$\mathrm{\Δ}{\∂}_E=|\∂-{\∂}_E|$ $\mathrm{\Δ}{\∂}_M=|\∂-{\∂}_M|$

When

This image patch is the moon;

This image patch is the earth;

Other does not satisfy design conditions, stops follow-up work;

T in the formula _ΔBe angle threshold, size is determined by the nominal orbit error;

(2) according to the inertia attitude data of star sensor output, half angle, moon vector or the earth vector of ultraviolet sensors output carries out track and estimates;

(3) only carry out track and estimate, or carry out track according to the output valve of the measured value of ultraviolet sensors and star sensor and estimate according to the measured value of ultraviolet sensors.

2, Benyue segment base according to claim 1 is characterized in that in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor: after track is estimated, also carry out orbit error and estimate.

3, Benyue segment base according to claim 2 is in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor, it is characterized in that: described orbit error estimates to adopt track value and ground predicted value from host computer relatively to obtain orbit error, or takes the photograph in the image position according to the earth under the predicted orbit, moon image space and actual institute and compare and carry out the orbit error estimation.

4, according to the described Benyue segment base of claim 1 in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor, it is characterized in that: during M=2, the recognition methods of the earth, the moon is as follows in described second step:

Find the solution the average gray value G of two target image patches ₁, G ₂,

Work as G ₂＞G ₁, mark 2 image patches are the earth, mark 1 image patch is the moon;

Work as G ₂＜G ₁, mark 1 image patch is the earth, mark 2 image patches are the moon.

5, Benyue segment base according to claim 1 and 2 is characterized in that in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor: described step (2) middle orbit is estimated to exist only to occur the moon in the image and two kinds of situations of the earth only occur, wherein:

When only the moon occurring in the image, implementation procedure is:

The first step, calculate the satellite vector under month heart inertial system:

V _SC＝C _IB×V _M

In second step, calculate a satellite and a month heart distance L _M:

$L_M=R_M/\sin {\∂}_M$

The 3rd step, the orbital position under month heart inertial system:

[X _MI?Y _MI?Z _MI]＝L _M·V _SC

Wherein, R _MBe the moon radius of a ball;

C _IBThe satellite body of exporting for star sensor is tied to inertial system attitude transition matrix;

V _M,

Be respectively moon vector, moon nominal half angle under the ultraviolet sensors output body series;

Only occur terrestrial time in image, implementation procedure is:

The first step is calculated the satellite vector under the Earth central inertial system, and computing formula is:

V _SCE＝C _IB×V _E

In second step, calculate satellite and geocentric distance L _E:

$L_E=R_E/\sin {\∂}_E$

The 3rd step, the orbital position under the Earth central inertial system:

[X _EI?Y _EI?Z _EI]＝L _E·V _SCE

Wherein, R _EBe earth radius;

C _IBThe satellite body of exporting for star sensor is tied to inertial system attitude transition matrix;

V _E,

Be respectively earth vector, earth nominal half angle under the ultraviolet sensors output body series.

6, Benyue segment base according to claim 1 and 2 is characterized in that in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor: only carrying out track according to the measured value of ultraviolet sensors and estimate that implementation procedure is in the described step (3):

The first step is according to the moon vector V of ultraviolet sensors measurement _M, moon nominal half angle

The phases of the moon, earth vector V _E, earth nominal half angle

The ground phase, known numeric value moon radius R _M, earth radius R _ECalculate satellite and geocentric distance L _E, satellite and month heart distance L _MAnd solar vector V _S

In second step, calculate the poor of ground month vector, V _EM=V _E-V _M

The 3rd step is according to the solar vector V that calculates _S, the difference V of month vector _EMFind the solution satellite body and be tied to inertial system attitude transition matrix C _IB

In the 4th step, utilize the described transition matrix C that calculates _IB, calculate the moon heart inertial system and Earth central inertial system under orbital position [X _MIY _MIZ _MI] and [X _EIY _EIZ _EI].

7, Benyue segment base according to claim 6 is characterized in that in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor: the orbital position [X under described month heart inertial system and the Earth central inertial system _MIY _MIZ _MI] and [X _EIY _EIZ _EI] computing formula is:

[X _MI?Y _MI?Z _MI]＝L _M·V _SC

[X _EI?Y _EI?Z _EI]＝L _E·V _SCE

Wherein, L _M-satellite and a month heart distance, $L_M=R_M/\sin {\∂}_M;$

L _E-satellite and geocentric distance, $L_E=R_E/\sin {\∂}_E;$

V _SCSatellite vector under the-month heart inertial system, V _SC=C _IB* V _M

V _SCESatellite vector under the-Earth central inertial system, V _SCE=C _IB* V _E

R _MBe the moon radius of a ball;

R _EBe earth radius;

C _IBFor the satellite body of finding the solution in above-mentioned the 3rd step is tied to inertial system attitude transition matrix;

V _M,

Be respectively moon vector, moon nominal half angle under the ultraviolet sensors output body series;

V _E,

Be respectively earth vector, earth nominal half angle under the ultraviolet sensors output body series.

8, Benyue segment base according to claim 6 is characterized in that in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor: described solar vector V _SComputing formula is:

$V_S=\frac{\sin {\mathrm{\θ}}_{\mathrm{sm}}}{\sin {\∂}_1}{V}_{\mathrm{AC},B}-(\sin {\mathrm{\θ}}_{\mathrm{sm}}\mathrm{ctg}{\∂}_1+\cos {\mathrm{\θ}}_{\mathrm{sm}})M_B$

Wherein, V _{Ac, B}For satellite body is a point vector in the bright arc in the hypograph;

Be the moon/earth apparent radius;

θ _SmBe moon heart vector/the earth's core vector and solar vector angle;

M _BFor satellite body is a heart vector/the earth's core vector next month.

9, Benyue segment base according to claim 1 and 2 is in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor, it is characterized in that: the output valve according to the measured value of ultraviolet sensors and star sensor in the described step (3) is carried out track and is estimated that implementation procedure is:

The first step utilizes the satellite body of star sensor output to be tied to inertial system attitude transition matrix C _IBCalculate the orbital position [X under month heart inertial system and the Earth central inertial system respectively _MIY _MIZ _MI] and [X _EIY _EIZ _EI];

Second step was that the lower railway position is transformed under moon heart inertial system with Earth central inertial, and formula is:

[X′ _MI?Y′ _MI?Z′ _MI]＝[X _EI?Y _EI?Z _EI]+[X _D?Y _D?Z _D]

Wherein, [X _DY _DZ _D] be tied to moon translational movement of heart inertial system for Earth central inertial;

The 3rd step, calculate the track estimated value, computing formula is:

[X _MI?Y _MI?Z _MI]＝([X′ _MI?Y′ _MI?Z′ _MI]+[X _MI?Y _MI?Z _MI])/2。

10, Benyue segment base according to claim 9 is characterized in that in the autonomous Orbit method of estimation of ultraviolet sensors, star sensor: the orbital position [X under described month heart inertial system and the Earth central inertial system _MIY _MIZ _MI] and [X _EIY _EIZ _EI] computing formula is:

[X _MI?Y _MI?Z _MI]＝L _M·V _SC

[X _EI?Y _EI?Z _EI]＝L _E·V _SCE

Wherein, L _M-satellite and a month heart distance, $L_M=R_M/\sin {\∂}_M;$

L _E-satellite and geocentric distance, $L_E=R_E/\sin {\∂}_E;$

V _SCSatellite vector under the-month heart inertial system, V _SC=C _IB* V _M

V _SCESatellite vector under the-Earth central inertial system, V _SCE=C _IB* V _E

R _M-month the radius of a ball;

R _E-earth radius;

C _IBThe satellite body of-star sensor output is tied to inertial system attitude transition matrix;

V _M,

-be respectively moon vector, moon nominal half angle under the ultraviolet sensors output body series;

V _E,

-be respectively earth vector, earth nominal half angle under the ultraviolet sensors output body series.

Images