CN105352500A - Adaptive satellite selection method and system with celestial body interference - Google Patents

Abstract

The invention discloses an adaptive satellite selection method and system with celestial body interference. The method comprises: obtaining the optical axis information of a star sensor sent by a starlight analog module and formed by an orbit; selecting a optimum interval of observing the refractive starlight; establishing a starlight observation window according to the position of a spacecraft, determining the area of the fixed star observed by the star sensor; analyzing the starlight refraction celestial body interference to select and optimize the starlight observation window; determining an observing star area of the spacecraft according to the obtained starlight observation window to obtain a selectable refraction star range of the spacecraft, predicting the direction and information of the fixed star of which starlight occurs refraction according to the given position and time information of the spacecraft by using a starlight refraction navigation error minimum satellite selection method, and selecting the adaptive navigational star. The adaptive satellite selection method and system with celestial body interference can exclude the celestial body interference according to the given position and time information of the spacecraft, and can predict the direction and information of the fixed star of which starlight refracting accurately, so that the function of selecting the adaptive navigational star is realized.

Description

Self-adaptation satellite selection method with Disturbance of celestial bodies and system

Technical field

The present invention relates to a kind of self-adaptation satellite selection method and system, particularly relate to a kind of self-adaptation satellite selection method with Disturbance of celestial bodies and system.

Background technology

The features such as along with improving constantly of precision of star sensor, starlight location and navigation technology reaches its maturity, and it is passive, hidden, navigation error does not accumulate in time more and more cause the attention of domestic and international experts and scholars.But due to the restriction of the factor such as cost, technical conditions, it is unpractical for frequently making a flight test to the test of Star navigation system method and research, and therefore usual room semi-physical simulation experiment porch is by experiment studied both at home and abroad.

Fig. 1 is laboratory starlight semi-physical simulation experiment porch configuration diagram in prior art, and Fig. 2 is this starlight semi-physical simulation platform principle schematic.This experiment simulation plateform system function is divided into three parts: track produces simulates part, star chart acquisition process and barycenter Extraction parts, star pattern matching and navigation calculation part with starlight.Its workflow is as follows: track generation module generator orbital data, and produces star sensor optical axis information, is transferred to starlight analog module; The satellite optical axis directional information (right ascension that starlight simulation softward (celestial body simulator) generates according to track generation module, declination), star module is selected by self-adaptation, generate the star map of specific visual field, export to star chart display screen and greyscale display, the former is for the direct display of star chart, and latter simulates starlight by area source simultaneously, the starlight of now simulating is received, as original star chart data by CCD sensor (or camera); Celestial body barycenter extraction module is proceeded to after carrying out background segment, linear filtering, distortion correction by image pre-processing module; Barycenter extraction module slightly extracts through barycenter, barycenter segmented positioning, after error compensation, accurate barycenter information is proceeded to importance in star map recognition module; Importance in star map recognition, by specific way of search, accurately obtains corresponding nautical star in navigation star database; And calculate the ephemeris information of the position angle of this navigation starlight, elevation angle information and this celestial body; Finally, navigational computer calculates speed and the positional information of spacecraft by band multi-model switching observation model and starlight spacer section continuous navigation algorithm.

In this experiment simulation plateform system, self-adaptation selects star module to play an important role, and it directly determines the precision of starlight location navigation.The satellite navigation location satellite selection method of prior art mainly contains: best geometric dilution of precision method, maximum extremity tetrahedron, maximum orthogonality sciagraphy etc.Said method is all some the location stars carrying out regioselective based on minimum geometric dilution of precision, select calculated amount in star process large, needs take a long time, lack a kind of for starlight refraction navigation effective Analysis and Screening Disturbance of celestial bodies and realize select the minimum quick high accuracy of star navigation error to select star strategy.

Summary of the invention

For overcoming the deficiency that above-mentioned prior art exists, the fundamental purpose of the present invention is to provide a kind of self-adaptation satellite selection method with Disturbance of celestial bodies and system, it is by determining that observation reflects between the optimal zone of starlight, according to Space Vehicle position, set up starlight observation window, determine the fixed star region that can be observed by star sensor, and carry out the ELIMINATION OF ITS INTERFERENCE of life light and gas light, given Space Vehicle position is achieved by starlight refraction navigation error minimum satellite selection method optimum, there is the fixed star direction of refraction and the object of fixed star information in temporal information and measurable starlight, and then achieve the function that self-adaptation chooses nautical star.

For reaching above-mentioned and other object, the present invention proposes a kind of self-adaptation satellite selection method with Disturbance of celestial bodies, comprises the steps:

Step one, obtains track and produces the star sensor optical axis information exported with starlight analog module;

Step 2, between the optimal zone of selected observation refraction starlight;

Step 3, according to Space Vehicle position, sets up starlight observation window, between the optimal zone of selected observation refraction starlight, determines the fixed star region that can be observed by star sensor;

Step 4, carries out the interference analysis of starlight astrographic, screening and optimizing starlight observation window;

Step 5, according to the starlight observation window that step 4 obtains, determine that the spacecraft TV star is with, obtain spacecraft and optionally reflect star scope, by the minimum selecting-star algorithm of starlight refraction navigation error, realize given Space Vehicle position, temporal information to predict that fixed star direction and the fixed star information object of refraction occur starlight, and then under realizing Disturbance of celestial bodies, adaptive optimal chooses nautical star.

Further, in step 3, this starlight observation window is:

Wherein, wherein Re is earth radius, and hg2=50km, hg1=20km, α are the refraction angle of starlight when air height 50km, for Space Vehicle position vector.

Further, this self-adaptation satellite selection method also comprises the steps: when catching fixed star refraction starlight, by the relation of the line of fixed star-spacecraft and the angle of the sun/moon-spacecraft and the sun/moon apparent radius, the star sensor sun/moon avoiding angle, the sun and the moon are got rid of outside star sensor visual field as intense light source.

Further, the line of fixed star-spacecraft and the angle of the sun/moon-spacecraft must meet the condition being greater than the sun/moon apparent radius and the star sensor sun/moon avoiding angle sum.

Further, for the calculating of the angle of fixed star-spacecraft line and the moon-spacecraft, obtaining monthly calendar information and Space Vehicle position information, by calculating the orientation of the moon under the celestial coordinate system centered by spacecraft, and then calculating the angle of fixed star-spacecraft line and the moon-spacecraft.

Further, this self-adaptation satellite selection method also comprises the steps: when catching fixed star refraction starlight, removes the interference of starlight refraction navigation by gas light according to the size of spacecraft to Horizon observation subtended angle.

For achieving the above object, the present invention also provides a kind of self-adaptation with Disturbance of celestial bodies to select star system, comprising:

Orbital data receiver module, produces the orbital datas such as the star sensor optical axis information exported with starlight analog module for obtaining track;

Refraction interval determination module, for determining that observation reflects between the optimal zone of starlight;

Best starlight observation window determination module, for according to Space Vehicle position, sets up starlight observation window, determines the fixed star region that can be observed by star sensor;

Select star module, for the best starlight observation window according to acquisition, determine that the spacecraft TV star is with, obtain spacecraft and optionally reflect star scope, utilize starlight to reflect navigation error minimum satellite selection method and realize given Space Vehicle position, temporal information predict the fixed star direction that starlight occurs to reflect and fixed star information object, and then realize self-adaptation and choose nautical star.

Further, this self-adaptation selects star system also to comprise a day moonlight ELIMINATION OF ITS INTERFERENCE module, within this day, moonlight ELIMINATION OF ITS INTERFERENCE module is when catching fixed star refraction starlight, by the relation of the line of fixed star-spacecraft and the angle of the sun/moon-spacecraft and the sun/moon apparent radius, the star sensor sun/moon avoiding angle, the sun and the moon are got rid of outside star sensor visual field as intense light source.

Further, this self-adaptation selects star system also to comprise gas light ELIMINATION OF ITS INTERFERENCE module, for when catching fixed star refraction starlight, removes the interference of starlight refraction navigation by gas light according to the size of spacecraft to Horizon observation subtended angle.

Further, the minimum satellite selection method of this starlight refraction navigation error reflects star by rapid screening, reduces the observational error between theory calculate apparent altitude and actual observation apparent altitude, realizes high precision navigation.

Compared with prior art, a kind of self-adaptation satellite selection method with Disturbance of celestial bodies of the present invention and system are by between the optimal zone of determining to observe refraction starlight, according to Space Vehicle position, set up starlight observation window, determine the fixed star region that can be observed by star sensor, and carry out the ELIMINATION OF ITS INTERFERENCE of life light and gas light, achieve given Space Vehicle position, temporal information and measurable starlight and the fixed star direction of refraction and the object of fixed star information occur, and then achieve the function that self-adaptation chooses nautical star.

Accompanying drawing explanation

Fig. 1 is laboratory starlight semi-physical simulation experiment porch configuration diagram in prior art;

Fig. 2 is this starlight semi-physical simulation platform principle schematic;

Fig. 3 is the flow chart of steps of a kind of self-adaptation satellite selection method with Disturbance of celestial bodies of the present invention;

Fig. 4 is that 20km ~ 50km highly reflects star regularity of distribution schematic diagram;

Fig. 5 is that track one week 20km ~ 50km highly reflects star regularity of distribution lab diagram;

Fig. 6 is refraction star observation window definition schematic diagram;

Fig. 7 is that the sun affects schematic diagram to star sensor;

Fig. 8 is the sun-earth-spacecraft relative position relation schematic diagram;

Fig. 9 is spacecraft visible subtended angle schematic diagram over the ground;

Figure 10 is gas light interference analysis schematic diagram;

Figure 11 is the elevation angle, the earth's core geometrical principle figure;

Figure 12 is the system architecture diagram that a kind of self-adaptation with Disturbance of celestial bodies of the present invention selects star system.

Embodiment

Below by way of specific instantiation and accompanying drawings embodiments of the present invention, those skilled in the art can understand other advantage of the present invention and effect easily by content disclosed in the present specification.The present invention is also implemented by other different instantiation or is applied, and the every details in this instructions also can based on different viewpoints and application, carries out various modification and change not deviating under spirit of the present invention.

Fig. 3 is the flow chart of steps of a kind of self-adaptation satellite selection method with Disturbance of celestial bodies of the present invention.As shown in Figure 3, a kind of self-adaptation satellite selection method with Disturbance of celestial bodies of the present invention, comprises the steps:

Step 301, obtains track and produces the star sensor optical axis information exported with starlight analog module;

Step 302, determines that observation reflects between the optimal zone of starlight.

When reflecting starlight through (refraction is highly less than 20km) during troposphere, because troposphere temperature variation is large, convective activity strong, atmospheric density change is unstable, and indirect responsive Horizon precision is low.And when starlight refraction is highly greater than 50km, stellar refraction angle is too small, is difficult to accurate observation.20km ~ 50km is Sudden warming in stratosphere scope, and temperature variation is slow therebetween, and atmospheric density is relatively stable, the enough large and substantially constant in the refraction angle under same refraction height, and therefore, 20km ~ 50km interval is considered to observe between the optimal zone of refraction starlight.

As shown in Figure 4, s represents Space Vehicle position, c ₁for enclosing air apart from earth surface 20km place one, c ₂for enclosing air apart from earth surface 50km place one.Spacecraft is at s ₁place can observe light s ₁a ₁, s ₁a ₂respectively with R _e+ 50km and R _ethe earth border of+20km height is tangent, s ₁a ₁with s ₁a ₂between angle for spacecraft is at s ₁place can be used for the starlight scope of the responsive Horizon that can observe of navigating.Spacecraft is at s ₂place can observe light s ₂b ₁, s ₂b ₂respectively with R _e+ 50km and R _ethe earth border of+20km height is tangent, s ₂b ₁with s ₂b ₂between angle for spacecraft is at s ₂place can be used for the starlight scope of the responsive Horizon that can observe of navigating.By figure geometric relationship, easily know along with os increases gradually, diminish gradually, namely spacecraft distance the earth's core is nearer, then reflect distribution range corresponding to star just wide, and refraction star number amount is also more; Spacecraft distance the earth's core is far away, then reflect distribution range corresponding to star just narrow, and refraction star number amount is also fewer.

Namely from the whole cycle of earth orbit, refraction satellite-based is originally zonal arrangement near orbit plane, as shown in Figure 5.The width of refraction star braid is with spacecraft when the distance of carving distance the earth's core is relevant, and spacecraft distance the earth's core is near, then reflect distribution range corresponding to star just wide, and refraction star number amount is also many.

Step 303, according to Space Vehicle position, sets up starlight observation window, determines the fixed star region that can be observed by star sensor.

Step 304, carries out the interference analysis of starlight astrographic, screening and optimizing starlight observation window.

As from the foregoing, because Space Vehicle position is different, refraction star braid varies in size; Therefore according to Space Vehicle position, set up starlight observation window, determine the fixed star region that can be observed by star sensor.

As shown in Figure 6, spacecraft is defined at position vector with any fixed star starlight direction between angle for window angle, its geometric representation is:

If u1 is fixed star starlight do not occur at air height 50km place reflect critical starlight vector direction, can be derived from window angle, coboundary by formula (3.1) is:

${\mathrm{\φ}}_{max}=\arcsin \frac{\mathrm{Re}+hg2}{\left|\stackrel{\→}{r}\right|}---\left(3.2\right)$

In formula, Re is earth radius, hg2=50km.

Separately set u2 the critical starlight vector direction that reflects not to occur as fixed star starlight at air height 20km place, then push away lower boundary window angle is by formula (3.1):

${\mathrm{\φ}}_{min}=\arcsin \frac{\mathrm{Re}+hg1}{\left|\stackrel{\→}{r}\right|}-\mathrm{\α}---\left(3.3\right)$

In formula, α is the refraction angle of starlight when air height 50km, hg1=20km.

Obtain spacecraft by (3.1), (3.2) two formulas and obtain the best observation window condition reflecting starlight through 20km ~ 50km Sudden warming in stratosphere:

If certain fixed star meets this best starlight observation window conditional (3.4), namely illustrate that this fixed star will reflect between 20m ~ 50km height, can be used as navigation refraction star.

Step 305, according to the starlight observation window that step 304 obtains, determine that the spacecraft TV star is with, obtain spacecraft and optionally reflect star scope, by the minimum satellite selection method of starlight refraction navigation error, realize given Space Vehicle position, temporal information to predict that fixed star direction and the fixed star information object of refraction occur starlight, and then realizes the function that self-adaptation chooses nautical star.

From Such analysis, spacecraft distance the earth's core is near, then reflect distribution range corresponding to star just wide, and refraction star number amount is also many, and spacecraft distance the earth's core is far away, then reflect distribution range corresponding to star just narrow, and refraction star number amount is just few.The instantaneous φ of definition spacecraft _max-φ _minfor now the TV star is with, represent that spacecraft now optionally reflects star scope.

Certainly, in the indirect responsive Horizon process of starlight refraction, star sensor will inevitably be subject to day, the moon, etc. the impact of light that sends of celestial body.Compared to the starlight brightness that star sensor is normally caught, the sun, the moon, gas light light intensity are all excessively strong, need star sensor to evade.Luminous celestial body will inevitably cause available navigation in spacecraft operational process to reflect the minimizing of star number amount to the restriction that star sensor grabs star orientation, even affects navigation accuracy.Thus evading of Disturbance of celestial bodies in spacecraft in orbit process is just extremely necessary.

Preferably, after step 305, also comprise the steps: when catching fixed star refraction starlight, by the relation of the line of fixed star-spacecraft and the angle of the sun/moon-spacecraft and the sun/moon apparent radius, the star sensor sun/moon avoiding angle, the sun and the moon are got rid of outside star sensor visual field as intense light source.

When catching fixed star refraction starlight, the sun must be got rid of outside star sensor visual field as intense light source, so, as in Fig. 7, the line of fixed star-spacecraft and the angle theta of the sun-spacecraft _star-sunmust meet and be greater than sun apparent radius γ and star sensor sun avoiding angle θ _avoiding-sunthe condition of sum, that is:

θ _star-sun≥γ+θ _avoiding-sun(3.5)

Known sun distance Space Vehicle position is relatively far away, therefore angle θ _star-suncan calculate in the celestial coordinate system centered by the earth's core.

For the moon, no matter it is bright is dark, all may have an impact to the fixed star closed on, thus for its analytical approach of the moon and the sun similar.Difference is, moon distance Space Vehicle position is comparatively near, and the spacecraft relative moon position that moves in-orbit changes greatly.For spacecraft, the moon directly can not calculate the centre of sphere angle of itself and star formation region in the celestial coordinate system centered by the earth's core, need to obtain monthly calendar information and Space Vehicle position information, by calculating the orientation of the moon under the celestial coordinate system centered by spacecraft, and then calculate the angle of fixed star-spacecraft line and the moon-spacecraft.

Preferably, the self-adaptation satellite selection method of the present invention also comprises the steps: when catching fixed star refraction starlight, removes the interference of starlight refraction navigation by gas light according to the size of spacecraft to Horizon observation subtended angle.

The generation of gas light, on the one hand from the radiation of the earth, comes from earth surface reflection solar radiation on the other hand.Catch for fixed star refraction starlight for star sensor, rear one source is main interference factors.The core utilizing refraction Star navigation system is responsive Horizon, and due to the existence of circle in morning and evening, whether to be in night hemisphere most important for the Horizon of navigation moment starlight process.

As shown in Figure 8, when the sun is in position O1, look from spacecraft G position to the earth, the sun is blocked by the earth completely, and now star sensor visual field border C and D is all regions at night, and itself is not luminous, and starlight refraction navigation can work.

When the sun is in position O2, look from spacecraft G position to the earth, sunshine part is blocked by the earth, and now star sensor visual field border hemisphere in existing daytime also has night hemisphere, and at night hemisphere D place, starlight refraction navigation still can work.

When the sun is in position O3, look from spacecraft G position to the earth, the star sensor visual field border earth is daytime completely, and now starlight refraction navigation cannot work.

Known according to analysis, indirect responsive Horizon observes subtended angle size relevant with spacecraft to Horizon by the influence degree of gas light, and this subtended angle is relevant with spacecraft orbit height.

Spacecraft over the ground maximum visible subtended angle asks method as follows:

As seen from Figure 9, as can be seen here, spacecraft orbit height is higher, and the visible earth angular of spacecraft is less, and starlight refraction navigation affects less by gas light.

In Figure 10, O-SUN is solar vector direction, and O-SAT is spacecraft direction vector, and circle GOH is earth terminator.Seen by figure, the earth half portion towards the sun is in hemisphere in daytime, and the earth half portion of the sun is in night hemisphere dorsad.Consider that earth atmosphere diffuse reflection is on the impact of catching refraction star, if the point of contact F on SAT-STAR line and earth border is positioned at the night hemisphere of the earth, then catches and reflect star and do not affect by gas light, otherwise, be then subject to the irreflexive impact of air and cannot responsive Horizon.Therefore, can with point of contact be positioned at daytime hemisphere or night hemisphere weigh the impact of current time gas light for the responsive Horizon of caught refraction star.

Can be obtained by following formula by the known F point coordinate of geometric relationship in Figure 10:

In formula 3.6, space Vehicle position vector, R _efor earth radius, h _afor refraction height.After the coordinate obtaining F point, calculate the distance r between refraction point F and the sun _sUN-F, the some A on the circle in morning and evening of calculating simultaneously and the distance of the sun:

$r_{SUN-A}=\sqrt{{r_{O-SUN}}^2+{(R_e+h_a)}^2}---\left(3.7\right)$

If can r be met _sUN-F> r _sUN-A, then illustrate that the refraction starlight of catching does not affect by gas light.

Particularly, the minimum satellite selection method of starlight refraction navigation error utilized in step 305 of the present invention, specific as follows: the fixed star stellar refraction angle R that star sensor measures is measured as systematic perspective by the apparent altitude ha calculated, but due to the noise effect of star sensor measurement accuracy, and light propagates factor impacts such as being disturbed in an atmosphere, cause there is an error between theory calculate apparent altitude and actual observation apparent altitude, namely actual spacecraft is at position vector r ₀the observed reading at place by calculated value H (r ₀) and apparent altitude measurement noise composition

$\tilde{H}\left(r\right)=H\left(r_0\right)+\mathrm{\γ}---\left(3.8\right)$

γ is apparent altitude measurement noise.

Non-linear observation equation linearization is reflected to starlight,

$h_a=\sqrt{r^2-u^2}+utan\left(R\right)-R_e-a---\left(3.9\right)$

By it at position vector r ₀place does first order Taylor and launches

$\tilde{H}\left(r\right)=H\left(r_0\right)+\frac{\∂h_a}{\∂x}\mathrm{\Δ}x+\frac{\∂h_a}{\∂y}\mathrm{\Δ}y+\frac{\∂h_a}{\∂z}\mathrm{\Δ}z+V---\left(3.10\right)$

V is residual sequence.

The front starlight refractive direction of refraction is had for refraction star 1 by Figure 11 it reflects at apparent altitude ha1, and its Space Vehicle position vector is in starlight direction upper projected size is

u ₁＝rcos(θ ₁)

(3.11)

Edge, direction

Formula (3.11) is substituted into refraction observation equation (3.9), and arrangement can be expressed as formula (3.12)

$\begin{array}{c}{h}_a=f(x,y,z,\mathrm{\θ},R)\\ =\sqrt{x^2+y^2+z^2}(\sin \mathrm{\θ}+cos\mathrm{\θ}tan(R\left)\right)-R_e-a\end{array}---\left(3.12\right)$

Can obtain like this in formula (3.10)

$\left\{\begin{array}{c}\frac{\∂h}{\∂x}=\frac{x}{\sqrt{x^2+y^2+z^2}}(\sin \mathrm{\θ}+\cos \mathrm{\θ}\tan (R\left)\right)\\ \frac{\∂h}{\∂y}=\frac{y}{\sqrt{x^2+y^2+z^2}}(\sin \mathrm{\θ}+\cos \mathrm{\θ}\tan (R\left)\right)\\ \frac{\∂h}{\∂z}=\frac{z}{\sqrt{x^2+y^2+z^2}}(\sin \mathrm{\θ}+\cos \mathrm{\θ}\tan (R\left)\right)\end{array}\right.$

Existing starlight refraction navigation accuracy height navigation average position error can realize 200m.In suppositive mood (3.10), Taylor's single order launches Xiang Mo is 1, and by can be calculated within site error 500m, the magnitude of residual sequence V is 10 ^-5~ 10 ^-4, as can be seen here, when star error effect is selected in analysis, can principal element be caught, ignore residual sequence, simplify most preferably star model.

Therefore at position vector r ₀place's simultaneous formula (3.9) and formula (3.10), can obtain

$\mathrm{\γ}\≈\frac{\∂h}{\∂x}\mathrm{\Δ}x+\frac{\∂h}{\∂y}\mathrm{\Δ}y+\frac{\∂h}{\∂z}\mathrm{\Δ}z---\left(3.13\right)$

Formula (3.12) reflects apparent altitude measurement noise and is applied to Space Vehicle position vector r ₀the error effect of all directions.

If vectorial and error vector for following formula:

$\stackrel{\→}{p}=\[\begin{array}{ccc}\frac{\∂h}{\∂x}& \frac{\∂h}{\∂y}& \frac{\∂h}{\∂z}\end{array}\]$

$\mathrm{\Δ}\stackrel{\→}{r}=\[\begin{array}{ccc}\mathrm{\Δ}x& \mathrm{\Δ}y& \mathrm{\Δ}z\end{array}\]$

Then apparent altitude measurement noise γ can be expressed as

$\mathrm{\γ}=\stackrel{\→}{p}\·\mathrm{\Δ}\stackrel{\→}{r}=\left|\stackrel{\→}{p}\right|\left|\mathrm{\Δ}\stackrel{\→}{r}\right|cos\mathrm{\α}$

α represents vector and error vector between angle, arrangement can obtain

$\left|\mathrm{\Δ}\stackrel{\→}{r}\right|=\frac{\mathrm{\γ}}{\left|p\right|\·\cos \mathrm{\α}}\≥\frac{\mathrm{\γ}}{\left|p\right|}$

To make at position vector r ₀navigation error minimum, needs obtain and meet position vector r ₀locate maximum | p|.

If optimization function

J＝max|p|

Maximization optimal function

J＝sinθ+cosθ·tanR

Distortion arrangement can obtain

$J=\sqrt{1+{\tan }^{2}R}(\frac{1}{\sqrt{1+{\tan }^{2}R}}sin\mathrm{\θ}+\frac{\tan R}{\sqrt{1+{\tan }^{2}R}}cos\mathrm{\θ})=\sqrt{1+{\tan }^{2}R}sin(\mathrm{\θ}+\mathrm{\β})---\left(3.14\right)$

Wherein

$\mathrm{\β}=\arccos \frac{1}{\sqrt{1+{\tan }^{2}R}}$

When reflecting height hg and meeting 20km<hg<50km, according to starlight Continuous Observation model, stellar refraction angle R should meet 3.52 " <R<334.51 ", therefore β ≈ 0.Can be seen by formula (3.14), it is monotonically increasing function that optimal function is positioned at 0 ~ 90 ° of interval in the earth's core elevation angle theta.

And then spacecraft can be obtained when at position vector r ₀the optimum observation refraction in place rule, even spacecraft can observe some refraction stars simultaneously, selects the maximum refraction star of now the earth's core elevation angle theta to carry out navigation calculation and can reach best navigation accuracy.

Figure 12 is the system architecture diagram that a kind of self-adaptation with Disturbance of celestial bodies of the present invention selects star system.As shown in figure 12, a kind of self-adaptation with Disturbance of celestial bodies of the present invention selects star system, comprising: orbital data receiver module 110, refraction interval determination module 111, best starlight observation window determination module 112 and select star module 113.

Wherein, orbital data receiver module 110 produces the orbital datas such as the star sensor optical axis information exported with starlight analog module for obtaining track; Refraction interval determination module 111, for determining that observation reflects between the optimal zone of starlight, in present pre-ferred embodiments, is air height 20km ~ 50km between this best refracting sphere; Best starlight observation window determination module 112 is for according to Space Vehicle position, set up starlight observation window, determine the fixed star region that can be observed by star sensor, to go forward side by side the analysis of planet anaclasis Disturbance of celestial bodies, screening and optimizing starlight observation window, i.e. best starlight observation window, in the present invention, the best observation window condition that spacecraft obtains through 20km ~ 50km Sudden warming in stratosphere refraction starlight is:

Wherein, ${\mathrm{\&phi;}}_{max}=\arcsin \frac{\mathrm{Re}+hg2}{\left|\stackrel{\&RightArrow;}{r}\right|},$ ${\mathrm{\&phi;}}_{min}=\arcsin \frac{\mathrm{Re}+hg1}{\left|\stackrel{\&RightArrow;}{r}\right|}-\mathrm{\&alpha;},$ Wherein Re is earth radius, and hg2=50km, hg1=20km, α are the refraction angle of starlight when air height 50km, for Space Vehicle position.

Select star module 113, for the best starlight observation window according to acquisition, determine that the spacecraft TV star is with, obtain spacecraft and optionally reflect star scope, the minimum satellite selection method of navigation error realizes given Space Vehicle position, temporal information to predict that fixed star direction and the fixed star information object of refraction occur starlight, and then realizes the function that self-adaptation chooses nautical star to utilize starlight to reflect.

In the present invention, because spacecraft distance the earth's core is near, the distribution range that refraction star is corresponding is just wide, and refraction star number amount is also many, and spacecraft distance the earth's core is far away, and the distribution range that refraction star is corresponding is just narrow, and refraction star number amount is just few.Therefore, the instantaneous φ of spacecraft is defined _max-φ _minfor now the TV star is with, represent that spacecraft now optionally reflects star scope.

Preferably, the self-adaptation of the present invention selects star system also to comprise a day moonlight ELIMINATION OF ITS INTERFERENCE module 114, within this day, moonlight ELIMINATION OF ITS INTERFERENCE module 114 is when catching fixed star refraction starlight, by the relation of the line of fixed star-spacecraft and the angle of the sun/moon-spacecraft and the sun/moon apparent radius, the star sensor sun/moon avoiding angle, the sun and the moon are got rid of outside star sensor visual field as intense light source.

When catching fixed star refraction starlight, the sun must be got rid of outside star sensor visual field as intense light source.So, the line of fixed star-spacecraft and the angle theta of the sun-spacecraft _star-sun(as Fig. 7) must meet and be greater than sun apparent radius γ and star sensor sun avoiding angle θ _avoiding-sunthe condition of sum, that is:

θ _star-sun≥γ+θ _avoiding-sun

Known sun distance Space Vehicle position is relatively far away, therefore angle θ _star-suncan calculate in the celestial coordinate system centered by the earth's core.

For the moon, no matter it is bright is dark, all may have an impact to the fixed star closed on, thus for its analytical approach of the moon and the sun similar.Difference is, moon distance Space Vehicle position is comparatively near, and the spacecraft relative moon position that moves in-orbit changes greatly.For spacecraft, the moon directly can not calculate the centre of sphere angle of itself and star formation region in the celestial coordinate system centered by the earth's core, need to obtain monthly calendar information and Space Vehicle position information, by calculating the orientation of the moon under the celestial coordinate system centered by spacecraft, and then calculate the angle of fixed star-spacecraft line and the moon-spacecraft.

Preferably, the present invention also comprises gas light ELIMINATION OF ITS INTERFERENCE module 115, for when catching fixed star refraction starlight, removes the interference of starlight refraction navigation by gas light according to the size of spacecraft to Horizon observation subtended angle.Particularly, the refraction starlight of catching does not affect by gas light, need meet:

r _SUN-F＞r _SUN-A

Wherein, r _sUN-Ffor the distance between refraction point F (point of contact on spacecraft-fixed star line and earth border) and the sun, r _sUN-Afor the some A on morning and evening circle and the distance of the sun.

In sum, a kind of self-adaptation satellite selection method with Disturbance of celestial bodies of the present invention and system are by between the optimal zone of determining to observe refraction starlight, according to Space Vehicle position, set up starlight observation window, determine the fixed star region that can be observed by star sensor, and carry out the ELIMINATION OF ITS INTERFERENCE of life light and gas light, utilize starlight to reflect the minimum satellite selection method of navigation error and achieve the object that fixed star direction and the fixed star information reflected occurs for given Space Vehicle position, temporal information and measurable starlight, and then achieve the function that self-adaptation chooses nautical star.The present invention selects star process computation amount little, and computing time is few, can realize fast-moving star light guide location.

Above-described embodiment is illustrative principle of the present invention and effect thereof only, but not for limiting the present invention.Any those skilled in the art all without prejudice under spirit of the present invention and category, can carry out modifying to above-described embodiment and change.Therefore, the scope of the present invention, should listed by claims.

Claims (10)

1. the self-adaptation satellite selection method with Disturbance of celestial bodies, comprises the steps:

Step one, obtains track and produces the star sensor optical axis information exported with starlight analog module;

Step 2, between the optimal zone of selected observation refraction starlight;

Step 3, according to Space Vehicle position, sets up starlight observation window, between the optimal zone of selected observation refraction starlight, determines the fixed star region that can be observed by star sensor;

Step 4, carries out the interference analysis of starlight astrographic, screening and optimizing starlight observation window;

Step 5, according to the starlight observation window that step 4 obtains, determine that the spacecraft TV star is with, obtain spacecraft and optionally reflect star scope, by the minimum satellite selection method of starlight refraction navigation error, realize given Space Vehicle position, temporal information to predict that fixed star direction and the fixed star information object of refraction occur starlight, and then under realizing Disturbance of celestial bodies, adaptive optimal chooses nautical star.

2. a kind of self-adaptation satellite selection method with Disturbance of celestial bodies as claimed in claim 1, is characterized in that, in step 3, this best starlight observation window is:

Wherein, ${\mathrm{\&phi;}}_{max}=arcsin\frac{\mathrm{Re}+hg2}{\left|\stackrel{\&RightArrow;}{r}\right|},$ ${\mathrm{\&phi;}}_{min}=arcsin\frac{\mathrm{Re}+hg1}{\left|\stackrel{\&RightArrow;}{r}\right|}-\mathrm{\&alpha;},$ Wherein Re is earth radius, and hg2=50km, hg1=20km, α are the refraction angle of starlight when air height 50km, for Space Vehicle position vector.

3. a kind of self-adaptation satellite selection method with Disturbance of celestial bodies as claimed in claim 1, it is characterized in that, this self-adaptation satellite selection method also comprises the steps: when catching fixed star refraction starlight, by the relation of the line of fixed star-spacecraft and the angle of the sun/moon-spacecraft and the sun/moon apparent radius, the star sensor sun/moon avoiding angle, the sun and the moon are got rid of outside star sensor visual field as intense light source.

4. a kind of self-adaptation satellite selection method with Disturbance of celestial bodies as claimed in claim 3, is characterized in that: the line of fixed star-spacecraft and the angle of the sun/moon-spacecraft must meet the condition being greater than the sun/moon apparent radius and the star sensor sun/moon avoiding angle sum.

5. a kind of self-adaptation satellite selection method with Disturbance of celestial bodies as claimed in claim 4, it is characterized in that: for the calculating of the angle of fixed star-spacecraft line and the moon-spacecraft, obtain monthly calendar information and Space Vehicle position information, by calculating the orientation of the moon under the celestial coordinate system centered by spacecraft, and then calculate the angle of fixed star-spacecraft line and the moon-spacecraft.

6. a kind of self-adaptation satellite selection method with Disturbance of celestial bodies as claimed in claim 4, it is characterized in that, this self-adaptation satellite selection method also comprises the steps: when catching fixed star refraction starlight, removes the interference of starlight refraction navigation by gas light according to the size of spacecraft to Horizon observation subtended angle.

7. the self-adaptation with Disturbance of celestial bodies selects a star system, comprising:

Orbital data receiver module, produces the orbital datas such as the star sensor optical axis information exported with starlight analog module for obtaining track;

Refraction interval determination module, for determining that observation reflects between the optimal zone of starlight;

Best starlight observation window determination module, for according to Space Vehicle position, sets up starlight observation window, determines the fixed star region that can be observed by star sensor, planet anaclasis Disturbance of celestial bodies analysis of going forward side by side, screening and optimizing starlight observation window;

Select star module, for the best starlight observation window according to acquisition, determine that the spacecraft TV star is with, obtain spacecraft and optionally reflect star scope, utilize starlight to reflect navigation error minimum satellite selection method and realize given Space Vehicle position, temporal information predict the fixed star direction that starlight occurs to reflect and fixed star information object, and then realize self-adaptation and choose nautical star.

8. a kind of self-adaptation with Disturbance of celestial bodies as claimed in claim 7 selects star system, it is characterized in that: this self-adaptation selects star system also to comprise a day moonlight ELIMINATION OF ITS INTERFERENCE module, within this day, moonlight ELIMINATION OF ITS INTERFERENCE module is when catching fixed star refraction starlight, by the relation of the line of fixed star-spacecraft and the angle of the sun/moon-spacecraft and the sun/moon apparent radius, the star sensor sun/moon avoiding angle, the sun and the moon are got rid of outside star sensor visual field as intense light source.

9. a kind of self-adaptation with Disturbance of celestial bodies as claimed in claim 7 selects star system, it is characterized in that: this self-adaptation selects star system also to comprise gas light ELIMINATION OF ITS INTERFERENCE module, for when catching fixed star refraction starlight, remove the interference of starlight refraction navigation by gas light according to the size of spacecraft to Horizon observation subtended angle.

10. a kind of self-adaptation with Disturbance of celestial bodies as claimed in claim 7 selects star system, it is characterized in that: the minimum satellite selection method of this starlight refraction navigation error reflects star by rapid screening, reduce the observational error between theory calculate apparent altitude and actual observation apparent altitude, realize high precision navigation.