CN102564454B - Measuring data simulation method for autonomous navigation system based on sun-earth-moon azimuth information - Google Patents

Abstract

A measuring data simulation method for an autonomous navigation system based on sun-earth-moon azimuth information respectively comprises simulation for reference pulse, earth pulse, moon pulse and sun pulse. The simulation and measurement data accurately and comprehensively reflects the influence on sun-earth-moon integrated sensor measurement caused by various complex factors, including the influence on the compression of the earth and the shape change of the earth. According to the measuring data simulation method disclosed by the invention, the simulation of pulse information of the sun-earth-moon integrated sensor and the quantitative assessment for system errors of the sensor and the verification for the performances of the navigation system can be effectively carried out in a number value artificial form.

Description

Based on the measurement data analogy method of autonomous navigation system of day ground moon azimuth information

Technical field

The present invention relates to a kind of autonomous navigation simulation test method, particularly a kind of method for numerical simulation based on ground day, integration moon sensor metrical information, belongs to autonomous navigation technology field.

Background technology

Autonomous navigation technology refers to that satellite is not when relying on ground system and supporting, only relies on spaceborne measuring equipment to determine position and the speed of satellite in-orbit in real time, also claims autonomous Orbit to determine.For satellite system, independent navigation is conducive to reducing satellite to the degree of dependence on ground, improves system survivability, such as wartime, when ground control station suffers destruction and the interference of enemy, still can complete determination and the maintenance of track, this has very important significance to military satellite.In addition, independent navigation effectively can also alleviate the burden of ground control station, reduces ground and supports cost, thus reduce the development cost of whole space program.Independent navigation is that satellite realizes from the basic premise of main control and basis, is also one of the gordian technique of structure constellation, Space-based network.

The MANS autonomous navigation system of Microcosm company of U.S. development, utilize special integrated sensor according to day, ground, the inflight measurement data of the moon determine track and the three-axis attitude of spacecraft in real time, autonomous navigation system (Anthony J.Autonomous Space Navigation Experiment [C] the .AIAA 92-1710 in complete meaning, 1992. and Collins J T and Conger R E.MANS:Autonomous Navigation and Orbit Control for Communication Satellites [C] .AIAA 94-1127-CP, 1994. and Hosken R W, Wertz J R.Microcosm autonomous navigation system on-orbit operation [C] .Keystone, Colorado:The 18th Annual AAS Guidance and Control Conference, AAS 95-074, 1995.).Because the sun, the earth, the moon are nature celestial bodies, its characteristics of motion is known and is easily identified, navigating resources is unrestricted.Therefore, study this autonomous navigation technology and there is important engineer applied value." space test platform-D " spacecraft that USAF is launched as far back as in March, 1994 just carries and the system have been flight test (Wertz J R.Implementing autonomous orbit control [C] .Breckenridge in-orbit, Colorado:Proceedings of the Annual AAS Guidance and Control Conference, AAS 96-004,1996.).Beijing Control Engineering Inst. to day ground moon autonomous navigation system navigation sensor feature and the navigation algorithm of system carried out theoretical analysis and research (Li Jie, Chen Yiqing. the new development [J] of Spacecraft Autonomous Navigation Technology. control engineering, 1997, (1): 76 ~ 81.).

But, because traditional also exists the low problem of navigation accuracy based on the Autonomous Navigation Algorithm measured by ground moon day, greatly affect and constrain development and the application (TaiFrank of the autonomous navigation system measured based on ground moon day, Noerdlinger Peter D.A Low Cost Autonomous Navigation System [C] .AAS 89-001, 1989. and Daniele Mortari.Moon-Sun Attitude Sensor [J] .Journal of Spacecraft and Rockets, 1997, 34 (3): 360 ~ 364. and Li Jilu, Chen Yiqing, Sun Chengqi. a kind of navigation accuracy analysis [J] of MANS autonomous navigation system. control engineering, 1998, (5): 1 ~ 6.).For this navigational system,

For based on day ground the moon autonomous navigation system, the principal element restricting its navigation accuracy derives from the determination in direction, the earth's core and moon heart direction.And to affect the key factor determined in direction, the earth's core be compression of the earth, affecting the factor determined in moon heart direction is the phases of the moon.And the impact of these factors is all embodied directly on the pulse data of sensor.Therefore the analogue system setting up ground day, the navigation moon is necessary.But current semi-physical simulation verification system also cannot simulate the impact that compression of the earth is measured day ground moon integration sensor; Yet cannot realize the high-precision analog to moon change of shape, therefore, be necessary the measuring principle studying navigational system, the pulse of sensor is simulated, and simulating, verifying is carried out to air navigation aid, set up the autonomous navigation simulation verification system based on metrical information.

Summary of the invention

Technology of the present invention is dealt with problems and is: overcome existing deficiency, provides the measurement data analogy method of a kind of day ground moon autonomous navigation system, can be used in the amendment scheme studying Navigation system error, improves the navigation performance of day ground moon autonomous navigation system.

Technical solution of the present invention is: based on the measurement data analogy method of autonomous navigation system of day ground moon azimuth information, realize as follows:

(1) ground day, integration moon sensor reference pulse, reference pulse t _refprovided by following formula:

t _ref＝B _R/ω _rot+kT _s(1)

Wherein: B _rfor the plane of symmetry is relative to the drag angle of surving coordinate system X-axis, ω _rotfor scanning rotating speed, T _sfor the scan period, k is the scanning number of turns;

(2) ground day, integration moon sensor earth impulse, step is as follows:

The first step: the positional information r being obtained aircraft by dynamics of orbits model integration, under being projected in inertia is wherein subscript represent transposition, subscript I represents this variable drop under inertial system, and defines gesture mode, provides the attitude R of aircraft _iSE;

Second step: within a scan period, get any time t, in conjunction with attitude information, calculates the vector of unit length of this moment scanning sight line expression in inertial coordinates system

${\widehat{\mathrm{\ρ}}}^I=\left[\begin{array}{c}\sin \left({\mathrm{\γ}}_i\right)\cos ({\mathrm{\ω}}_{\mathrm{rot}}(t-t_{\mathrm{ref}})-B_{\mathrm{Ri}})\\ \sin \left({\mathrm{\γ}}_i\right)\sin ({\mathrm{\ω}}_{\mathrm{rot}}(t-t_{\mathrm{ref}})-B_{\mathrm{Ri}})\\ \cos \left({\mathrm{\γ}}_i\right)\end{array}\right]---\left(2\right)$

Wherein: γ _ifor the semi-cone angle of scanned cone, B _rifor scanning head is relative to the drag angle of sensor reference line, i=1, the 2 corresponding little vertebra of difference and augers;

3rd step: definition Fun:=b ²-4ac also calculates the size of Fun, wherein:

$a={\widehat{\mathrm{\ρ}}}_1^2+{\widehat{\mathrm{\ρ}}}_2^2+{\widehat{\mathrm{\ρ}}}_3^2/{(1-\mathrm{ee})}^2$

$b=2{\widehat{\mathrm{\ρ}}}_{1}x+2{\widehat{\mathrm{\ρ}}}_{2}y+2{\widehat{\mathrm{\ρ}}}_{3}z/{(1-\mathrm{ee})}^2$

$c=x^2+y^2+z^2/{{(1-\mathrm{ee})}^2-R}_E^2$

In above formula: ee is compression of the earth, R _efor the equatorial radius of earth reference ellipsoid;

4th step: if Fun=0, then t is the pulse moment.Otherwise, then get the another one moment, repeat the first step to the 4th step, calculate corresponding Fun; By the Fun value in two moment, utilizing the time t that dichotomy finds Fun=0 corresponding, is also the pulse moment;

(3) ground day, integration moon sensor moon pulse simulation, step is as follows:

The first step: utilize ephemeris to differentiate the phases of the moon by current date;

Second step: the positional information r being obtained aircraft by dynamics of orbits model integration, under being projected in inertia is obtain the moon position under inertial system by ephemeris, and define gesture mode, provide the attitude R of aircraft _iSE;

3rd step: utilize coordinate transform to calculate the expression of star moon vector under sensor surving coordinate system, thus position angle φ can be obtained _mwith elevation angle δ _m;

4th step: calculate and do not consider the moon pulse t that the phases of the moon affects _1moonand t _2moon:

${\mathrm{\φ}}_m={\mathrm{\ω}}_{\mathrm{rot}}(\frac{t_{1\mathrm{moon}}+t_{2\mathrm{moon}}}{2}-t_{\mathrm{ref}})$ (3)

${\mathrm{\σ}}_m=\frac{1}{2}({\mathrm{\ω}}_{\mathrm{rot}}(t_{2\mathrm{moon}}-t_{1\mathrm{moon}})-\mathrm{\θ})$

Wherein: σ _m=arcsin (tan δ _m/ cos β _s), β _sfor sensor slit is relative to the pitch angle of scanning ball meridian, θ is the centre of sphere angle of Slot bottom;

5th step: if the current phases of the moon is full moon, then the pulse that the 4th step obtains is the output moon pulse of sensor, if crescent or half moon, then need the basis of the pulse exported in the 4th step to be added delay δ t as output pulse, wherein δ t is that following formula provides:

$\mathrm{\&delta;t}=\left\{\begin{array}{cc}\frac{2\mathrm{\&delta;}{t}_{\max }}{\mathrm{\&pi;}}{\mathrm{\&theta;}}_x& ({\mathrm{\&theta;}}_x<\mathrm{\&pi;}/2)\\ \frac{2{\mathrm{\&delta;t}}_{\max }}{\mathrm{\&pi;}}(\mathrm{\&pi;}-{\mathrm{\&theta;}}_x)& ({\mathrm{\&theta;}}_x\&GreaterEqual;\mathrm{\&pi;}/2)\end{array}\right.---\left(4\right)$

Wherein: θ _xfor scan axis and x _mthe angle of namely vertical with life direction in normal society plane axle, δ t _maxfor maximum-delay amount.δ t < 0 when the moon at the first quarter, δ t > 0 when the moon at the last quarter;

(4) ground day, integration moon sensor sun pulse simulation, step is as follows:

The first step: the positional information r being obtained aircraft by dynamics of orbits model integration, under being projected in inertia is obtain the moon position under inertial system by ephemeris, and define gesture mode, provide the attitude R of aircraft _iSE;

Second step: utilize coordinate transform to calculate the expression of star day vector under sensor surving coordinate system, thus the position angle φ of the sun can be obtained _swith elevation angle δ _s;

3rd step: calculate sun pulse t _1sunand t _2sun:

${\mathrm{\&phi;}}_s={\mathrm{\&omega;}}_{\mathrm{rot}}(\frac{t_{1\mathrm{sun}}+t_{2\mathrm{sun}}}{2}-t_{\mathrm{ref}})$ (5)

${\mathrm{\&sigma;}}_s=\frac{1}{2}({\mathrm{\&omega;}}_{\mathrm{rot}}(t_{2\mathrm{sun}}-t_{1\mathrm{sun}})-\mathrm{\&theta;})$

Wherein: σ _s=arcsin (tan δ _s/ cos β _s), β _sfor sensor slit is relative to the pitch angle of scanning ball meridian, θ is the centre of sphere angle of Slot bottom;

Calculate ground day, the integration moon reference pulse of sensor, earth impulse, moon pulse and sun pulse by above-mentioned, thus complete the simulation of ground day, integration moon sensor pulse information.

The present invention's advantage is compared with prior art: as can be seen from the document delivered and patent, at present to based on day ground moon information autonomous navigation method mathematical simulation adopt the direct modeling moon heart, the earth's core and solar direction, do not simulate the direct output pulse of sensor, therefore cannot consider the impact of compression of the earth and moon phase variety.The direct modeling of the present invention output pulse data of sensor, and consider the impact of compression of the earth and the phases of the moon.Analogue measurement data accurately and all sidedly can reflect the impact that complicated factor (compression of the earth, the phases of the moon) is measured day ground moon integration sensor.

Accompanying drawing explanation

Fig. 1 is the structure of navigation sensor;

Fig. 2 is the scheme of installation of navigation sensor;

Fig. 3 is the visual field of navigation sensor;

Fig. 4 is basic pulse schematic diagram of the present invention;

Fig. 5 is Horizon sweep vector geometric relationship figure;

Fig. 6 is infrared horizon pulse simulation process flow diagram of the present invention;

The membership function of Fig. 7 corresponding to the phases of the moon of the present invention;

Fig. 8 is the responsive moon schematic diagram of sensor;

Fig. 9 is the projection of moon light in sensor surving coordinate system;

Scan axis and the moon orientation schematic diagram when Figure 10 is the moon at the first quarter of the present invention;

Figure 11 is scanning rotor shaft direction of the present invention and pulse daley schematic diagram;

Figure 12 is moon pulse simulation process flow diagram of the present invention;

Figure 13 is the responsive sun schematic diagram of sensor;

Figure 14 is the projection of sunray in sensor surving coordinate system;

Figure 15 is sun pulse simulation process flow diagram of the present invention;

Figure 16 is the earth impulse (not considering compression of the earth) that the scanned cone of semi-cone angle of the present invention 38 ° scans;

Figure 17 to be semi-cone angle of the present invention be earth impulse (not considering compression of the earth) that the scanned cone of 73 ° sweeps to;

Figure 18 is sun pulse of the present invention (place of countless certificate shows that satellite is in shadow region);

Figure 19 is moon pulse of the present invention (place of countless certificate shows that satellite is in shadow region);

Figure 20 is the error angle of with not having the moon of ellipticity angle of the present invention;

Figure 21 is the error angle of with not having the day of ellipticity angle of the present invention;

Figure 22 is the earth impulse (consideration compression of the earth) that the scanned cone of semi-cone angle of the present invention 38 ° scans;

Figure 23 is the earth impulse (consideration compression of the earth) that the scanned cone of semi-cone angle of the present invention 73 ° scans;

Figure 24 is the angle error angle, moon ground that there is ellipticity of the present invention;

Figure 25 is the angle error angle, day ground that there is ellipticity of the present invention;

Pulse (band postpones) when Figure 26 is the moon at the first quarter of the present invention;

Moon ground error angle when Figure 27 is the moon at the first quarter of the present invention;

Figure 28 is realization flow figure of the present invention.

Embodiment

First, the measure geometry relation of ground day, integration moon sensor is analyzed, analyzes sensor scanning one and enclose the pulse number obtained; The analysis result of measure geometry can obtain the analogy method of reference pulse very intuitively; According to integrated sensor measure geometry relation, utilize earth ellipsoid equation, set up the simulation to earth impulse, provide the pulse moment that sensor swept into, scanned out the earth; Then set up phases of the moon model, according to track, ephemeris and attitude information, utilize sensor measure geometry relation determination sensor slit visual field to sweep to moon pulse; Last according to solar direction, utilize sensor measure geometry relation determination sensor slit visual field to sweep to sun pulse.

One, sensor measure geometry is analyzed

Navigation sensor by double cone scanning type infrared earth sensor and two fan-shaped slit visual fields scan-type day, the moon sensor form.Its structural drawing as shown in Figure 1.

Conical scanning formula infrared earth sensor has single optical scanning head.Utilize mirror structure to obtain two infrared visual fields, after scanning, the track of infrared visual field is two coaxial circular cones.Optical head scanning one circle, pyroelectric detector can detect that at most four Horizons pass through signal, and the moment occurred by signal can determine the orientation of the earth's core direction vector relative to satellite, and can try to achieve the distance of satellite to the earth's core.The basis of double cone scanning type infrared senstive device adds two visible light sensors.In the scanning process of optical head, fan-shaped visual field is inswept space and the scanning spherical zone region of pivoting angle between 23 ° and 87 °.Utilize si-photodiode detector responsively can arrive the sun and the moon, can in the hope of the orientation of its direction vector relative to satellite according to the moment that the sun, the moon occur in fan-shaped visual field.Detecting device has multiple Intensity threshold, can distinguish the sun and moon signal, and can reject earth signal.

Autonomous navigation system adopts two navigation sensors, and its mounting structure relative to satellite as shown in Figure 2.The geometric relationship of navigation sensor visual field relation is shown in Fig. 3.As shown in the figure, two circular cone infrared scan visual fields are designated as respectively infrared scan cone 1 and infrared scan cone 2, corresponding semi-cone angle is designated as γ respectively ₁, γ ₂, γ here ₁=38 °, γ ₂=73 °.Two fan-shaped slit visible light sensor visual fields are designated as slit visual field 1 and slit visual field 2 respectively, the plane of symmetry of slit visual field 1 and slit visual field 2 is designated as M ₁-M ₂face.The focal length that slit visual field 1 is ahead of slit visual field 2 on sensor equatorial plane is designated as θ, slit visual field 1 and slit visual field 2 are designated as β relative to the angle of inclination of scanning rotating shaft, here θ=4 °, β=16 °, if two diameters be the infrared visual field 1 of 2.5 ° and infrared visual field 2 on sensor equatorial plane along direction of scanning relative to M ₁-M ₂the delayed angular distance in face is respectively B _r1, B _r2.

In the process of navigation sensor scanning one circle, when infrared scan 1 and infrared scan 2 sweep, scan out the earth, the sun, the moon are swept in slit visual field 1 and slit visual field 2, and the plane of symmetry of slit visual field 1 and slit visual field 2 by the reference point that is connected with sensor time, navigation sensor all provides relevant pulse signal, can obtain 9 pulse moment thus.

As shown in figure 28, the present invention is implemented as follows implementation procedure of the present invention:

Two, the simulation of reference pulse

This pulse refers to the pulse moment of plane of symmetry M1-M2 by reference point X-axis, is designated as t _ref, as shown in Figure 4, can be calculated by following formula:

t _ref＝B _R/ω _rot+kT _s(6)

Wherein: B _rfor the plane of symmetry is relative to X _sEdrag angle, ω _rotfor scanning rotating speed, T _sfor the scan period, k is the scanning number of turns.

Three, earth sensor pulse simulation

As shown in Figure 5, for the unit vector of infrared scan probe, for t, its component under the first sensor surving coordinate system can be expressed as:

$\widehat{\mathrm{\&rho;}}=\left[\begin{array}{c}\sin \left({\mathrm{\&gamma;}}_i\right)\cos ({\mathrm{\&omega;}}_{\mathrm{rot}}(t-t_{\mathrm{ref}})-B_{\mathrm{Ri}})\\ \sin \left({\mathrm{\&gamma;}}_i\right)\sin ({\mathrm{\&omega;}}_{\mathrm{rot}}(t-t_{\mathrm{ref}})-B_{\mathrm{Ri}})\\ \cos \left({\mathrm{\&gamma;}}_i\right)\end{array}\right]---\left(7\right)$

be the vector that earth Equivalent ellipsoid face is pointed in the earth's core, meet ellipsoid equation:

${x_e}^2+{y_e}^2+{z_e}^2/{(1-\mathrm{ee})}^2=R_E^2---\left(8\right)$

Wherein: ee is compression of the earth.

The vector making infrared scan pop one's head in and earth Equivalent ellipsoid face intersection length are l, can be obtained by the geometric relationship of Fig. 5:

$R_E=r+R_{\mathrm{ISE}}l\widehat{\mathrm{\&rho;}}---\left(9\right)$

Wherein, for sensor surving coordinate is tied to the rotation matrix of inertial coordinates system, R _sEIfor inertial coordinate is tied to the rotation matrix of sensor surving coordinate system, when pulse simulation, attitude is given in advance.Therefore be known. be the expression of vector under inertial system pointing to satellite from the earth's core, can be obtained by dynamics of orbits integration.

Order (9) are updated in ellipsoid equation (8) and have:

${(x+l{\widehat{\mathrm{\&rho;}}}_1)}^2+(y+l{\widehat{\mathrm{\&rho;}}}_2)y^2+{(z+l{\widehat{\mathrm{\&rho;}}}_3)}^2/{(1-\mathrm{ee})}^2=R_E^2---\left(10\right)$

Equation (10) is the quadratic equation about l, can obtain its abbreviation:

al ²+bl+c＝0 (11)

Wherein:

$a={\widehat{\mathrm{\&rho;}}}_1^2+{\widehat{\mathrm{\&rho;}}}_2^2+{\widehat{\mathrm{\&rho;}}}_3^2/{(1-\mathrm{ee})}^2$

$b=2{\widehat{\mathrm{\&rho;}}}_{1}x+2{\widehat{\mathrm{\&rho;}}}_{2}y+2{\widehat{\mathrm{\&rho;}}}_{3}z/{(1-\mathrm{ee})}^2$

$c=x^2+y^2+z^2/{{(1-\mathrm{ee})}^2-R}_E^2$

When infrared sweep vector scanning is less than horizontal circle time, (11) do not have real root; When just sweeping horizontal circle edge, (11) have two equal real roots; When sweeping to earth ellipsoid inside, (11) have two real roots do not waited.

From the character of the root of quadratic equation, for the pulse moment sweeping into, scan out the earth, (11) formula has:

Fun＝b ²-4ac＝0

For the pulse moment sweeping the earth, Fun passes through 0 point with the form increased progressively; For the pulse moment scanning out the earth, Fun passes through 0 point with the form of successively decreasing.

Earth infrared pulse modeling process chart is shown in Fig. 6:

Four, moon pulse simulation

(1) phases of the moon model

For the pulse of the moon, only consider four kinds of phases of the moons: crescent, the moon at the first quarter, full moon and the moon at the last quarter, when the phases of the moon is crescent time, do not observe the moon, can not export moon pulse; When crescent or half moon, the moon photocentre observed and moon barycenter do not overlap, and therefore need to add a small delay (it be negative that the moon at the first quarter postpones, and moon at the last quarter delay is just) on the basis of principle; Delay need not be added when full moon.Eight phases are divided on the phases of the moon is strict, p is made to be that the moon illuminates the area of part and the number percent of the total area, these eight phases of the moons are respectively: crescent (new moon p=0), crescent (waxing crescent moon 0 < p < 0.5), the moon at the first quarter (first quarter half moon p=0.5), gradually waxing gibbous (waxing gibbous moon 0.5 < p < 1), full moon (full moon p=1), gradually waning gibbous (waning gibbous moon 0.5 < p < 1), the moon at the last quarter (last quarter half moon p=0.5), waning moon (waning cresent moon 0 < p < 0.5).4 phases are only considered in simulation due to moon pulse, and this just needs to utilize the principle of fuzzy mathematics to be these four kinds of phases of the moons by other phases of the moon fuzzy approximations.Regard the phases of the moon as logical variable, its logical value has: crescent, crescent or half moon and full moon.Using p estimating as the phases of the moon, p is continuous print, needs to carry out obfuscation (fuzzification) to it.As shown in Figure 7, membership function (Z-type) μ of crescent _n(p) be:

${\mathrm{\&mu;}}_N\left(p\right)=\left\{\begin{array}{cc}1& p{<p}_1\\ \frac{p_2-p}{p_2-p_1}& p_1\&le;p{\&le;p}_2\\ 0& p>p_3\end{array}\right.---\left(12\right)$

Membership function (∏ type) μ of crescent or half moon _q(p) be:

${\mathrm{\&mu;}}_Q\left(p\right)=\left\{\begin{array}{cc}0& p\&le;p_1\\ \frac{p-p_1}{p_2-p_1}& p_1<p<p_2\\ 1& p_2\&le;p\&le;p_3\\ \frac{p_4-p}{p_4-p_3}& p_3<p<p_4\\ 0& p{\&GreaterEqual;p}_4\end{array}\right.---\left(13\right)$

Membership function (Z-type) μ of full moon _f(p) be:

${\mathrm{\&mu;}}_F\left(p\right)=\left\{\begin{array}{cc}1& p{<p}_3\\ \frac{p-p_3}{p_4-p_3}& p_3<p{<p}_4\\ 0& p_4\&le;p\&le;p_5\end{array}\right.---\left(14\right)$

From the union of fuzzy set, phases of the moon during pulse simulation differentiates as follows:

$\left\{\begin{array}{cc}\mathrm{new\; moon}& p_0\&le;p\&le;\frac{p_1+p_2}{2}\\ \mathrm{quarter\; half\; moon}& \frac{p_1+p_2}{2}<p<\frac{p_3+p_4}{2}\\ \mathrm{full\; moon}& \frac{p_3+p_4}{2}\&le;p{\&le;p}_5\end{array}\right.---\left(15\right)$

Wherein p _iimplication as shown in Figure 7, represent μ _n(p), μ _q(p), μ _fthe frontier point of (p).

(2) moon pulse simulation during full moon

When full moon, the photocentre of the moon overlaps with barycenter, and the moon can be approximated to be luminous point.As shown in Figure 8, Q ₁q ₂and Q ₃q ₄be that two slit sensors are used for the responsive sun and moon signal, two slits are symmetrical, are β=16 ° relative to the pitch angle of scanning ball meridian.The centre of sphere angle of Slot bottom is θ=4 °.In t, the coordinate of moon direction vector under surving coordinate system can be tried to achieve by ephemeris and coordinate transform, thus two position angle δ _m, φ _mknown (see Fig. 9).Work as GQ ₁₂with GS ₁when overlapping, this moment is exactly the pulse moment t that slit visual field 1 scans the moon _1moon; Work as GQ ₃₄with GS ₁when overlapping, this moment is exactly the pulse moment t that slit visual field 2 scans the sun _2moon.

Operation steps is as follows:

I). calculate the ground moon vector of t, and the expression of geaster vector under inertial system;

Ii). utilize coordinate transform to calculate the expression of star moon vector under sensor surving coordinate system, thus can to two position angle δ _m, φ _m;

Iii). calculate σ _m=arcsin (tan δ _m/ cos β _s);

Iv). obtain the pulse moment by following two equations:

${\mathrm{\&phi;}}_m={\mathrm{\&omega;}}_{\mathrm{rot}}(\frac{t_{1\mathrm{moon}}+t_{2\mathrm{moon}}}{2}-t_{\mathrm{ref}})$ (16)

${\mathrm{\&sigma;}}_m=\frac{1}{2}({\mathrm{\&omega;}}_{\mathrm{rot}}(t_{2\mathrm{moon}}-t_{1\mathrm{moon}})-\mathrm{\&theta;})$

V). export pulse moment t _1moon, t _2moon.

(3) moon pulse simulation during upper lower edge

Pulse simulation under this situation and full moon situation is similar, but distinguishes to some extent again with full moon.When the upper moon at the last quarter, the barycenter of the moon does not overlap with the photocentre of scanning.Time simulation mechanism produces pulse, delay should be added on the basis of full sized pules.The following detailed description of the computing method postponed.

As shown in Figure 10.As scan axis Z _sEwith Ox _mtime parallel, sensor from top to bottom scans first quarter moon, and photocentre overlaps with barycenter, and pulse does not postpone.Work as Z _sEat o-y _mz _mwhen plane, sensor scans first quarter moon from left to right, photocentre p _mwith barycenter o _mdo not overlap, need in full sized pules, add one and postpone δ t (when the moon at the first quarter δ t < 0, δ t > 0 when the moon at the last quarter).Work as Z _sEbe parallel to oz _mor oy _mwhen obtaining, δ t is now maximum is designated as δ t _max, provide δ t below _maxapproximate calculation method.As shown in Figure 10, o _mp _m≈ R _m/ 2, R _mfor the moon radius of a ball, oo _m=384400km, thus:

$\&angle;p_m{\mathrm{oo}}_m=\frac{o_m{p}_m}{{\mathrm{oo}}_m}---\left(17\right)$

So: $\mathrm{\&delta;}{t}_{\max }=\frac{\&angle;p_m{\mathrm{oo}}_m}{{\mathrm{\&omega;}}_{\mathrm{rot}}}=8.9949e-005s$

Because scan axis direction is there is very complicated relation on the impact of lower edge on the moon, be difficult to describe by accurate mathematical model.According to scan geometry figure, the scan mechanism of approaching to reality can only be carried out to a certain extent to it.If θ _xfor scan axis and x _mangle, according to scanning patter, by approximate following for pulse daley model (see Figure 11): z _sewith x _m(or-x _m) when overlapping, δ t=0 (θ _x=0 (π)); z _seat y _mz _mwhen plane, δ t=δ t _max(θ _x=pi/2); Pulse daley and θ _xlinear.

Therefore the delay of pulse can represent with a simple linear model:

$\mathrm{\&delta;t}=\left\{\begin{array}{c}\frac{2{\mathrm{\&delta;t}}_{\max }}{\mathrm{\&pi;}}{\mathrm{\&theta;}}_x({\mathrm{\&theta;}}_x<\mathrm{\&pi;}/2)\\ \frac{2\mathrm{\&delta;}{t}_{\max }}{\mathrm{\&pi;}}(\mathrm{\&pi;}-{\mathrm{\&theta;}}_x){\mathrm{\&theta;}}_x\&GreaterEqual;\mathrm{\&pi;}/2\end{array}\right.---\left(18\right)$

(δ t < 0 when the moon at the first quarter, δ t > 0 when the moon at the last quarter)

Moon pulse simulation process flow diagram is shown in Figure 12.

Five, sun pulse simulation

As shown in figure 13, in t, the coordinate of solar direction vector under surving coordinate system can be tried to achieve by ephemeris and coordinate transform, thus two position angle δ _s, φ _sknown (see Figure 14).Work as GQ ₁₂with GS ₁when overlapping, this moment is exactly the pulse moment t that slit visual field 1 scans the sun _1sun; Work as GQ ₃₄with GS ₁when overlapping, this moment is exactly the pulse moment t that slit visual field 2 scans the sun _2sun.

Operation steps is as follows:

I). calculate the ground day vector of t, and the expression of geaster vector under inertial system;

Ii). utilize coordinate transform to calculate the expression of star day vector under sensor surving coordinate system, thus two position angle δ can be obtained _s, φ _s;

Iii). calculate σ _s=arcsincot (δ _s)/cos β

Iv). obtain the pulse moment by following two equations:

${\mathrm{\&phi;}}_s={\mathrm{\&omega;}}_{\mathrm{rot}}(\frac{t_{1\mathrm{sun}}+t_{2\mathrm{sun}}}{2}-t_{\mathrm{ref}})$ (19)

${\mathrm{\&sigma;}}_s=\frac{1}{2}({\mathrm{\&omega;}}_{\mathrm{rot}}(t_{2\mathrm{sun}}-t_{1\mathrm{sun}})-\mathrm{\&theta;})$

V). export pulse moment t _1sun, t _2sun

Sun pulse simulation process flow diagram is shown in Figure 15.

Six, emulation and error evaluation

The relevant physical parameter of ground day, integration moon sensor is described as follows: it is ω that day ground moon integration sensor scans rotating speed _rot=240r/min.Two semi-cone angle of circular cone Horizon sensor infrared horizon visual field circular cone are respectively γ ₁=38 ° and γ ₂=73 °.Scanning head is relative to M ₁m ₂lag angle be B _r=0 °, ground day, integration moon sensor surving coordinate system is γ relative to the setting angle of system _i=π/6, β _i=π/6, α _i=π/6.Fan-shaped life sensor is β relative to the pitch angle with scanning rotating shaft _s=16 °, slit visual field 1 is ahead of angular distance θ=4 ° of slit visual field 2 in ground day, integration moon sensor equatorial plane.

Initial six key elements of actual trajcctorics are: major semi-axis is 6878km, and orbital eccentricity is 0, and orbit inclination is 92 °, and right ascension of ascending node is π/6, and argument of perigee is 0 °, and very near angle is 0 °.

(1) pulse precision assessment

When the initial time of track is taken as 6 days 0 May in 2012 0 point 0 second.The phases of the moon is at this time full moon, and therefore moon pulse does not postpone.If do not consider compression of the earth, the earth impulse obtained is shown in Figure 16 and Figure 17.Sun pulse and moon pulse are shown in Figure 18 and Figure 19 respectively.

In order to verify the precision of pulses generation, owing to not considering compression of the earth when earth orientation is determined, so by not considering the earth impulse of compression of the earth and day, moon pulse calculates day ground angle, the moon ground angle compared with the day ground angle of reality, moon ground angle, the error angle obtained is shown in Figure 20 and Figure 21.As can be seen from the figure come, the precision of error angle reaches 10 ^-10degree, thus demonstrate the precision of pulse.

(2) ellipticity impact that orientation, the earth's core is determined

When the initial time of track is taken as 6 days 0 May in 2012 0 point 0 second.If consideration compression of the earth, the earth impulse obtained is shown in Figure 22 and Figure 23, and sun pulse and moon pulse are still as Figure 18 and Figure 19.

If calculate moon star ground, day star ground by pulse not carry out ellipticity correction time azimuthal, then the error result between itself and true bearing angle is shown in Figure 24 and Figure 25.As can be seen from Figure, if do not carry out ellipticity correction, the moon ground angle maximum error be 0.13 day ground angle maximum error be 0.12 °.

(3) impact determined of the moon relative moon heart orientation

The moon at the first quarter is similar with the principle of the delay of moon at the last quarter paired pulses, just there is positive and negative relation.Here the simulation result during moon at the first quarter is only provided.When the initial time of emulation is taken as 28 days 0 May in 2012 0 point 0 second.Moon pulse as shown in figure 26.The moon that pulse obtains ground angle with the real moon ground angle error angle curve see Figure 27, by figure the known moon ground angle maximum error angle be 0.12 degree, and moon ground angle with being greater than the real moon angle calculated.

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

Claims (1)

1., based on the measurement data analogy method of autonomous navigation system of day ground moon azimuth information, it is characterized in that realizing as follows:

(1) ground day, integration moon sensor reference pulse, reference pulse t _refprovided by following formula:

t _ref＝B _R/ω _rot+kT _s(1)

Wherein: B _rfor the plane of symmetry is relative to the drag angle of surving coordinate system X-axis, ω _rotfor scanning rotating speed, T _sfor the scan period, k is the scanning number of turns;

(2) ground day, integration moon sensor earth impulse, step is as follows:

The first step: the positional information r being obtained aircraft by dynamics of orbits model integration, under being projected in inertia is wherein subscript represent transposition, subscript I represents this variable drop under inertial system, and defines gesture mode, provides the attitude R of aircraft _iSE;

Second step: within a scan period, get any time t, in conjunction with attitude information, calculates the vector of unit length of this moment scanning sight line expression in inertial coordinates system

${\widehat{\mathrm{\&rho;}}}^I=\left[\begin{array}{c}\sin \left({\mathrm{\&gamma;}}_i\right)\cos ({\mathrm{\&omega;}}_{\mathrm{rot}}(t-t_{\mathrm{ref}})-B_{\mathrm{Ri}})\\ \sin \left({\mathrm{\&gamma;}}_i\right)\sin ({\mathrm{\&omega;}}_{\mathrm{rot}}(t-t_{\mathrm{ref}})-B_{\mathrm{Ri}})\\ \cos \left({\mathrm{\&gamma;}}_i\right)\end{array}\right]---\left(2\right)$

Wherein: γ _ifor the semi-cone angle of scanned cone, B _rifor scanning head is relative to the drag angle of sensor reference line, i=1, the 2 corresponding little vertebra of difference and augers;

3rd step: definition Fun=b ²-4ac also calculates the size of Fun, wherein:

$a={\widehat{\mathrm{\&rho;}}}_1^2+{\widehat{\mathrm{\&rho;}}}_2^2+{\widehat{\mathrm{\&rho;}}}_3^2/{(1-\mathrm{ee})}^2$

$b=2{\widehat{\mathrm{\&rho;}}}_{1}x+2{\widehat{\mathrm{\&rho;}}}_{2}y+2{\widehat{\mathrm{\&rho;}}}_{3}z/{(1-\mathrm{ee})}^2$

$c=x^2+y^2+z^2/{(1-\mathrm{ee})}^2-R_E^2$

In above formula: ee is compression of the earth, R _efor the equatorial radius of earth reference ellipsoid;

4th step: if Fun=0, then t is the pulse moment; Otherwise, then get the another one moment, repeat the first step to the 4th step, calculate corresponding Fun; By the Fun value in two moment, utilizing the time t that dichotomy finds Fun=0 corresponding, is also the pulse moment;

(3) ground day, integration moon sensor moon pulse simulation, step is as follows:

The first step: utilize ephemeris to differentiate the phases of the moon by current date;

Second step: the positional information r being obtained aircraft by dynamics of orbits model integration, under being projected in inertia is obtain the moon position under inertial system by ephemeris, and define gesture mode, provide the attitude R of aircraft _iSE;

3rd step: utilize coordinate transform to calculate the expression of star moon vector under sensor surving coordinate system, thus position angle φ can be obtained _mwith elevation angle δ _m;

4th step: calculate and do not consider the moon pulse t that the phases of the moon affects _1moonand t _2moon:

${\mathrm{\&phi;}}_m={\mathrm{\&omega;}}_{\mathrm{rot}}(\frac{t_{1\mathrm{moon}}+t_{2\mathrm{moon}}}{2}-t_{\mathrm{ref}})$ (3)

Wherein: σ _m=arcsin (tan δ _m/ cos β _s), β _sfor sensor slit is relative to the pitch angle of scanning ball meridian, for the centre of sphere angle of Slot bottom;

5th step: if the current phases of the moon is full moon, then the pulse that the 4th step obtains is the output moon pulse of sensor, if crescent or half moon, then need the basis of the pulse exported in the 4th step to be added delay δ t as output pulse, wherein δ t is that following formula provides:

$\mathrm{\&delta;t}=\left\{\begin{array}{cc}\frac{2\mathrm{\&delta;}{t}_{\max }}{\mathrm{\&pi;}}{\mathrm{\&theta;}}_x& ({\mathrm{\&theta;}}_x<\mathrm{\&pi;}/2)\\ \frac{2\mathrm{\&delta;}{t}_{\max }}{\mathrm{\&pi;}}(\mathrm{\&pi;}-{\mathrm{\&theta;}}_x)& ({\mathrm{\&theta;}}_x\&GreaterEqual;\mathrm{\&pi;}/2)\end{array}\right.---\left(4\right)$

Wherein: θ _xfor scan axis and x _mthe angle of namely vertical with life direction in normal society plane axle, δ t _maxfor maximum-delay amount; δ t < 0 when the moon at the first quarter, δ t > 0 when the moon at the last quarter;

(4) ground day, integration moon sensor sun pulse simulation, step is as follows:

The first step: the positional information r being obtained aircraft by dynamics of orbits model integration, under being projected in inertia is obtain the moon position under inertial system by ephemeris, and define gesture mode, provide the attitude R of aircraft _iSE;

Second step: utilize coordinate transform to calculate the expression of star day vector under sensor surving coordinate system, thus the position angle φ of the sun can be obtained _swith elevation angle δ _s;

3rd step: calculate sun pulse t _1sunand t _2sun:

${\mathrm{\&phi;}}_s={\mathrm{\&omega;}}_{\mathrm{rot}}(\frac{t_{1\mathrm{sun}}+t_{2\mathrm{sun}}}{2}-t_{\mathrm{ref}})$ (5)

Wherein: σ _s=arcsin (tan δ _s/ cos β _s), β _sfor sensor slit is relative to the pitch angle of scanning ball meridian, for the centre of sphere angle of Slot bottom;

Calculate ground day, the integration moon reference pulse of sensor, earth impulse, moon pulse and sun pulse by above-mentioned, thus complete the simulation of ground day, integration moon sensor pulse information.