CN102927983B - Method for protecting earth sensor probe under GEO (Geosynchronous Orbit) satellite pitch offset - Google Patents

Abstract

The invention discloses a method for protecting an earth sensor probe under GEO (Geosynchronous Orbit) satellite pitch offset, which comprises the following steps of: (1) determining the offset quantity of a satellite pitch attitude; (2) determining the protection parameters of the earth sensor probe; (3) determining the components of a solar unit vector S and a moony center unit vector M in a star coordinate system at current time; (4) computing the included angle theta SM from the moon to a satellite to the sun, the included angle theta ME of the moon to the satellite to the earth and the included angle theta SE from the sun to the satellite to the earth; (5) respectively determining the protection strategies of the earth sensor probe through the disturbance of the moon and the sun on the earth sensor probe; and (6) prohibiting the corresponding earth sensor probe from participating in attitude computation according to the protected probe determined from the step (5). The method disclosed by the invention can be used for determining the protected earth sensor probe of the satellite under the condition that the satellite pitch attitude is offset for a long time.

Description

The biased lower earth sensor probe guard method of a kind of GEO satellite pitching

Technical field

The present invention relates to the biased lower earth sensor probe guard method of a kind of GEO satellite pitching.

Background technology

Geostationary orbit (GEO) satellite distribution is at earth equatorial plane, and orbit altitude is about 36000km, and satellite uses roll attitude and the pitch attitude of swing type infrared earth sensor instrumented satellite.Swing type infrared earth sensor has four probes, scans earth edge respectively, determines the attitude of satellite by calculating.Before and after spring and fall equinoxes turnover ground shadow, the sun enters infrared earth sensor scanning probe visual field, or the moon enters earth sensor probe scanning field of view, both all can disturb infrared earth sensor to pop one's head in, and the attitude of infrared earth sensor is exported and produces error.

The probe guard method of existing satellite is:

1. satellite is when rolling, pitch attitude are zero, arranges the protection zone of infrared earth sensor four probe.

2. ground or star pass through to calculate, draw earth sensor a certain probe protection start time, guard time length and guard time interval.

3. implement earth sensor probe guard period at needs, satellite independently can carry out probe protection, or realizes the protection of infrared earth sensor probe by the method for ground transmission telecommand.

There is some deficiency following in said method:

The setting prerequisite of infrared earth probe protection zone is the rolling of satellite, pitch attitude is zero; when Satellite Payloads needs to make pitch attitude carry out long-term being biased; the probe protection zone of infrared earth sensor there occurs change, and original computing method are no longer applicable.

Summary of the invention

Technology of the present invention is dealt with problems and is: for the deficiencies in the prior art; provide the biased lower earth sensor probe guard method of a kind of GEO satellite pitching; the situation adopting the present invention can be biased for a long time in satellite pitch attitude, determines the protected earth sensor probe of satellite.

Technical solution of the present invention is:

The biased lower earth sensor probe guard method of a kind of GEO satellite pitching, comprises the following steps:

(1) amount of bias of satellite pitch attitude is determined;

(2) determine earth sensor probe Protection parameters, described earth sensor probe Protection parameters f, a, b, c, d are determined by following formula,

f＝ρecos45°，a＝f+g+dM，b＝f-dM，c＝f+dM，d＝f-g-dM；

Wherein, ρ e is earth apparent radius, and dM is earth sensor probe protection threshold value, value 3 °, and g is earth sensor probe sweep limit, value 5.3;

(3) determine that sun unit vector S and moon central element vector M is at the component S of celestial body coordinate system under current time _x, S _y, S _z, M _x, M _y, M _z;

(4) angle theta of the moon-satellite-sun is calculated _sM, the moon-satellite-earth angle theta _mE, the sun-satellite-earth angle theta _sE;

(5) earth sensor probe Preservation tactics is determined in the moon and the interference of the sun to earth sensor probe respectively, wherein, to the interference of earth sensor probe, step (51) is then proceeded to for the moon, then proceed to step (52) for the sun to the interference of earth sensor probe, described earth sensor probe comprises the first probe, the second probe, the 3rd probe, the 4th probe;

(51) if θ _sM< 120 °, or θ _sMwhen>=120 °, and θ _mE< ρ e-dE or θ _mE> ρ a+1.4* (dM+ θ), or θ _sMwhen>=120 °, and | M _x| > sin (a) or | M _x| < sin (d) or | M _y| < sin (b) or | M _y| > sin (c), then the moon is noiseless to earth sensor probe, proceeds to step (6), otherwise, determine protected probe according to following situation and proceed to step (6):

If M _x< 0 and M _y< 0, then need protection first to pop one's head in;

If M _x< 0 and M _y> 0, then need protection second to pop one's head in;

If M _x> 0 and M _y< 0, then need protection the 3rd probe;

If M _x> 0 and M _y> 0, then need protection the 4th probe;

Wherein, dE is earth sensor probe protection threshold value, value 2 °; θ is satellite pitch attitude angle of eccentricity absolute value;

(52) if θ _sE< ρ e-dE or θ _sE> ρ a+1.4* (dM+ θ), or | S _x| > sin (a) or | S _x| < sin (d) or | S _y| < sin (b) or | S _y| > sin (c), then the sun is noiseless to earth sensor probe, proceeds to step (6), otherwise, determine protected probe according to following situation and proceed to step (6):

If S _x< 0 and S _y< 0, then need protection first to pop one's head in;

If S _x< 0 and S _y> 0, then need protection second to pop one's head in;

If S _x> 0 and S _y< 0, then need protection the 3rd probe;

If S _x> 0 and S _y> 0, then need protection the 4th probe;

(6) according to the protection probe determined in step (5), forbid that corresponding earth sensor probe participates in Attitude Calculation.

The present invention compared with prior art tool has the following advantages:

The present invention overcomes the deficiency of original infrared earth sensor probe guard method; provide the method for infrared earth sensor probe protection under a kind of GEO satellite pitching bias conditions, overcome existing method non-serviceable problem when satellite pitch attitude is biased.Further, the present invention can be different according to the angle that satellite pitch attitude is biased, calculated the parameter of probe protection by amendment, realize the method for the infrared earth sensor probe protection of satellite in different pitching angle of eccentricity situations.

Accompanying drawing explanation

Fig. 1 is process flow diagram of the present invention;

Embodiment

Just by reference to the accompanying drawings the present invention is described further below.

As described in Figure 1, be process flow diagram of the present invention, detailed process is as follows:

(1) amount of bias of satellite pitch attitude is determined, pitching angle of eccentricity absolute value representation when this amount of bias adopts the long-term pitching of satellite to be biased.

(2) under above-mentioned bias conditions, determine earth sensor probe Protection parameters, this Protection parameters comprises f, a, b, c, d, and each parameter adopts following formula to determine respectively.

f＝ρecos45°，a＝f+g+dM，b＝f-dM，c＝f+dM，d＝f-g-dM；

Wherein, ρ e is earth apparent radius; 45 ° is earth sensor probe scanning latitude; DM is probe protection threshold value, and value 3 ° of effects expand protection domain; G is earth sensor probe sweep limit, value 5.3.

Infrared earth sensor probe, along pitch orientation scanning, has two kinds of scan modes: wide scanning and narrow scan.Generally, transfer orbit runs or catches terrestrial time and uses wide scan mode, uses narrow scan mode when geostationary orbit.If infrared earth sensor is operated in wide scan mode; there will be the situation that infrared earth sensor two probe is subject to the sun or moon interference simultaneously; infrared earth sensor cannot normally work when two probes are disturbed simultaneously; so should make quickly to be operated in narrow scan mode at infrared earth sensor probe guard period; under GEO satellite infrared earth sensor narrow scan working method, scanning field of view is 5.3 degree, and biased 3 degree of satellite pitching does not exceed earth sensor visual field.

(3) need the moment determining to protect earth sensor probe, sun unit vector S and moon central element vector M is at the component of orbital coordinate system.

Solar direction is at the component S of orbital coordinate system

$\left[\begin{array}{c}{S}_{\mathrm{ox}}\\ S_{\mathrm{oy}}\\ S_{\mathrm{oz}}\end{array}\right]=\mathrm{Aou}*\mathrm{Aoii}*\left[\begin{array}{ccc}0& 1& 0\\ 0& 0& -1\\ -1& 0& 0\end{array}\right]\left[\begin{array}{c}\mathrm{Xs}\\ \mathrm{Ys}\\ \mathrm{Zs}\end{array}\right]$

Moon unit vector is at the component M of orbital coordinate system

$\left[\begin{array}{c}{M}_{\mathrm{ox}}\\ M_{\mathrm{oy}}\\ M_{\mathrm{oz}}\end{array}\right]=\mathrm{Aou}*\mathrm{Aoii}*\left[\begin{array}{ccc}0& 1& 0\\ 0& 0& -1\\ -1& 0& 0\end{array}\right]\left[\begin{array}{c}\mathrm{Xm}\\ \mathrm{Ym}\\ \mathrm{Zm}\end{array}\right]$

Wherein: transition matrix:

$\mathrm{Aou}=\left[\begin{array}{ccc}\cos \left(u\right)& 0& \sin \left(u\right)\\ 0& 1& 0\\ -\sin \left(u\right)& 0& \cos \left(u\right)\end{array}\right];$ $\mathrm{Aoii}=\left[\begin{array}{ccc}\cos \left(i\right)\cos \left(\mathrm{\&Omega;}\right)& -\sin \left(i\right)& \cos \left(i\right)\sin \left(\mathrm{\&Omega;}\right)\\ \sin \left(i\right)\cos \left(\mathrm{\&Omega;}\right)& \cos \left(i\right)& \sin \left(i\right)\sin \left(\mathrm{\&Omega;}\right)\\ -\sin \mathrm{\&Omega;}& 0& \cos \left(\mathrm{\&Omega;}\right)\end{array}\right]$

I is inclination of satellite orbit, unit: rad;

Ω is right ascension of satellite ascending node, unit: rad;

U is satellite t track argument, unit: rad;

Xs, Ys, Zs are that sun unit vector is at inertial coordinates system component;

Xm, Ym, Zm are that moon unit vector is at inertial coordinates system component.

The nominal attitude of GEO satellite operation track normal mode is: θ=θ b, φ=φ b, ψ=ψ b=0 °

Under nominal attitude, earth unit vector E, moon unit vector M, sun unit vector S can be expressed as at the component of celestial body coordinate system:

$\left[\begin{array}{c}{E}_x\\ E_y\\ E_z\end{array}\right]=A_{\mathrm{bo}}\&CenterDot;\left[\begin{array}{c}{E}_{\mathrm{ox}}\\ E_{\mathrm{oy}}\\ E_{\mathrm{oz}}\end{array}\right],$ $\left[\begin{array}{c}{M}_x\\ M_y\\ M_z\end{array}\right]=A_{\mathrm{bo}}\&CenterDot;\left[\begin{array}{c}{M}_{\mathrm{ox}}\\ M_{\mathrm{oy}}\\ M_{\mathrm{oz}}\end{array}\right]\left[\begin{array}{c}{S}_x\\ S_y\\ S_z\end{array}\right]=A_{\mathrm{bo}}\&CenterDot;\left[\begin{array}{c}{S}_{\mathrm{ox}}\\ S_{\mathrm{oy}}\\ S_{\mathrm{oz}}\end{array}\right]$

Wherein: transition matrix

$A_{\mathrm{bo}}=\left[\begin{array}{ccc}1& 0& 0\\ 0& \cos {\mathrm{\&phi;}}_b& \sin {\mathrm{\&phi;}}_b\\ 0& -\sin {\mathrm{\&phi;}}_b& \cos {\mathrm{\&phi;}}_b\end{array}\right]\left[\begin{array}{ccc}\cos {\mathrm{\&theta;}}_b& 0& -\sin {\mathrm{\&theta;}}_b\\ 0& 1& 0\\ \sin {\mathrm{\&theta;}}_b& 0& \cos {\mathrm{\&theta;}}_b\end{array}\right]$

(4) angle theta of the moon-satellite-sun is calculated _sM, the moon-satellite-earth angle theta _mE, the sun-satellite-earth angle theta _sE.

θ _SM＝arccos(M·S)；θ _ME＝arccos(M·E)；θ _SE＝arccos(S·E)。

In above formula, " " represents dot product; E is the component of earth unit vector in orbital coordinate system, is expressed as

$\left[\begin{array}{c}{E}_{\mathrm{ox}}\\ E_{\mathrm{oy}}\\ E_{\mathrm{oz}}\end{array}\right]=\left[\begin{array}{c}0\\ 0\\ 1\end{array}\right].$

(5) earth sensor probe Preservation tactics is determined in the moon and the interference of the sun to earth sensor probe respectively; wherein, in step (51) and step (52), independent judgment is carried out to the interference of earth sensor probe and the sun respectively to the interference of earth sensor probe for the moon.First probe, the second probe, the 3rd probe, the 4th probe are comprised for earth sensor probe described in GEO satellite

The interference of the moon to earth sensor probe adopts following process:

If θ _sM< 120 °, or θ _sMwhen>=120 °, and θ _mE< ρ e-dE or θ _mE> ρ a+1.4* (dM+ θ), or θ _sMwhen>=120 °, and | M _x| > sin (a) or | M _x| < sin (d) or | M _y| < sin (b) or | M _y| > sin (c), then the moon is noiseless to earth sensor probe.Otherwise, select to need to close to carry out the probe protected according to following situations:

If M _x< 0 and M _y< 0, then need protection first to pop one's head in;

If M _x< 0 and M _y> 0, then need protection second to pop one's head in;

If M _x> 0 and M _y< 0, then need protection the 3rd probe;

If M _x> 0 and M _y> 0, then need protection the 4th probe;

Wherein, dE is earth sensor probe protection threshold value, and effect expands protection domain, value 2 °; θ is for representing satellite pitch attitude angle of eccentricity absolute value;

The interference of the sun to earth sensor probe adopts following process:

If θ _sE< ρ e-dE or θ _sE> ρ a+1.4* (dM+ θ), or | S _x| > sin (a) or | S _x| < sin (d) or | S _y| < sin (b) or | S _y| > sin (c), then the sun is noiseless to earth sensor probe.Otherwise, select to need to close to carry out the probe protected according to following situation:

If S _x< 0 and S _y< 0, then need protection first to pop one's head in;

If S _x< 0 and S _y> 0, then need protection second to pop one's head in;

If S _x> 0 and S _y< 0, then need protection the 3rd probe;

If S _x> 0 and S _y> 0, then need protection the 4th probe;

(6) after determined the probe needing to protect according to step (5), the interference of the comprehensive moon and the sun, closes and needs protected probe.

GEO satellite orbit period is about 24 hours, therefore pitch attitude is biased θ often increases+1 °, and ground quick probe protection start time, end time postpone about: (satellite orbit period * 3600)/360 seconds, namely 240 seconds.In like manner, pitch attitude is biased often increases-1 °, ground quick probe protection start time, about 240 seconds in advance end time.

The unspecified part of the present invention belongs to general knowledge as well known to those skilled in the art.

Claims (1)

1. the biased lower earth sensor probe guard method of GEO satellite pitch attitude, its feature is comprising the following steps:

(1) amount of bias of satellite pitch attitude is determined;

(2) determine earth sensor probe Protection parameters, described earth sensor probe Protection parameters f, a, b, c, d are determined by following formula,

f＝ρecos45°，a＝f+g+dM，b＝f-dM，c＝f+dM，d＝f-g-dM；

Wherein, ρ e is earth apparent radius, and dM is earth sensor probe protection threshold value, value 3 °, and g is earth sensor probe sweep limit, value 5.3;

(3) determine that sun unit vector S and moon central element vector M is at the component S of celestial body coordinate system under current time _x, S _y, S _z, M _x, M _y, M _z;

(4) angle theta of the moon-satellite-sun is calculated _sM, the moon-satellite-earth angle theta _mE, the sun-satellite-earth angle theta _sE;

(5) earth sensor probe Preservation tactics is determined in the moon and the interference of the sun to earth sensor probe respectively, wherein, to the interference of earth sensor probe, step (51) is then proceeded to for the moon, then proceed to step (52) for the sun to the interference of earth sensor probe, described earth sensor probe comprises the first probe, the second probe, the 3rd probe, the 4th probe;

(51) if θ _sM< 120 °, or θ _sMwhen>=120 °, and θ _mE< ρ e-dE or θ _mE> ρ a+1.4* (dM+ θ), or θ _sMwhen>=120 °, and | M _x| > sin (a) or | M _x| < sin (d) or | M _y| < sin (b) or | M _y| > sin (c), then the moon is noiseless to earth sensor probe, proceeds to step (6), otherwise, determine protected probe according to following situation and proceed to step (6):

If M _x< 0 and M _y< 0, then need protection first to pop one's head in;

If M _x< 0 and M _y> 0, then need protection second to pop one's head in;

If M _x> 0 and M _y< 0, then need protection the 3rd probe;

If M _x> 0 and M _y> 0, then need protection the 4th probe;

Wherein, dE is earth sensor probe protection threshold value, value 2 °; θ is for representing satellite pitch attitude angle of eccentricity absolute value;

(52) if θ _sE< ρ e-dE or θ _sE> ρ a+1.4* (dM+ θ), or | S _x| > sin (a) or | S _x| < sin (d) or | S _y| < sin (b) or | S _y| > sin (c), then the sun is noiseless to earth sensor probe, proceeds to step (6), otherwise, determine protected probe according to following situation and proceed to step (6):

If S _x< 0 and S _y< 0, then need protection first to pop one's head in;

If S _x< 0 and S _y> 0, then need protection second to pop one's head in;

If S _x> 0 and S _y< 0, then need protection the 3rd probe;

If S _x> 0 and S _y> 0, then need protection the 4th probe;

(6) according to the protection probe determined in step (5), forbid that corresponding earth sensor probe participates in Attitude Calculation.