CN103808330A - Satellite ephemeris calculation method for star sensor field experiment - Google Patents

Satellite ephemeris calculation method for star sensor field experiment Download PDF

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Publication number
CN103808330A
CN103808330A CN201210439239.XA CN201210439239A CN103808330A CN 103808330 A CN103808330 A CN 103808330A CN 201210439239 A CN201210439239 A CN 201210439239A CN 103808330 A CN103808330 A CN 103808330A
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calculate
test
moment
satellite
ascending node
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CN103808330B (en
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张小伟
袁荣钢
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SHANGHAI AEROSPACE CONTROL ENGINEERING INSTITUTE
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SHANGHAI AEROSPACE CONTROL ENGINEERING INSTITUTE
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    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C25/00—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/02—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by astronomical means

Abstract

The invention discloses a satellite ephemeris calculation method for star sensor field experiment. The method comprises the following steps: determining and calculating input and output quantity parameters; and calculating output quantity by using the input quantity, according to formula of spherical triangle and astronomical ephemeris calculation. Compared with the prior art, the invention has the beneficial effects of fixing the problem of inconsistency of satellite orbit ephemeris in ground stationary condition and orbit flight condition, and realizing the simulation of real flight track of satellite in experiment. The method is verified to have correct results and good effect, and can be used in star sensor field experiment, and system grade outfield experiment of sun sensor and other optical parts.

Description

A kind of star sensor field trial satellite ephemeris computing method
Technical field
The present invention relates to a kind of star sensor field trial technology, especially the track ephemeris computation method of satellite in the time carrying out the field trial of star sensor system level.
Background technology
Be mainly product-level for the field trial of star sensor at present, star sensor is connected with spaceborne computer, the system-level test in outfield of its access satellite attitude control system closed loop was not also carried out.But for fully installation polarity, serviceability and the system software culminant star sensor data processing of checking star sensor in satellite attitude control system and the correctness of attitude algorithm, it is necessary carrying out the system-level test in star sensor outfield.
System-level test is exactly the field trial place at broad view, select fine night, mounting means on analog satellite, star sensor is arranged on turntable, and be connected with spaceborne computer, allow star sensor be operated in and approximate in orbit operating mode, output inertia attitude data is to spaceborne computer, and spaceborne computer calculates the track profile of satellite again in conjunction with track almanac data according to inertia attitude data.
The problem existing in test is at present: the situation of the ephemeris of satellite and orbit parameter and true flight is in-orbit inconsistent, test location longitude and latitude is exactly substar longitude and latitude corresponding to test moment satellite orbital position, it is definite known quantity, the semi-major axis of track, excentricity, orbit inclination is got satellite actual value in-orbit, these several parameters are normal values for circular orbit satellite, above-mentioned parameter can be used as constant and is arranged in advance in spaceborne computer software, and track argument, right ascension of ascending node, when test, on the star of spaceborne computer, the parameter such as clock can not directly obtain, therefore can not real simulation situation in-orbit.
Summary of the invention
The technical problem to be solved in the present invention is to provide a kind of star sensor field trial satellite ephemeris computing method.It can meet star sensor field trial needs.
For solving the problems of the technologies described above, a kind of star sensor field trial satellite ephemeris computing method of the present invention, are to utilize known inclination of satellite orbit i, test location latitude
Figure 201210439239X100002DEST_PATH_IMAGE001
and longitude
Figure 804947DEST_PATH_IMAGE002
, the method for the track argument that corresponding time on-test in local moment of calculating ascending node of orbit is spaceborne computer clock, satellite position in track segmental arc and the right ascension of ascending node of reflection orbital plane and equatorial plane relative position.
Described star sensor field trial satellite ephemeris computing method, it comprises the steps:
1) determine and calculate input and output amount parameter
Determine test location longitude
Figure 476100DEST_PATH_IMAGE002
, latitude
Figure 774226DEST_PATH_IMAGE001
, inclination of satellite orbit i, ascending node local time t d as input quantity, moment on-test t 0 and track argument corresponding to moment u 0 , right ascension of ascending node for output quantity;
2) calculate test moment track argument u 0
Utilize orbit inclination ifirst calculate orbital plane and equatorial plane angle, utilize its angle, location latitude
Figure 470393DEST_PATH_IMAGE001
, calculate test moment track argument according to spherical triangle formula u 0 ;
3) calculate moment on-test t 0
Utilize orbit inclination ifirst calculate orbital plane and equatorial plane angle, recycle its angle and calculate ascending node and meridian ellipse angle according to spherical triangle formula, utilize its angle and ascending node local time t d , longitude
Figure 222897DEST_PATH_IMAGE002
calculate moment on-test t 0 ;
4) calculate right ascension of ascending node
Figure 996818DEST_PATH_IMAGE003
Utilization calculates t 0 calculate Julian date
Figure 924323DEST_PATH_IMAGE004
, and then calculate t 0 moment Greenwich sidereal time angle
Figure 201210439239X100002DEST_PATH_IMAGE005
, utilize t 0 , t d ,
Figure 629891DEST_PATH_IMAGE005
calculate right ascension of ascending node
Figure 275636DEST_PATH_IMAGE003
.
The method that the present invention adopts, compared with prior art, its advantage and beneficial effect are: the present invention proposes a kind of general star sensor outfield system-level test satellite ephemeris computation method, solved satellite under the static operating mode in ground with the inconsistent problem of flight track ephemeris in-orbit, realized the simulation of process of the test Satellite Live Flying track.The method empirical tests result is correct, respond well, except can be applicable to, star sensor field trial, also to can be applicable to the system-level field trial of other optical measurement such as sun sensor.
 
Accompanying drawing explanation
Below with reference to drawings and Examples, the invention will be further described.
Satellite orbital position and attitude embodiment schematic diagram when Fig. 1 is test.
embodiment
As shown in Figure 1, satellite orbital position and attitude schematic diagram during for test, in figure, symbol implication is as follows:
Figure 587669DEST_PATH_IMAGE006
for the earth's core;
n, srepresent respectively earth north, the South Pole;
cfor ascending node;
Place face is earth equatorial plane;
Figure DEST_PATH_IMAGE007
for test location warp;
for ascending node place warp;
Figure DEST_PATH_IMAGE009
place face is satellite orbit plane;
for orbital plane with
Figure 641841DEST_PATH_IMAGE007
intersection point, tests location;
Figure DEST_PATH_IMAGE011
for the celestial body coordinate system in zero attitude situation.
Known input quantity is:
Orbit inclination;
bpoint longitude and latitude
Figure 695247DEST_PATH_IMAGE002
,
Figure 659661DEST_PATH_IMAGE001
;
Ascending node of orbit local time t d .
Output quantity to be asked is:
Track argument u 0 ;
Moment on-test t 0 ;
Right ascension of ascending node
Figure 140321DEST_PATH_IMAGE003
.
Computation process is as follows:
From trirectangular spherical triangle
Figure 432762DEST_PATH_IMAGE012
,
Figure 913029DEST_PATH_IMAGE014
, then by spherical triangle formula
Figure DEST_PATH_IMAGE015
Can obtain
So obtain track argument u 0 :
Again by spherical triangle formula
Figure 383510DEST_PATH_IMAGE018
Can obtain
Beijing time, corresponding longitude was
Figure 992608DEST_PATH_IMAGE020
While test, should be when Beijing
t 0 = t d +
Figure DEST_PATH_IMAGE021
If the date is respectively when test year, month, day, first calculate Julian date
Figure 450135DEST_PATH_IMAGE004
Figure 998928DEST_PATH_IMAGE022
Figure DEST_PATH_IMAGE023
Greenwich sidereal time angle is:
Right ascension of ascending node
Figure 416319DEST_PATH_IMAGE003
for:
Figure DEST_PATH_IMAGE025
Spaceborne computer carries out Orbit simulation computing by the parameter result that said method calculates gained, operation result can keep with satellite the Live Flying track when the overtesting location consistent, analog satellite situation in orbit more truly, for the demonstration test of star sensor provides operating condition of test more true to nature.

Claims (2)

1. star sensor field trial satellite ephemeris computing method, utilize known inclination of satellite orbit i, test location latitude
Figure 183926DEST_PATH_IMAGE001
and longitude , it is characterized in that: be the calculating of spaceborne computer clock, the track argument of reflection satellite position in track segmental arc and the right ascension of ascending node of reflection orbital plane and equatorial plane relative position to corresponding time on-test in the local moment of ascending node of orbit.
2. star sensor field trial satellite ephemeris computing method according to claim 1, is characterized in that comprising the steps:
1) determine and calculate input and output amount parameter
Determine test location longitude
Figure 943120DEST_PATH_IMAGE002
, latitude
Figure 338330DEST_PATH_IMAGE001
, inclination of satellite orbit i, ascending node local time t d as input quantity, moment on-test t 0 and track argument corresponding to moment u 0 , right ascension of ascending node
Figure 464286DEST_PATH_IMAGE003
for output quantity;
2) calculate test moment track argument u 0
Utilize orbit inclination ifirst calculate orbital plane and equatorial plane angle, utilize its angle, location latitude
Figure 286749DEST_PATH_IMAGE001
, calculate test moment track argument according to spherical triangle formula u 0 ;
3) calculate moment on-test t 0
Utilize orbit inclination ifirst calculate orbital plane and equatorial plane angle, recycle its angle and
Figure 616099DEST_PATH_IMAGE001
calculate ascending node and meridian ellipse angle according to spherical triangle formula, utilize its angle and ascending node local time t d , longitude
Figure 628048DEST_PATH_IMAGE002
calculate moment on-test t 0 ;
4) calculate right ascension of ascending node
Figure 296927DEST_PATH_IMAGE003
Utilization calculates t 0 calculate Julian date
Figure 352608DEST_PATH_IMAGE004
, and then calculate t 0 moment Greenwich sidereal time angle
Figure 106937DEST_PATH_IMAGE005
, utilize t 0 , t d ,
Figure 152604DEST_PATH_IMAGE005
calculate right ascension of ascending node
Figure 675989DEST_PATH_IMAGE003
.
CN201210439239.XA 2012-11-07 2012-11-07 A kind of star sensor field trial satellite ephemeris computational methods Active CN103808330B (en)

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104965208A (en) * 2015-04-01 2015-10-07 北京理工雷科电子信息技术有限公司 Ephemeris for simulation of satellite navigation signals, and almanac obtaining method
CN108710379A (en) * 2018-06-14 2018-10-26 上海卫星工程研究所 Fixed statellite is imaged Yaw steering angle computational methods
CN113532423A (en) * 2021-07-22 2021-10-22 哈尔滨工业大学 Astronomical platform autonomous positioning method based on tilt angle sensor

Citations (2)

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Publication number Priority date Publication date Assignee Title
KR20060076548A (en) * 2004-12-29 2006-07-04 한국항공우주연구원 Satellite magnetometer bias correction method using orbit geometry
CN101082494A (en) * 2007-06-19 2007-12-05 北京航空航天大学 Self boundary marking method based on forecast filtering and UPF spacecraft shading device

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Publication number Priority date Publication date Assignee Title
KR20060076548A (en) * 2004-12-29 2006-07-04 한국항공우주연구원 Satellite magnetometer bias correction method using orbit geometry
CN101082494A (en) * 2007-06-19 2007-12-05 北京航空航天大学 Self boundary marking method based on forecast filtering and UPF spacecraft shading device

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Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104965208A (en) * 2015-04-01 2015-10-07 北京理工雷科电子信息技术有限公司 Ephemeris for simulation of satellite navigation signals, and almanac obtaining method
CN108710379A (en) * 2018-06-14 2018-10-26 上海卫星工程研究所 Fixed statellite is imaged Yaw steering angle computational methods
CN108710379B (en) * 2018-06-14 2021-02-02 上海卫星工程研究所 Method for calculating yaw guide angle in imaging of stationary satellite
CN113532423A (en) * 2021-07-22 2021-10-22 哈尔滨工业大学 Astronomical platform autonomous positioning method based on tilt angle sensor

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