CN102081156A - Driving mechanism for synthetic aperture radar satellite - Google Patents
Driving mechanism for synthetic aperture radar satellite Download PDFInfo
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- CN102081156A CN102081156A CN2009101996748A CN200910199674A CN102081156A CN 102081156 A CN102081156 A CN 102081156A CN 2009101996748 A CN2009101996748 A CN 2009101996748A CN 200910199674 A CN200910199674 A CN 200910199674A CN 102081156 A CN102081156 A CN 102081156A
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- satellite
- driving mechanism
- synthetic aperture
- aperture radar
- angle
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Abstract
The present invention relates to a solar cell array driving mechanism oblique mounting technology, belongs to the field of radar satellite configuration design, and concretely discloses a driving mechanism for a synthetic aperture radar satellite. Two driving mechanisms are respectively installed on a positive and a negative Y separator plate of the satellite body and obliquely installed at an angle of 15 degrees together with the surfaces with solar cells bonded thereon, and the solar cell arrays of the two driving mechanisms are installed in an antisymmetric manner. The driving mechanism for synthetic aperture radar satellite of the invention is obliquely installed by an angle according to the angle of oblique flying of the satellite, thereby reducing the influence on the satellite control precision caused by the interference moment, wherein the interference moment is generated by that the solar cell arrays tracks and toward the sun, and enlarging the effective illuminating area of the solar battery arrays.
Description
Technical field
The present invention relates to a kind of driving mechanism of Synthetic Aperture Radar satellite.
Background technology
So-and-so satellite is first X-band Synthetic Aperture Radar satellite of China, has two driving mechanisms, be installed in respectively satellite ± the Y dividing plate on.Because the antenna as satellite load needs tiltedly side-looking over the ground, and antenna is a formal dress, therefore needs satellite body tiltedly to fly here and guarantees.In order to guarantee the energy supply of whole star, simultaneously in order to reduce solar battery array to following the tracks of of the influence of the directed disturbance torque that produces day to the satellite control accuracy, just need so that correct the angle that satellite tiltedly flies, make solar battery array lighting best driving mechanism angle mount on satellite.Yet because the reference for installation (the solar battery array plain flange face of driving mechanism and the intersection point of driving mechanism axis) of sun battle array is virtual, can't directly measure and determine, therefore can't on installing plate, directly locate, hole, install with traditional way, even loaded onto reluctantly, also to face the problem of frequent dismounting, big and the repeatable accuracy of surveying work amount is difficult to guarantee, also can't effectively control for the erection stress between solar battery array and the driving mechanism.
Summary of the invention
At the above-mentioned deficiency of prior art, technical matters to be solved by this invention provides a kind of driving mechanism of Synthetic Aperture Radar satellite.The present invention tiltedly flies angle according to satellite, with driving mechanism angle mount certain angle, has reduced solar battery array to following the tracks of the influence of the directed disturbance torque that produces to the satellite control accuracy day, has enlarged the effective illuminating area of solar battery array.
In order to solve the problems of the technologies described above, the driving mechanism of a kind of Synthetic Aperture Radar satellite provided by the invention, its innovative point be, two driving mechanisms be installed in respectively the satellite celestial body ± the Y dividing plate on, and and solar cell piece paster face be 15 ° of angle mounts together; The solar battery array antisymmetry of two driving mechanisms is installed.
Among the present invention, the installation site of driving mechanism on ± Y dividing plate is by locating piece and frock is installed determines, locus with the three-coordinate instrument datum mark is installed behind the frock dress star, locating piece is close to behind the frock dress star dress star locating device as driving mechanism.
The present invention can reduce solar battery array to following the tracks of the influence of the directed disturbance torque that produces to the satellite control accuracy day, has enlarged satellite effective illuminating area of solar battery array in orbit the time, reduces the illumination angle, improves light conditions; In addition, special tooling cooperates the use of locating piece, has simplified the work of precision measure greatly, makes easy accessibility, has avoided the work that repeats.
Description of drawings
Fig. 1 is the principle schematic of driving mechanism of the present invention.
Fig. 2 is the installation design sketch of the present invention on satellite.
Fig. 3 is a special tooling engineering drawing of the present invention.
Fig. 4 be the present invention on satellite at rail flight synoptic diagram (heading be vertical paper inwards).
Wherein: 1 is-the Y driving mechanism; 2 are+the Y driving mechanism; 3,4 is solar cell piece paster face; 10 is the satellite celestial body; 20 is solar battery array.
Embodiment
Below in conjunction with accompanying drawing and specific embodiment the present invention is described in further detail:
As shown in Figure 1, 2, satellite useful load (antenna) requires tiltedly side-looking over the ground, correspondingly except requiring to the angle mount of floor (above infrared earth sensor, observing and controlling antenna are installed, number passes antennas etc.), in order to reduce the influence that 20 pairs of solar battery arrays are followed the tracks of the directed disturbance torque that produces to the satellite control accuracy day as much as possible, guarantee the illumination of solar battery array 20 to greatest extent, take all factors into consideration factors such as satellite transit track, driving mechanism 1,2 is done 15 ° of designs of angle mount.According to solar battery array 20 designing requirements, the one side of pasting battery sheet 3,4 when rounding state must be towards celestial body 10 outsides.Consider the situation of 3 blocks of plates, after solar battery array 20 launches, if make both wings battery sheet towards unanimity, both wings sun battle array symmetry is installed, but, can bring the inconsistent of sun battle array root hinge start point like this, both wings solar battery array duration of run is asynchronous, thereby control exerts an influence to the attitude of satellite.Therefore, select both wings solar battery array 20 antisymmetry that (as shown in Figure 2) is installed, farthest guarantee the synchronism of both wings solar battery array 20 durations of run.When solar battery array 20 launches a wing down, a wing upwards at first makes zero before therefore needing to follow the tracks of 1,2 pairs of days of driving mechanism, makes its both wings in the same way.
As shown in Figure 3, reference point (cross line place) after the new conversion that the use special tooling provides, can determine driving mechanism and the solar battery array mutual relationship between compressing a little accurately and easily, and its rear end that is installed on the celestial body is identical with real driving mechanism physical dimension, after utilizing it to determine the position of locating piece, just locating piece can be permanently affixed on the celestial body plate, take off special tooling and can driving mechanism be installed according to locating piece.After this, state is fixed, and the installation each time of driving mechanism all is benchmark with locating piece on the star, and convenient and swift, repeatable accuracy also is guaranteed.
As shown in Figure 4, reality shows that in rail flight result the present invention can satisfy the performance requirement of each product, and it is big to day directed tracing generation disturbance torque well to have solved solar battery array, it is little that solar battery array is effectively shone area, contradictions such as solar irradiation angle, light application time instability.Created a new method for designing for the configuration design of follow-up radar satellite, it will be very extensive that this invention is used in this area.
Those skilled in the art can carry out various changes and distortion and not break away from the spirit and scope of the present invention on driving mechanism angle mount technical foundation of the present invention.Like this, if these modifications of the present invention and distortion belong within the scope of claim of the present invention and equivalent technologies thereof, then the invention is intended to also comprise these changes and distortion interior.
Claims (1)
1. the driving mechanism of a Synthetic Aperture Radar satellite is characterized in that, two driving mechanisms be installed in respectively the satellite celestial body ± the Y dividing plate on, and and solar cell piece paster face be 15 ° of angle mounts together; The solar battery array antisymmetry of two driving mechanisms is installed.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN 200910199674 CN102081156B (en) | 2009-11-30 | 2009-11-30 | Driving mechanism for synthetic aperture radar satellite |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN 200910199674 CN102081156B (en) | 2009-11-30 | 2009-11-30 | Driving mechanism for synthetic aperture radar satellite |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN102081156A true CN102081156A (en) | 2011-06-01 |
| CN102081156B CN102081156B (en) | 2013-12-18 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN 200910199674 Active CN102081156B (en) | 2009-11-30 | 2009-11-30 | Driving mechanism for synthetic aperture radar satellite |
Country Status (1)
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| CN (1) | CN102081156B (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105035364A (en) * | 2015-07-24 | 2015-11-11 | 上海卫星工程研究所 | Solar array driving swinging method for low-dip-angle orbit radar satellite |
| WO2016079945A1 (en) * | 2014-11-18 | 2016-05-26 | 川崎重工業株式会社 | Radar satellite and radar satellite system using same |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5626315A (en) * | 1990-11-30 | 1997-05-06 | Aerospatiale Societe Nationale Industrielle | Apparatus controlling the pitch attitude of a satellite by means of solar radiation pressure |
| US5895014A (en) * | 1996-07-31 | 1999-04-20 | Hughes Electronics Corporation | Satellite solar array and method of biasing to reduce seasonal output power fluctuations |
| US6439511B1 (en) * | 2000-07-26 | 2002-08-27 | Hughes Electronics Corporation | Thermal shock avoidance for satellite solar panels |
| US6637702B1 (en) * | 2002-04-24 | 2003-10-28 | Lockheed Martin Corporation | Nested beam deployable solar array |
-
2009
- 2009-11-30 CN CN 200910199674 patent/CN102081156B/en active Active
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5626315A (en) * | 1990-11-30 | 1997-05-06 | Aerospatiale Societe Nationale Industrielle | Apparatus controlling the pitch attitude of a satellite by means of solar radiation pressure |
| US5895014A (en) * | 1996-07-31 | 1999-04-20 | Hughes Electronics Corporation | Satellite solar array and method of biasing to reduce seasonal output power fluctuations |
| US6439511B1 (en) * | 2000-07-26 | 2002-08-27 | Hughes Electronics Corporation | Thermal shock avoidance for satellite solar panels |
| US6637702B1 (en) * | 2002-04-24 | 2003-10-28 | Lockheed Martin Corporation | Nested beam deployable solar array |
Non-Patent Citations (1)
| Title |
|---|
| 陈烈民: "《航天器结构与机构》", 31 January 2005 * |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2016079945A1 (en) * | 2014-11-18 | 2016-05-26 | 川崎重工業株式会社 | Radar satellite and radar satellite system using same |
| CN105035364A (en) * | 2015-07-24 | 2015-11-11 | 上海卫星工程研究所 | Solar array driving swinging method for low-dip-angle orbit radar satellite |
| CN105035364B (en) * | 2015-07-24 | 2017-05-03 | 上海卫星工程研究所 | Solar array driving swinging method for low-dip-angle orbit radar satellite |
Also Published As
| Publication number | Publication date |
|---|---|
| CN102081156B (en) | 2013-12-18 |
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