CN104197839A - Compensation method for spacecraft assembly accuracy influenced by gravity and temperature - Google Patents
Compensation method for spacecraft assembly accuracy influenced by gravity and temperature Download PDFInfo
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- CN104197839A CN104197839A CN201410514480.3A CN201410514480A CN104197839A CN 104197839 A CN104197839 A CN 104197839A CN 201410514480 A CN201410514480 A CN 201410514480A CN 104197839 A CN104197839 A CN 104197839A
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Abstract
The invention discloses a compensation method for the spacecraft assembly accuracy influenced by gravity and temperature based on a standard cubic mirror. The compensation method comprises the following steps: as for the compensation for a gravity deformation influence, the way of station distribution is adopted through theodolites, the pose angle matrix of a standard mirror with zero gravity are measured in comparison with a reference standard mirror before and after a balancing weight is additionally assembled respectively to calculate to obtain a warped matrix, and the pose matrix measured when equipment is mounted is corrected according to the warped matrix; the compensation method further comprises the following steps: as for the compensation for a temperature deformation influence, the standard cubic mirror on the equipment to be measured is measured at different temperatures compared with the pose angle matrix of a reference standard cubic mirror, and finally a corrected function is obtained through matching according to obtained matrix and temperature data. According to the invention, when the spacecraft works in outer space, temperature deformation can be corrected to obtain an on-orbit mounting pose matrix according to an on-orbit temperature, the corrected function and the initial mounting matrix after gravity compensation at the temperature of 20 DEG C.
Description
Technical field
The invention belongs to commercial measurement technical field, being specifically related to many transits of a kind of comprehensive utilization builds a station to form and measures network, analyzing spacecraft assembly precision by the benchmark prism square measurement on spacecraft structure plate is subject to effect of weightlessness and compensates, by under different temperatures to being related between the benchmark prism square between equipment that control survey obtains the temperature variant curve of assembly precision, provides the method for correction matrix according to curve.
Background technology
In spacecraft process, because the structural slab that affects spacecraft that is subject to erecting equipment gravity can produce distortion, thereby the attitude data that causes equipment to be installed is inaccurate, can not truly reflect the relative attitude data of spacecraft under weightlessness of space environment.Spacecraft, in the time that space moves, because temperature variation in space is larger, can cause spacecraft structure distortion, thereby affect the assembly precision of precision equipment simultaneously, and then affects its normal function performance.Therefore need to, in the spacecraft process of the test of ground, affected by gravity, temperature on its assembly precision and analyze and compensate.
The measuring technique of transit alignment fiducial prism square is extensively used in spacecraft develops, as the paper " the prism square attitude measurement based on autocollimation principle " that the Yang Zhen of information engineering university of PLA delivers in " mapping circular ", the paper " application of prism square in space-vehicle antenna general assembly is measured " that the Bai Hongwu of Xi’an Branch of the Chinese Academy of Space Technology delivers in " space electronic technology ".But the impact that these measuring methods can not abatement apparatus gravity at present and temperature deformation impact.
Also there is no at present open source literature report for the removing method of the impact that in spacecraft process, equipment gravity causes.The distortion producing for temperature variation under space environment, there is scholar to adopt photogrammetric method both at home and abroad, the paper " satellite antenna distortion photogrammetric technology under vacuum low-temperature environment " of delivering in " optical technology " as the Jiang Shanping of Beijing Satellite Environment Engineering Research Institute.Photogrammetricly need to paste some cooperative target punctuates at tested body structure surface, put the variation of position distribution with structure by measurement target drone, obtain by the distortion of geodesic structure.What adopt photogrammetric reflection is the internal modification of integral structure, not high for the measure of the change precision of the assembly precision between the instrument and equipment that in large-scale assembly, size is less, and correction matrix can not be provided.
In the present invention, propose based on the build a station experimental technique of measuring basis prism square of transit, the impact that can solve the mutual assembly precision between the equipment causing due to equipment gravity in spacecraft process provides correction matrix, also the deflection of the assembly precision between equipment under different temperatures environment after spacecraft be can analyze, deformation curve and compensation matrix provided.
Summary of the invention
The object of the present invention is to provide a kind of compensation method that is subject to the impact of gravity, temperature variation by measurement and experimental amount fractional analysis assembly precision between distinct device in spacecraft process.
For reaching above object, the technical solution used in the present invention is:
The compensation method that a kind of spacecraft assembly precision based on benchmark prism square is affected by gravity and temperature, comprise while carrying out gravity deformation impact compensation, in spacecraft process, near the equipment under test setting position of spacecraft structural slab one side, be fixedly installed a zero-g reference mirror, adopt the mode at four transit cloth stations, installing balancing weight fore-and-aft survey zero-g reference mirror additional with respect to the attitude angle matrix that is positioned at opposite side reference data mirror on spacecraft structure plate respectively, calculate the distortion matrix of the spacecraft structure plate producing because of equipment under test weight, the attitude matrix measuring after equipment under test being installed according to distortion matrix is revised, also comprise while carrying out temperature deformation impact compensation, on the spacecraft structure plate that equipment under test is set, paste heating temperature measuring equipment, key position is heated, heating is from 20 DEG C of room temperatures, after 10 DEG C of every risings of temperature, temperature keeps 10 minutes, under this state, measure the benchmark prism square that arranges on the equipment under test attitude angle matrix with respect to the reference data prism square arranging on spacecraft structure plate opposite side, be elevated to 90 DEG C always, last according to the matrix and the temperature data that obtain, matching obtains correction function, spacecraft is in the time that space moves, according to temperature in-orbit, initial installation matrix when correction function and 20 DEG C after gravity compensation, can obtain Installation posture matrix in-orbit to temperature deformation correction.
Wherein, described gravity deformation affects compensation process and comprises:
1 in spacecraft structure plate one side the position of Devices to test to be installed a zero-g reference mirror is set.
2) set up two mutually orthogonal reflectings surface that collimate respectively the reference data prism square arranging on spacecraft structure plate opposite side with transit T1, the T2 of alignment function, set up two mutually orthogonal reflectings surface that collimate respectively zero-g benchmark prism square with transit T3, the T4 of alignment function, transit T1 takes aim at mutually to form with transit T2, T3, T4 respectively and measures network.
3 are not having under the state that equipment and balancing weight are set to be measured, measure the cosine matrix A of zero-g reference mirror with respect to the attitude angle of reference data mirror
1.
4 transit T3, transit T4, under collimating status, install balancing weight 5, gather the collimation data of transit T3, transit T4, calculate the cosine matrix A of zero-g reference mirror with respect to the attitude angle of reference data mirror
2; Calculate the distortion correction matrix of structural slab according to the cosine matrix that balancing weight front and back are installed
5 pull down zero-g benchmark prism square and balancing weight, and equipment under test is installed.
6 set up two mutually orthogonal reflectings surface that collimate respectively the tested benchmark prism square arranging on equipment under test with transit T3, the transit T4 of alignment function, and transit T3, T4 and transit T1 again take aim at mutually to form and measure network.
7 gather collimation data, calculate the cosine matrix B of tested benchmark prism square with respect to the attitude angle of reference data mirror
2, use correction matrix C to B
2revise, obtain the attitude angle cosine matrix of equipment under test relatively tested benchmark prism square of its benchmark prism square under state of weightlessness
Wherein, described temperature deformation affects compensation process and comprises:
1 installs equipment under test on spacecraft structure plate, and paste heating temperature measuring equipment, and heating temperature measuring equipment comprises heating plate and temperature element, heating plate is arranged on spacecraft structure plate its temperature is added to heat control, and measure temperature by temperature element, temperature control is at 20-90 DEG C.
2 set up two mutually orthogonal reflectings surface that collimate respectively the reference data prism square arranging in spacecraft structure plate one side with transit T1, the T2 of alignment function, set up two mutually orthogonal reflectings surface that collimate respectively the tested benchmark prism square arranging on spacecraft structure plate opposite side with transit T3, the T4 of alignment function, transit T2, T3, T4 and transit T1 take aim at mutually to form and measure network.
3 at room temperature, gathers transit collimation data, calculates the attitude angle cosine matrix B of tested benchmark prism square with respect to reference data prism square
2.
4 heat spacecraft structure plate by heating temperature measuring equipment, in the time of 30 DEG C, 40 DEG C, 50 DEG C, 60 DEG C, 70 DEG C, 80 DEG C, 90 DEG C, stablize 10 minutes respectively, and gather the collimation angle of transit T3, T4, calculate the attitude angle cosine matrix B of tested benchmark prism square with respect to reference data prism square
3, B
4, B
5, B
6, B
7, B
8, B
9.
5 obtain matrix function f (t according to temperature and the matrix fitting measuring
n, B
n), at spacecraft in orbit in process, according to the spacecraft the measuring initial Installation posture matrix after gravity compensation when structure temperature, matrix function and 20 DEG C in-orbit, the attitude angle cosine matrix of tested benchmark prism square relative reference benchmark prism square when matching obtains in-orbit.
In the method for the invention, measure and experimental provision mainly comprises: reference data prism square, equipment under test benchmark prism square, zero-g benchmark prism square, 4 transits with alignment function, balancing weight, heating and temperature collecting devices, control system.
In this article, term " benchmark prism square " is a prism square as benchmark, it is an orthogonal hexahedron, every two adjacent mirror surface normals are mutually orthogonal, three mutually orthogonal minute surface normals represent the coordinate axis sensing of equipment (equipment under test) body coordinate system, and the weight of benchmark prism square is no more than 5g.
Term " reference data prism square " is also that it is as the benchmark prism square with reference to coordinate system as a prism square with reference to benchmark.Term " equipment under test benchmark prism square " be one on equipment under test for characterizing the benchmark prism square of equipment under test coordinate system.
Equally, term " zero-g benchmark prism square " is also a prism square, and it is for the benchmark prism square of auxiliary zero-g experiment.Wherein, transit monitors the reflecting surface of tested benchmark prism square for collimating, wherein two mutually orthogonal reflectings surface of two collimation reference data prism squares, other two for collimating two mutually orthogonal reflectings surface of tested benchmark prism square.
In addition, the copper billet of balancing weight for simulating tested erecting equipment weight, quality is identical with equipment under test.Heating and temperature collecting device produce distortion and gather heating the temperature reaching for test product being carried out to spot heating.
The present invention has following beneficial effect:
Spacecraft assembly deflections of the present invention is analyzed and compensation method, the mutual attitude relation between equipment after spacecraft can effectively analyzed is subject to the impact of weightlessness of space, temperature variation, and provide correction matrix, can obtain accurately the assembly precision data of spacecraft in-orbit time, provide powerful support for for accurate attitude control and the mapping etc. of spacecraft provide.The method is applied in the spacecraft such as High Resolution Remote Sensing Satellites, the telstar development of China.
Brief description of the drawings
Fig. 1 is the schematic diagram of tested project in method of the present invention, and wherein tested benchmark prism square 4 sticks on equipment under test 3, and reference data prism square 1 is fixedly mounted on spacecraft structure plate 2 with equipment under test 3.
Wherein, 1-reference data prism square, 2-structural slab, 3-equipment under test, the tested benchmark prism square of 4-;
Fig. 2 is that in method of the present invention, gravity deformation affects compensating measure schematic diagram, has fixed reference data prism square 1, zero-g benchmark prism square 6 on spacecraft structure plate 2, sets up the transit T with alignment function
1, T
2collimate respectively two mutually orthogonal reflectings surface of reference data prism square 1.Set up the transit T with alignment function
3, T
4collimate respectively two mutually orthogonal reflectings surface of zero-g benchmark prism square.T
1, T
2, T
3, T
4form and measure network.Balancing weight 5 can be fixed near zero-g benchmark prism square position, and balancing weight 5 is identical in quality with equipment under test.
Wherein: 1-reference data prism square, 2-structural slab, 5-balancing weight (identical with measured piece), 6-zero-g reference mirror, T1-transit 1,, T2-transit 2, T3-transit 3, T4-transit 4.
Fig. 3 is the schematic diagram that in method of the present invention, temperature affects compensating measure; On the basis of Fig. 1, on spacecraft structure plate 2, fixedly paste heating temperature measuring equipment 7.Set up the transit T with alignment function
1, T
2two mutually orthogonal reflectings surface that collimate respectively reference data prism square 1, set up the transit T with alignment function
3, T
4collimate respectively two mutually orthogonal reflectings surface of tested benchmark prism square.T
1, T
2, T
3, T
4form and measure network.Heating temperature measuring equipment 7 can add gentle thermometric to spacecraft structure plate 2 key positions as required.
Wherein: 1-reference data prism square, 2-structural slab, 3-equipment under test, the tested benchmark prism square of 4-, 7-heat temperature measuring equipment, T1-transit 1,, T2-transit 2, T3-transit 3, T4-transit 4.
Embodiment
What below introduce is the embodiment as content of the present invention, below by embodiment, content of the present invention is further illustrated.Certainly, describing following embodiment is only the content of example different aspect of the present invention, and should not be construed as the restriction scope of the invention.
Fig. 1 has shown the schematic diagram of tested project in method of the present invention, and wherein tested benchmark prism square 4 sticks on equipment under test 3, and reference data prism square 1 is fixedly mounted on spacecraft structure plate 2 with equipment under test 3.
In the method for the invention, for gravity deformation experiment analytical method as shown in Figure 2, implementation step is as follows:
1) on spacecraft structure plate 2, a zero-g reference mirror 6 is installed in the position of equipment to be installed.
2) set up two mutually orthogonal reflectings surface that collimate respectively reference data prism square 1 with transit T1, the T2 of alignment function.Set up two mutually orthogonal reflectings surface that collimate respectively zero-g benchmark prism square 6 with transit T3, the T4 of alignment function.T1 takes aim at mutually to form with T2, T3, T4 respectively and measures network.
3) do not having under the state of erecting equipment 3 and counterweight 5, measuring the cosine matrix A of zero-g reference mirror 6 with respect to the attitude angle of reference data mirror 1
1.
4) transit T3, T4, under collimation monitored state, install balancing weight 5, gather the collimation data of transit T3, T4, calculate the cosine matrix A of zero-g reference mirror 6 with respect to the attitude angle of reference data mirror 1
2.
5) calculate the distortion correction matrix of structural slab
6) zero-g benchmark prism square 6 and balancing weight 5 are pulled down, equipment under test 3 is installed.
7) set up two mutually orthogonal reflectings surface that collimate respectively the tested benchmark prism square 4 on equipment under test 3 with transit T3, the T4 of alignment function, T3, T4 and T1 again take aim at mutually to form and measure network.
8) gather collimation data, calculate the cosine matrix B of tested benchmark prism square 4 with respect to the attitude angle of reference data mirror 1
2.
9) use correction matrix C to B
2revise, obtain the attitude angle cosine matrix of equipment under test 3 relatively tested benchmark prism square 1 of its benchmark prism square 4 under state of weightlessness
For temperature deformation experiment analytical method as shown in Figure 3, implementation step is as follows:
1) equipment under test 3 is installed on spacecraft structure plate 2, and is pasted heating temperature measuring equipment 7 (heating plate is arranged on structural slab its temperature is controlled, and temperature control is at 20-90 DEG C).
2) set up two mutually orthogonal reflectings surface that collimate respectively reference data prism square 1 with transit T1, the T2 of alignment function, set up two mutually orthogonal reflectings surface that collimate respectively tested benchmark prism square 4 with transit T3, the T4 of alignment function.T2, T3, T4 and T1 take aim at mutually to form and measure network.
3) under the normal temperature environment of laboratory (20 DEG C), gather transit collimation data, calculate the attitude angle cosine matrix B of tested benchmark prism square 4 with respect to reference data prism square 1
2.
4) by heating temperature measuring equipment 7, the key position of spacecraft structure plate 2 is heated, in the time of 30 DEG C, 40 DEG C, 50 DEG C, 60 DEG C, 70 DEG C, 80 DEG C, 90 DEG C, stablize 10 minutes respectively, and gather the collimation angle of transit T3, T4, calculate the attitude angle cosine matrix B of tested benchmark prism square 4 with respect to reference data prism square 1
3, B
4, B
5, B
6, B
7, B
8, B
9.
5) obtain matrix function f (t according to temperature and the matrix fitting measuring
n, B
n).
At spacecraft in orbit in process, during according to 20 DEG C, the spacecraft structure temperature measuring, matrix function and ground, carried out the revised initial matrix B of gravity deformation
1, matching obtains in-orbit the weightless and attitude angle cosine matrix of tested benchmark prism square 4 relative reference benchmark prism squares 1 under temperature conditions in-orbit.
Although above the specific embodiment of the present invention has been given to describe in detail and explanation; but what should indicate is; we can conception according to the present invention carry out various equivalence changes and amendment to above-mentioned embodiment; when its function producing does not exceed spiritual that instructions and accompanying drawing contain yet, all should be within protection scope of the present invention.
Claims (3)
1. the compensation method that the spacecraft assembly precision based on benchmark prism square is affected by gravity and temperature, comprise while carrying out gravity deformation impact compensation, near equipment under test setting position in spacecraft process in spacecraft structural slab one side, be fixedly installed a zero-g reference mirror, adopt the mode at four transit cloth stations, installing balancing weight fore-and-aft survey zero-g reference mirror additional with respect to the attitude angle matrix that is positioned at opposite side reference data mirror on spacecraft structure plate respectively, calculate the distortion matrix of the spacecraft structure plate producing because of equipment under test weight, the attitude matrix measuring after equipment under test being installed according to distortion matrix is revised, also comprise while carrying out temperature deformation impact compensation, on the spacecraft structure plate that equipment under test is set, paste heating temperature measuring equipment, key position is heated, heating is from 20 DEG C of room temperatures, after 10 DEG C of every risings of temperature, temperature keeps 10 minutes, under this state, measure the benchmark prism square that arranges on the equipment under test attitude angle matrix with respect to the reference data prism square arranging on spacecraft structure plate opposite side, be elevated to 90 DEG C always, last according to the matrix and the temperature data that obtain, matching obtains correction function, spacecraft is in the time that space moves, according to temperature in-orbit, initial installation matrix when correction function and 20 DEG C after gravity compensation, can obtain Installation posture matrix in-orbit to temperature deformation correction.
2. the method for claim 1, wherein described gravity deformation affects compensation process and comprises:
1) in spacecraft structure plate one side, the position of Devices to test to be installed arranges a zero-g reference mirror;
2) set up two mutually orthogonal reflectings surface that collimate respectively the reference data prism square arranging on spacecraft structure plate opposite side with transit T1, the T2 of alignment function, set up two mutually orthogonal reflectings surface that collimate respectively zero-g benchmark prism square with transit T3, the T4 of alignment function, transit T1 takes aim at mutually to form with transit T2, T3, T4 respectively and measures network;
3) do not having under the state of Devices to test and balancing weight, measuring the cosine matrix A of zero-g reference mirror with respect to the attitude angle of reference data mirror
1;
4) transit T3, transit T4, under collimating status, install balancing weight 5, gather the collimation data of transit T3, transit T4, calculate the cosine matrix A of zero-g reference mirror with respect to the attitude angle of reference data mirror
2; Calculate the distortion correction matrix of structural slab according to the cosine matrix that balancing weight front and back are installed
5) zero-g benchmark prism square and balancing weight are pulled down, equipment under test is installed;
6) set up two mutually orthogonal reflectings surface that collimate respectively the tested benchmark prism square arranging on equipment under test with transit T3, the transit T4 of alignment function, transit T3, T4 and transit T1 again take aim at mutually to form and measure network;
7) gather collimation data, calculate the cosine matrix B of tested benchmark prism square with respect to the attitude angle of reference data mirror
2, use correction matrix C to B
2revise, obtain the attitude angle cosine matrix of equipment under test relatively tested benchmark prism square of its benchmark prism square under state of weightlessness
3. system as claimed in claim 1 or 2, wherein, described temperature deformation affects compensation process and comprises:
1) on spacecraft structure plate, equipment under test is installed, and paste heating temperature measuring equipment, and heating temperature measuring equipment comprises heating plate and temperature element, heating plate is arranged on spacecraft structure plate its temperature is added to heat control, and measure temperature by temperature element, temperature control is at 20-90 DEG C;
2) set up two mutually orthogonal reflectings surface that collimate respectively the reference data prism square arranging in spacecraft structure plate one side with transit T1, the T2 of alignment function, set up two mutually orthogonal reflectings surface that collimate respectively the tested benchmark prism square arranging on spacecraft structure plate opposite side with transit T3, the T4 of alignment function, transit T2, T3, T4 and transit T1 take aim at mutually to form and measure network;
3) at 20 DEG C of room temperatures, gather transit collimation data, calculate the attitude angle cosine matrix B of tested benchmark prism square with respect to reference data prism square
2;
4) by heating temperature measuring equipment, spacecraft structure plate is heated, in the time of 30 DEG C, 40 DEG C, 50 DEG C, 60 DEG C, 70 DEG C, 80 DEG C, 90 DEG C, stablize 10 minutes respectively, and gather the collimation angle of transit T3, T4, calculate the attitude angle cosine matrix B of tested benchmark prism square with respect to reference data prism square
3, B
4, B
5, B
6, B
7, B
8, B
9;
5) obtain matrix function f (t according to temperature and the matrix fitting measuring
n, B
n), so that at spacecraft in orbit in process, according to the spacecraft structure temperature and the matrix function that measure, matching obtains the attitude angle cosine matrix of tested benchmark prism square relative reference benchmark prism square.
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Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106546440A (en) * | 2016-10-18 | 2017-03-29 | 航天东方红卫星有限公司 | A kind of test method of the checking heat control system performance suitable for Gravity Satellite |
| CN108387547A (en) * | 2017-12-29 | 2018-08-10 | 华中科技大学 | A kind of survey spectrum compensation method of infared spectrum association remote sensing equipment |
| CN109631828A (en) * | 2019-01-23 | 2019-04-16 | 中国科学院长春光学精密机械与物理研究所 | The detection method of the reference axis angle of adjacent space coordinate system based on block prism |
| CN112033298A (en) * | 2020-08-25 | 2020-12-04 | 大连理工大学 | Spacecraft cabin body weight deformation measurement method based on motionless datum points |
| CN112284279A (en) * | 2020-09-21 | 2021-01-29 | 北京空间飞行器总体设计部 | High-precision spacecraft structure size deformation measurement method |
| CN113804429A (en) * | 2021-10-28 | 2021-12-17 | 北京卫星环境工程研究所 | Precision compensation method for equipment in spacecraft sealed cabin in rail pressure difference environment cabin |
| CN114636386A (en) * | 2022-02-28 | 2022-06-17 | 浙江时空道宇科技有限公司 | Angle measuring method, device, system and computer readable storage medium |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN85101015A (en) * | 1985-04-01 | 1986-08-27 | 北京科学仪器厂 | Gravity deformation, temperature deformation, microclimate compensating for variations method and compensation mechanism |
| KR20050057755A (en) * | 2003-12-11 | 2005-06-16 | 한국항공우주연구원 | Theodolite |
| CN102032918A (en) * | 2010-10-20 | 2011-04-27 | 郑州辰维科技股份有限公司 | Method for calibrating direction of three-probe start sensor |
| CN102692238A (en) * | 2012-06-12 | 2012-09-26 | 中国人民解放军92537部队 | Reset precision detection method of ship inertia navigated impact isolator |
| CN103499318A (en) * | 2013-10-21 | 2014-01-08 | 中国科学院光电技术研究所 | Method for measuring dead weight deflection of optical element |
| CN104038295A (en) * | 2014-06-06 | 2014-09-10 | 西安电子科技大学 | Deformed array antenna scattering performance analyzing method based on electromechanical coupling |
-
2014
- 2014-09-29 CN CN201410514480.3A patent/CN104197839B/en not_active Expired - Fee Related
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN85101015A (en) * | 1985-04-01 | 1986-08-27 | 北京科学仪器厂 | Gravity deformation, temperature deformation, microclimate compensating for variations method and compensation mechanism |
| KR20050057755A (en) * | 2003-12-11 | 2005-06-16 | 한국항공우주연구원 | Theodolite |
| CN102032918A (en) * | 2010-10-20 | 2011-04-27 | 郑州辰维科技股份有限公司 | Method for calibrating direction of three-probe start sensor |
| CN102692238A (en) * | 2012-06-12 | 2012-09-26 | 中国人民解放军92537部队 | Reset precision detection method of ship inertia navigated impact isolator |
| CN103499318A (en) * | 2013-10-21 | 2014-01-08 | 中国科学院光电技术研究所 | Method for measuring dead weight deflection of optical element |
| CN104038295A (en) * | 2014-06-06 | 2014-09-10 | 西安电子科技大学 | Deformed array antenna scattering performance analyzing method based on electromechanical coupling |
Non-Patent Citations (2)
| Title |
|---|
| 杨振等: "多种互瞄条件下立方镜准直传递方法研究", 《宇航计测技术》 * |
| 王智等: "测绘相机立方镜与星敏立方镜转换矩阵的标定", 《光学精密工程》 * |
Cited By (12)
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|---|---|---|---|---|
| CN106546440A (en) * | 2016-10-18 | 2017-03-29 | 航天东方红卫星有限公司 | A kind of test method of the checking heat control system performance suitable for Gravity Satellite |
| CN106546440B (en) * | 2016-10-18 | 2018-10-09 | 航天东方红卫星有限公司 | A kind of test method of verification heat control system performance suitable for Gravity Satellite |
| CN108387547A (en) * | 2017-12-29 | 2018-08-10 | 华中科技大学 | A kind of survey spectrum compensation method of infared spectrum association remote sensing equipment |
| CN108387547B (en) * | 2017-12-29 | 2019-09-27 | 华中科技大学 | A kind of survey spectrum compensation method of infared spectrum association remote sensing equipment |
| CN109631828A (en) * | 2019-01-23 | 2019-04-16 | 中国科学院长春光学精密机械与物理研究所 | The detection method of the reference axis angle of adjacent space coordinate system based on block prism |
| CN112033298A (en) * | 2020-08-25 | 2020-12-04 | 大连理工大学 | Spacecraft cabin body weight deformation measurement method based on motionless datum points |
| CN112033298B (en) * | 2020-08-25 | 2021-08-20 | 大连理工大学 | Spacecraft cabin body weight deformation measurement method based on motionless datum points |
| CN112284279A (en) * | 2020-09-21 | 2021-01-29 | 北京空间飞行器总体设计部 | High-precision spacecraft structure size deformation measurement method |
| CN112284279B (en) * | 2020-09-21 | 2022-06-21 | 北京空间飞行器总体设计部 | High-precision spacecraft structure size deformation measurement method |
| CN113804429A (en) * | 2021-10-28 | 2021-12-17 | 北京卫星环境工程研究所 | Precision compensation method for equipment in spacecraft sealed cabin in rail pressure difference environment cabin |
| CN113804429B (en) * | 2021-10-28 | 2023-11-10 | 北京卫星环境工程研究所 | Method for compensating equipment precision of spacecraft sealed cabin in rail pressure difference environment cabin |
| CN114636386A (en) * | 2022-02-28 | 2022-06-17 | 浙江时空道宇科技有限公司 | Angle measuring method, device, system and computer readable storage medium |
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