CN110967005A - Imaging method and imaging system for on-orbit geometric calibration through star observation - Google Patents

Imaging method and imaging system for on-orbit geometric calibration through star observation Download PDF

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CN110967005A
CN110967005A CN201911271977.6A CN201911271977A CN110967005A CN 110967005 A CN110967005 A CN 110967005A CN 201911271977 A CN201911271977 A CN 201911271977A CN 110967005 A CN110967005 A CN 110967005A
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star
satellite
detector
gyroscope
orbit
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CN110967005B (en
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余达
刘金国
吴国栋
姜肖楠
王钢
裴君妍
张雨
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Changchun Institute of Optics Fine Mechanics and Physics of CAS
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    • G—PHYSICS
    • G01—MEASURING; TESTING
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    • G—PHYSICS
    • G01—MEASURING; TESTING
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    • G01—MEASURING; TESTING
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Abstract

An imaging method and an imaging system for performing on-orbit geometric calibration through viewing a star relate to an imaging system for performing on-orbit geometric calibration in a satellite observation technology, and solve the problem of how to combine space optical imaging and an electronic technology to ensure that an on-orbit internal orientation element of a high-precision imaging system is calibrated, and comprise a satellite platform, a star sensor, a linear array camera and a gyroscope; the star sensor, the line array camera and the gyroscope are all installed on the satellite platform, the star sensor measures the in-orbit attitude of the satellite, and the gyroscope outputs the angular speed of the satellite. The imaging method adopts a combined star observation mode of star sensor and gyroscope, and sets proper integration time according to the imaging parameters of the linear array camera, the star to be observed and the like; setting the constant-speed swing speed of the satellite according to the set integration time and the allowable image shift amount; and selecting a proper gyroscope according to the flutter characteristic of the satellite, performing combined extended Kalman filtering with the star sensor, accurately acquiring the attitude variation among the linear array lines, and improving the geometric strength of measurement.

Description

Imaging method and imaging system for on-orbit geometric calibration through star observation
Technical Field
The invention relates to an imaging system for on-orbit geometric calibration in a satellite observation technology, in particular to a method for on-orbit geometric calibration through star observation and an imaging system.
Background
Most of the existing on-orbit calibration is based on the real coordinates of control points of a calibration field, is influenced by various factors such as the orbital regression period of a satellite, weather and the like, has low calibration efficiency, and wastes time and labor. The relative angular distance between the star points in the sky is unchanged, and the star points are natural geometric calibration sources. If the area array detector is used for observing the star for calibration, the method is relatively easy to realize and has high geometric strength; most detectors applied to earth observation are linear array detectors, a star-viewing image needs to be synthesized through multiple swing imaging, and the geometric intensity is low.
The invention aims to solve the technical problem of how to combine space optical imaging and electronic technology and ensure that the in-orbit element calibration of a high-precision imaging system is carried out on track.
Disclosure of Invention
The invention provides an imaging method and an imaging system for on-orbit geometric calibration through star observation, aiming at solving the problem of how to combine space optical imaging and electronic technology to ensure the on-orbit inner orientation element calibration of a high-precision imaging system.
An imaging method for on-orbit geometric calibration through star observation is realized by the following steps:
step one, setting the integral time of a detector according to the imaging parameters of a linear array camera and the observed stars and the like, specifically:
the method comprises the following steps of firstly, setting a constraint relation between a focal ratio F of an optical system of the linear array camera and a pixel size a of a detector as follows:
2.44λF＞a
wherein λ is the average wavelength of light;
step one and two, setting MVStar etc. light integration time t in detectorintThe number of star light charges generated on a single pixel is less than that of the full trap charge of the detector, the signal-to-noise ratio is greater than or equal to 2, and the formula is as follows:
Figure RE-GDA0002370348300000021
in the formula, sFWCNumber of electrons for a full well of the detector, QdarkDark charge generated for dark current, IdarkIs the dark current of the detector, QsIs the star light charge number;
secondly, determining the constant-speed swing speed of the satellite according to the set integration time of the detector and the allowable image motion amount; represented by the formula:
f×2tg(w×tint/2)≤κa
wherein f is the focal length of the optical system, w is the angular velocity of the satellite in uniform swing, and tintTaking the integral time as an integral time, a as the pixel size of the detector, and k as an allowable image shift coefficient;
determining parameters of a gyroscope, performing combined extended Kalman filtering on the gyroscope and a star sensor to acquire the attitude variation of the line-scan digital camera, and outputting the angular velocity of a satellite by the gyroscope;
selecting a gyroscope according to the flutter characteristics of the satellite, wherein the output frequency f of the gyroscope is requiredtuoluoGreater than the flutter frequency f of the satellitestar_zcPrecision delta of the gyroscopetuoluoLess than or equal to 0.1 pixel; represented by the formula:
ftuoluo≥fstar_zc
Figure RE-GDA0002370348300000022
the system for on-orbit geometric calibration through star observation comprises a satellite platform, a star sensor, a linear array camera and a gyroscope; the star sensor, the linear array camera and the gyroscope are all installed on a satellite platform, the star sensor measures the in-orbit attitude of the satellite, and the gyroscope outputs the angular speed of the satellite.
The invention has the beneficial effects that: the imaging system provided by the invention sets proper integration time according to the imaging parameters of the camera, so that the sufficiently high signal-to-noise ratio of star point centroid extraction is realized, and the centroid extraction precision is ensured. The attitude measurement precision in the dynamic imaging process is improved by using the high-precision gyroscope, so that the problem of weak geometric strength of linear array sweep calibration is solved.
The imaging method of the invention adopts a combined star observation mode of star sensor and gyroscope, and sets proper integration time according to the imaging parameters of the linear array camera, the star to be observed and the like; then setting the constant swing speed of the satellite according to the set integration time and the allowable image shift amount; and finally, selecting a proper gyroscope according to the flutter characteristic of the satellite, performing combined extended Kalman filtering with the satellite sensor, and accurately acquiring the attitude variation among the linear array lines, thereby improving the geometric strength of measurement.
Drawings
FIG. 1 is a structural diagram of an on-orbit geometric calibration system by observing stars according to the present invention.
Detailed Description
In the first embodiment, the present embodiment is described with reference to fig. 1, and an in-orbit geometric calibration system by observing a star mainly includes a satellite platform, a star sensor, a line camera, and a gyroscope. The star sensor, the linear array camera and the gyroscope are all installed on the satellite platform, the star sensor measures the in-orbit attitude of the satellite, and the gyroscope outputs the angular speed output by the satellite.
MVThe illuminance of the star and the like generated outside the atmosphere is as follows:
Figure RE-GDA0002370348300000033
in the formula E0＝2.54×10-6lux is the illuminance of a zero-isocenter outside the atmosphere, and the lux is the unit of illuminance;MVrepresenting stars, etc.
Requirement MVStar etc. light integration time t in detectorintThe number of possible star light charges generated on a single pixel cannot reach the number of full trap charges of the detector, and the signal-to-noise ratio cannot be lower than 2.
Figure RE-GDA0002370348300000031
sFWCNumber of electrons for a full well of the detector, QdarkDark charge generated for dark current, IdarkIs the dark current of the detector. The starlight charge QsRepresented by the formula:
Figure RE-GDA0002370348300000032
in the formula, τ0Is the transmittance of the optical system, D is the effective aperture of the optical system, λ is the average wavelength of light, h is the Planck constant, c is the speed of light, tintIntegration time of the detector, ηqThe quantum efficiency of the detector is shown, n is the number of pixels occupied by the diffuse speckles in one-dimensional direction, and k isλThe integral coefficient of the spectral range of the detector is 0-1.
Figure RE-GDA0002370348300000041
In the formula Em(lambda) is the spectral irradiance of the star point,λminthe minimum wavelength of light which can be received by the detector is the maximum wavelength of light which can be received by the detector.
The imaging system described in this embodiment performs in-orbit observation intra-satellite orientation element detection, and the constraints of the angular velocity and the integration time of the satellite uniform-velocity oscillation are as follows:
f×2tg(w×tint/2)≤κa
wherein f is the focal length of the optical system, w is the angular velocity of the satellite in uniform swing, and tintFor the integration time, a is the pixel size of the detector, and κ is the allowable image shift coefficient, which is greater than 0 and less than 3.
In the present embodiment, the output frequency f of the gyrotuoluoGreater than the flutter frequency f of the satellitestar_zcPrecision of the gyro deltatuoluoBetter than 0.1 pixel.
ftuoluo≥fstar_zc
Figure RE-GDA0002370348300000042
Wherein f is the focal length of the optical system, and a is the pixel size of the detector.
In this embodiment, the constraint relationship between the focal ratio F of the optical system of the line camera and the pixel size a of the detector is as follows:
2.44λF＞a
where λ is the average wavelength of light.
In this embodiment, the line-scan camera has two observation modes when viewing the star. One is that the satellite is in a low-speed uniform-speed swing state, and the balance of image shift amount and signal-to-noise ratio is carried out by selecting proper integral time; the other is that the satellite is in a stepping state, swings by an angle corresponding to one detector pixel each time, then gradually stays to a static state, takes a picture at the moment, and then starts swinging.
In a second specific embodiment, the second specific embodiment is a method for imaging an on-orbit geometric calibration system by observing a star, the method adopts a combined star observation mode of a star sensor and a gyroscope, firstly, an imaging parameter of a linear array camera is selected, the focal ratio and the central wavelength of an optical system of the linear array camera are required to be 2.44 times larger than the pixel size, and the imaging parameter is expressed as:
2.44λF＞a
where λ is the average wavelength of light.
Secondly, setting proper integration time according to the star of the pre-observation star and the spectral range of the detector to ensure that the linear array detector is not saturated and the signal-to-noise ratio is more than or equal to 2; namely:
setting MVStar etc. light integration time t in detectorintThe number of possible star light charges generated on a single pixel cannot reach the full of the detectorThe trap charge number, the signal-to-noise ratio, cannot be lower than 2.
Figure RE-GDA0002370348300000051
In the formula, sFWCNumber of electrons for a full well of the detector, QdarkDark charge generated for dark current, IdarkIs the dark current of the detector. The starlight charge QsRepresented by the formula:
Figure RE-GDA0002370348300000052
in the formula, τ0Is the transmittance of the optical system, D is the effective aperture of the optical system, λ is the average wavelength of light, h is the Planck constant, c is the speed of light, tintIntegration time of the detector, ηqThe quantum efficiency of the detector is shown, n is the number of pixels occupied by the diffuse speckles in one-dimensional direction, and k isλThe integral coefficient of the spectral range of the detector is 0-1.
Figure RE-GDA0002370348300000053
In the formula, Em(lambda) is the spectral irradiance of the star point,λminthe minimum wavelength of light which can be received by the detector is the maximum wavelength of light which can be received by the detector.
Then, setting the constant-speed swing speed of the satellite according to the set integration time and the allowable observing star image shift amount; namely: and (3) detecting the in-orbit observation satellite orientation elements, wherein the constraints of the angular velocity and the integral time of the uniform-speed swing of the satellite are as follows:
f×2tg(w×tint/2)≤κa
wherein f is the focal length of the optical system, w is the angular velocity of the satellite in uniform swing, and tintFor the integration time, a is the pixel size of the detector, and κ is the allowable image shift coefficient, which is greater than 0 and less than 3.
And finally, selecting a proper gyroscope according to the flutter characteristic of the satellite, wherein the output frequency of the gyroscope is greater than the flutter frequency of the satellite, the precision of the gyroscope is superior to 0.1 pixel, and the gyroscope and the star sensor are used for carrying out combined extended Kalman filtering to accurately acquire the attitude variation among linear array lines, so that the measured geometric strength is improved.
Output frequency f of the gyroscopetuoluoGreater than the flutter frequency f of the satellitestar_zcPrecision of the gyro deltatuoluoBetter than 0.1 pixel.
ftuoluo≥fstar_zc
Figure RE-GDA0002370348300000061
Wherein f is the focal length of the optical system, and a is the pixel size of the detector.
In this embodiment, the line-scan camera has two observation modes when viewing the star. One is that the satellite is in a low-speed uniform-speed swing state, and the balance of image shift amount and signal-to-noise ratio is carried out by selecting proper integral time; the other is that the satellite is in a stepping state, swings by an angle corresponding to one detector pixel each time, then gradually stays to a static state, takes a picture at the moment, and then starts swinging.
In the embodiment, the satellite platform adopts a satellite platform of a long-light satellite company, so that agile maneuvering can be realized; the star sensor adopts a high-precision star sensor of the aerospace 502; the linear array camera is designed based on a TDICMOS detector of a long-optical-time core company; the gyroscope is a high-speed gyroscope with domestic output frequency of 500 KHz.

Claims (6)

1. An imaging method for on-orbit geometric calibration through star observation is characterized in that: the imaging method is realized by the following steps:
step one, setting the integral time of a detector according to the imaging parameters of a linear array camera and the observed stars and the like, specifically:
the method comprises the following steps of firstly, setting a constraint relation between a focal ratio F of an optical system of the linear array camera and a pixel size a of a detector as follows:
2.44λF＞a
wherein λ is the average wavelength of light;
step one and two, setting MVStar etc. light integration time t in detectorintThe number of star light charges generated on a single pixel is less than that of the full trap charge of the detector, the signal-to-noise ratio is greater than or equal to 2, and the formula is as follows:
Figure RE-FDA0002382754210000011
in the formula, sFWCNumber of electrons for a full well of the detector, QdarkDark charge generated for dark current, IdarkIs the dark current of the detector, QsIs the star light charge number;
secondly, determining the constant-speed swing speed of the satellite according to the set integration time of the detector and the allowable image motion amount; represented by the formula:
f×2tg(w×tint/2)≤κa
wherein f is the focal length of the optical system, w is the angular velocity of the satellite in uniform swing, and tintTaking the integral time as an integral time, a as the pixel size of the detector, and k as an allowable image shift coefficient;
determining parameters of a gyroscope, performing combined extended Kalman filtering on the gyroscope and a star sensor to acquire the attitude variation of the line-scan digital camera, and outputting the angular velocity of a satellite by the gyroscope;
selecting a gyroscope according to the flutter characteristics of the satellite, wherein the output frequency f of the gyroscope is requiredtuoluoGreater than the flutter frequency f of the satellitestar_zcPrecision delta of the gyroscopetuoluoLess than or equal to 0.1 pixel; represented by the formula:
ftuoluo≥fstar_zc
Figure RE-FDA0002382754210000021
2. the imaging method for on-orbit geometric calibration by sight of a star as claimed in claim 1, wherein: in step one, MVThe illuminance of the star and the like generated outside the atmosphere is as follows:
Figure RE-FDA0002382754210000024
in the formula, E0Is the illuminance of a zero-iso star outside the atmosphere.
3. The imaging method for on-orbit geometric calibration by sight of a star as claimed in claim 1, wherein: in the first step, the number Q of the star light chargesRepresented by the formula:
Figure RE-FDA0002382754210000022
in the formula, τ0Is the transmittance of the optical system, D is the effective aperture of the optical system, λ is the average wavelength of light, h is the Planck constant, c is the speed of light, ηqThe quantum efficiency of the detector is shown, n is the number of pixels occupied by the diffuse speckles in one-dimensional direction, and k isλIs the integral coefficient of the spectral range of the detector.
4. The imaging method for on-orbit geometric calibration by sight of a star as claimed in claim 1, wherein: in the first step and the second step, the integral coefficient k of the spectral range of the detectorλThe value of (a) is between 0 and 1; represented by the formula:
Figure RE-FDA0002382754210000023
in the formula, Em(λ) is the spectral irradiance of the star point, λminIs the minimum wavelength of light received by the detector, lambdamaxThe maximum wavelength of light that can be received by the detector.
5. The imaging system of the imaging method for geometric calibration in orbit through observing a star as claimed in claim 1, wherein the system comprises a satellite platform, a star sensor, a line camera and a gyroscope; the star sensor, the linear array camera and the gyroscope are all installed on a satellite platform, the star sensor measures the in-orbit attitude of the satellite, and the gyroscope outputs the angular speed of the satellite.
6. The imaging system for in-orbit geometric calibration through star observation according to claim 5, wherein the line-scan camera has two observation modes when observing the star;
one is that the satellite is in a low-speed uniform-speed swing state, and the image shift amount and the signal-to-noise ratio are balanced by selecting the integral time of the detector;
the other is that the satellite is in a stepping state, swings by an angle corresponding to one detector pixel each time, then gradually moves to a static state, takes a picture at the moment, and then starts swinging.
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