CN111609857A - Space debris orbit determination traversal observation method and system - Google Patents

Space debris orbit determination traversal observation method and system Download PDF

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Publication number
CN111609857A
CN111609857A CN202010483246.4A CN202010483246A CN111609857A CN 111609857 A CN111609857 A CN 111609857A CN 202010483246 A CN202010483246 A CN 202010483246A CN 111609857 A CN111609857 A CN 111609857A
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satellite
axis
camera
observation
space debris
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陈起行
胡海鹰
朱永生
郑珍珍
董磊
盛蕾
王威
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Shanghai Engineering Center for Microsatellites
Innovation Academy for Microsatellites of CAS
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Shanghai Engineering Center for Microsatellites
Innovation Academy for Microsatellites of CAS
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    • 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
    • G01C21/24—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 specially adapted for cosmonautical navigation
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
    • G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
    • G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
    • G01S19/13—Receivers
    • G01S19/24—Acquisition or tracking or demodulation of signals transmitted by the system
    • G01S19/25—Acquisition or tracking or demodulation of signals transmitted by the system involving aiding data received from a cooperating element, e.g. assisted GPS
    • G01S19/258—Acquisition or tracking or demodulation of signals transmitted by the system involving aiding data received from a cooperating element, e.g. assisted GPS relating to the satellite constellation, e.g. almanac, ephemeris data, lists of satellites in view

Abstract

The invention discloses a method and a system for fixed-orbit traversal observation of space debris, which realize 'large-range scanning' from a starting deflection angle to an ending deflection angle by controlling a camera arranged in the direction of a satellite and a Y axis on a sun synchronous orbit in the morning and evening to rotate around a + Z axis so as to realize fixed-orbit traversal observation of space debris with a synchronous belt, and in the observation process, the angle of the satellite rotating around a satellite body and the + X axis is controlled by real-time calculation so as to ensure that the visual axis of the camera is always vertical to a geosynchronous belt with a specified latitude.

Description

Space debris orbit determination traversal observation method and system
Technical Field
The invention relates to the technical field of aerospace, in particular to a method for orbit determination, traversal and observation of space debris.
Background
At present, the quantity of various space aircrafts and fragments is greatly increased, the space activities are increasingly frequent, the satellite safety is seriously threatened, and particularly, the earth synchronous belt is a gathering area of a plurality of communication, broadcasting, meteorological, early warning and military reconnaissance satellites, so that the synchronous belt space fragment observation is increasingly important.
The space debris observation comprises two modes of a foundation and a space-based mode, wherein the space-based space debris observation has higher-limit spare coverage and observation timeliness compared with the foundation observation. But limited by the restriction of the sensor capability and the constraint of the track periodicity, the observation arc sections of the same observation object are distributed in space in a concentrated manner, so that the distribution field angle of the observation arc sections is smaller, and the main requirements of high-precision astronomical orbit determination on observation data are as follows: the continuous time of single observation is long, observes the arc length promptly, and the arc section is observed to many times is big and the arc section distribution homogeneity is high to many times of the distribution field angle of observing in the space, consequently, the space base is observed in orbit determination precision aspect, compares and is not showing the promotion with ground base observation data.
In order to improve the orbit determination precision of space-based observation, various solutions including joint orbit determination of various detection data, multi-satellite networking observation, carrying of a large-view-field detector and the like are provided by in-orbit models and scholars at home and abroad. The method comprises the following steps that an American space target monitoring satellite MSX comprehensively fixes the orbit by using SBV camera data and Millstone Hill radar data, and the orbit fixing precision is improved to be within 2km from the single observed 10km magnitude of an SBV camera; the subsequent SBSS system forms a constellation by 4 stars, and the target orbit determination precision of a high orbit space reaches 500 m; the 5-star constellation system of the newly built STARE nano satellite in the United states has the orbit determination precision index reaching 100m magnitude.
However, these methods mainly achieve the improvement of orbit determination accuracy by multi-source data fusion processing or a multi-station observation method using a multi-satellite system, and do not solve the problem of the improvement of orbit determination accuracy of single-satellite observation data of a single type load satellite.
Disclosure of Invention
Aiming at the problem of improving the cataloging and orbit determination capability of space debris in space-based space under the optical observation of a space-based single station, the invention provides a space debris orbit determination traversal observation method, which adopts a large-range scanning mode to observe the space debris through a camera arranged on a satellite, and comprises the following steps:
setting a satellite scanning starting deflection angle theta, a satellite scanning ending deflection angle and a satellite scanning speed according to observation requirements;
controlling the satellite to rotate around the satellite body and the X axis so that the visual axis of the camera is perpendicular to the earth synchronous belt at the specified latitude;
controlling the satellite to rotate around the satellite body and the Z axis, so that the included angle between the visual axis of the camera and the zero position of the satellite body and the Y axis is equal to the starting deflection angle theta at the observation starting time; and
and opening the camera, and simultaneously controlling the satellite to rotate around the satellite body and the Z axis according to the scanning speed until the included angle between the visual axis of the camera and the zero position of the satellite body and the Y axis is equal to theta so as to realize observation.
Further, the method further comprises the step of controlling the satellite to rotate around the + X axis of the satellite body when the camera observes, so that the visual axis of the camera is always perpendicular to the geosynchronous belt at the specified latitude.
Further, the satellite rotates around the + Z axis of the satellite body according to the scanning speed.
Further, the ending deflection angle is equal to- θ.
Further, the angle of the satellite rotating around the satellite body and the X axis is obtained by solving according to coordinate transformation and orbit geometry.
Further, when the camera performs observation, the exposure time is matched according to the scanning speed.
The invention also provides a space debris orbit determination traversal observation system, which comprises a camera and a controller, wherein the camera is installed on a satellite along the direction of the + Y axis of the satellite, and the controller is used for executing the observation method.
Further, the satellite is arranged on a morning and evening sun synchronous orbit, and the + Y axis reference of the satellite points to the normal line of the orbit surface in the light following direction.
Further, the field of view of the camera is not less than 2 ° × 2 °.
The invention provides a space debris orbit determination traversal observation method and a space debris orbit determination traversal observation system, which are suitable for high-precision orbit determination traversal observation and similar tasks of synchronous belt space debris. The method and the system can observe the synchronous belt area at the appointed latitude, can be popularized and applied to various tracks such as a sun synchronous track, a zero inclination angle, a small inclination angle and the like, and can be used for adaptively adjusting the installation shaft of the camera.
Drawings
To further clarify the above and other advantages and features of embodiments of the present invention, a more particular description of embodiments of the present invention will be rendered by reference to the appended drawings. It is appreciated that these drawings depict only typical embodiments of the invention and are therefore not to be considered limiting of its scope. In the drawings, the same or corresponding parts will be denoted by the same or similar reference numerals for clarity.
Fig. 1 is a schematic flow chart of a space debris orbit determination traversal observation method according to an embodiment of the present invention; and
fig. 2 is a schematic diagram illustrating a space debris tracking and traversing observation method according to an embodiment of the present invention.
Detailed Description
In the following description, the present invention is described with reference to examples. One skilled in the relevant art will recognize, however, that the embodiments may be practiced without one or more of the specific details, or with other alternative and/or additional methods, materials, or components. In other instances, well-known structures, materials, or operations are not shown or described in detail to avoid obscuring aspects of the invention. Similarly, for purposes of explanation, specific numbers, materials and configurations are set forth in order to provide a thorough understanding of the embodiments of the invention. However, the invention is not limited to these specific details. Further, it should be understood that the embodiments shown in the figures are illustrative representations and are not necessarily drawn to scale.
Reference in the specification to "one embodiment" or "the embodiment" means that a particular feature, structure, or characteristic described in connection with the embodiment is included in at least one embodiment of the invention. The appearances of the phrase "in one embodiment" in various places in the specification are not necessarily all referring to the same embodiment.
It should be noted that the embodiment of the present invention describes the process steps in a specific order, however, this is only for the purpose of illustrating the specific embodiment, and does not limit the sequence of the steps. Rather, in various embodiments of the present invention, the order of the steps may be adjusted according to process adjustments.
The invention provides a space debris orbit determination traversal observation method and a space debris orbit determination traversal observation system, aiming at the problem of how to improve the orbit determination precision of single satellite observation data of a single type load satellite.
Fig. 1 is a schematic flow chart of a space debris tracking and traversing observation method according to an embodiment of the present invention. As shown in fig. 1, in a space debris orbit determination traversal observation method, a camera mounted on a satellite is used to observe space debris by using a "large-scale scanning" mode as shown in fig. 2, wherein the satellite is deployed on a solar synchronous orbit in the morning and evening, and a + Y axis of the satellite points to a normal line of an orbit surface in a down-light direction. The observation method comprises the following steps:
first, in step 101, observation parameters are set. Setting observation parameters according to observation requirements, wherein the observation parameters comprise: satellite scan start deflection angle theta, end deflection angle, scan speed omega, and observation start time Ts(ii) a In one embodiment of the invention, said end deflection angle is equal to- θ;
next, at stepAnd 102, adjusting the initial attitude of the satellite. At the observation start time TsAdjusting the satellite attitude to enter an initial scanning state, wherein the adjustment of the satellite attitude comprises the following steps:
controlling the satellite to rotate around the satellite body + X axis by an angle gamma0A geosynchronous belt, such that the visual axis of the camera is perpendicular to a specified latitude, the latitude being denoted as Δ i; in one embodiment of the invention, said rotation angle γ0The method is obtained by solving according to the track geometry, and comprises the following specific steps:
first, the intersection point of the current camera visual axis of the satellite and the geosynchronous belt GEO with the latitude Δ i is taken as a starting observation point G0Then for the earth' S center O, the satellite position S and the starting observation point G0Formed triangle △ OSG0In the middle, according to the cosine theorem, there are:
Figure BDA0002518098210000041
wherein the content of the first and second substances,
|SG0l is the distance between the satellite and the observation point, and is marked as f0；
|OG0I is the distance from the geocentric to the earth synchronous belt, and the magnitude of the I is the right ascension R of the earth synchronous beltgeo；
The size of the distance from the earth center to the satellite is calculated by the coordinate R of the satellite in the J2000 coordinate system, wherein the coordinate R is [ R ═ R [ ]xRyRz]TIs a known value, then
Figure BDA0002518098210000042
And
from the angle definition of the vector, cos ∠ OSG can be obtained0With said angle gamma0The relationship between:
Figure BDA0002518098210000043
wherein n isosAnd
Figure BDA0002518098210000044
is unit vector coordinates under the inertial system:
Figure BDA0002518098210000045
and
according to the coordinate R of the satellite in the J2000 coordinate system and the transformation matrix M from the inertial system to the J2000 coordinate system1The calculation can obtain:
Figure BDA0002518098210000051
then, according to the vector relationship, it can be obtained:
OG0＝OS+SG0,
for the purpose of calculation, the coordinates of each vector in the following equation of the J2000 coordinate system are used, where:
Figure BDA0002518098210000052
wherein the content of the first and second substances,
Figure BDA0002518098210000053
Figure BDA0002518098210000054
and
Figure BDA0002518098210000055
wherein M is1Is a transformation matrix of the inertial system to the J2000 coordinate system, M2Is a transformation matrix of the inertial system to the body coordinate system of the observation state,
Figure BDA0002518098210000056
while
Figure BDA0002518098210000057
Is a vector of the satellite position to the starting observation point positionThe coordinates in the body coordinate system include:
Figure BDA0002518098210000058
in summary, the following results can be obtained: rgeosinΔi＝Rz+fsinγ0(ii) a And
finally, two equations obtained in the previous steps are combined, and the rotation angle gamma is solved0A value of (d); and
controlling the satellite to rotate around the satellite body and the Z axis, so that the included angle between the visual axis of the camera and the zero position of the satellite body and the Y axis is equal to the starting deflection angle theta at the observation starting time; and
finally, in step 103, an observation is made. And after the satellite enters an initial scanning state, a camera is started, and the satellite is controlled to rotate around the satellite body and the Z axis according to the scanning speed until the included angle between the visual axis of the camera and the zero position of the satellite body and the Y axis is equal to theta, so that orbit determination traversal observation of the synchronous belt at the specified latitude is realized. During the whole observation period, the angle of the satellite rotating around the + Z axis is 2 theta, and the observation time is 2 theta/omega. In one embodiment of the invention, the camera exposure time needs to be set according to the scanning speed. In another embodiment of the invention, in order to ensure that the visual axis of the camera is always perpendicular to the geosynchronous belt at a specified latitude during observation, the rotation angle gamma of the satellite around the satellite body plus the X axis needs to be calculated in real timetAnd controlling the satellite to rotate according to the rotation angle gammatRotating around the satellite body + X axis:
first, the intersection point of the current camera visual axis of the satellite and the geosynchronous belt GEO is taken as an observation point GtThen for the earth' S center O, the satellite position S and the observation point GtFormed triangle △ OSGtIn the middle, according to the cosine theorem, there are:
Figure BDA0002518098210000061
wherein the content of the first and second substances,
|SGtl is the satellite and the observationThe distance between the points is marked as f;
|OGti is the distance from the geocentric to the earth synchronous belt, and the magnitude of the I is the right ascension R of the earth synchronous beltgeo；
The size of the distance from the earth center to the satellite is calculated by the coordinate R of the satellite in the J2000 coordinate system, wherein the coordinate R is [ R ═ R [ ]xRyRz]TIs a known value, then
Figure BDA0002518098210000062
And
from the angle definition of the vector, cos ∠ OSG can be obtainedtWith said angle gammatThe relationship between:
Figure BDA0002518098210000063
wherein n isosAnd
Figure BDA0002518098210000064
is unit vector coordinates under the inertial system:
Figure BDA0002518098210000065
and
according to the coordinate R of the satellite in the J2000 coordinate system and the transformation matrix M from the inertial system to the J2000 coordinate system1The calculation can obtain:
Figure BDA0002518098210000066
then, according to the vector relationship, it can be obtained:
OGt＝OS+SGt,
for the purpose of calculation, the coordinates of each vector in the following equation of the J2000 coordinate system are used, where:
Figure BDA0002518098210000071
wherein the content of the first and second substances,
Figure BDA0002518098210000072
Figure BDA0002518098210000073
and
Figure BDA0002518098210000074
wherein M is1Is a transformation matrix of the inertial system to the J2000 coordinate system, M2Is a transformation matrix of the inertial system to the body coordinate system of the observation state,
Figure BDA0002518098210000075
while
Figure BDA0002518098210000076
The coordinates of the vector from the satellite position to the observation point position in the body coordinate system include:
Figure BDA0002518098210000077
in summary, the following results can be obtained: rgeosinΔi＝Rz+fsinγt(ii) a And
finally, two equations obtained in the previous steps are combined, and the rotation angle gamma is solvedtThe value of (c).
In order to implement the observation method in the above embodiment, an embodiment of the present invention further provides a space debris orbit determination traversal observation system, including a camera and a controller, where the camera is installed on a satellite along a satellite + Y axis direction, and the controller is configured to execute the above observation method. In order to realize the observation of synchronous belt space debris, the satellite is arranged on a solar synchronous orbit in the morning and evening, and a quasi-inertial system of the satellite is set according to the orbital characteristics, so that the + Y-axis reference of the satellite points to the normal line down-light direction of an orbital plane to determine the reference attitude of the satellite. To meet the observation requirements, the field of view of the camera is not less than 2 ° x 2 °.
While various embodiments of the present invention have been described above, it should be understood that they have been presented by way of example only, and not limitation. It will be apparent to persons skilled in the relevant art that various combinations, modifications, and changes can be made thereto without departing from the spirit and scope of the invention. Thus, the breadth and scope of the present invention disclosed herein should not be limited by any of the above-described exemplary embodiments, but should be defined only in accordance with the following claims and their equivalents.

Claims (10)

1. A space debris orbit determination traversal observation method is used for observing space debris through a camera installed on a satellite, and is characterized by comprising the following steps:
setting satellite scanning parameters according to observation requirements, wherein the parameters comprise a starting deflection angle theta, an ending deflection angle and a scanning speed;
controlling the satellite to rotate around the satellite body and the X axis so that the visual axis of the camera is perpendicular to the earth synchronous belt at the specified latitude;
controlling the satellite to rotate around a satellite body and a Z axis so that an included angle between a visual axis of the camera and a zero position of the satellite body and a Y axis is equal to the starting deflection angle; and
and opening the camera for observation, and simultaneously controlling the satellite to rotate around the satellite body and the Z axis until the included angle between the visual axis of the camera and the zero position of the satellite body and the Y axis is equal to the ending deflection angle.
2. The method of claim 1, wherein the scan parameters further comprise an observation start time.
3. The method of claim 1, wherein the ending deflection angle is of a size- θ.
4. The method of claim 1, wherein the satellite rotates about the satellite body + Z axis at the scan rate when the camera is observing.
5. The method of claim 1, further comprising controlling satellite rotation about satellite body + X axis by an angle γ while the camera is observingtSuch that the camera's visual axis is always perpendicular to the geosynchronous belt at a specified latitude.
6. The method of claim 5, wherein the satellite rotates about the satellite body + X axis by an angle γtAccording to the satellite S and the intersection point G of the current camera visual axis of the satellite and the geosynchronous belt GEOtAnd calculating the formed vector relation and geometric relation between the geocentrics O, wherein:
the vector relationship is expressed as follows:
OGt＝OS+SGt(ii) a And
the geometric relationship is represented as:
Figure FDA0002518098200000021
wherein the content of the first and second substances,
|OGt|＝Rgeo；
Figure FDA0002518098200000022
wherein [ R ]xRyRz]TCoordinates of the satellite in a J2000 coordinate system; and
Figure FDA0002518098200000023
wherein the content of the first and second substances,
Figure FDA0002518098200000024
and
Figure FDA0002518098200000025
wherein M is1Is a transformation matrix of the inertial system to the J2000 coordinate system.
7. The method of claim 6, wherein the vector relationship is calculated using coordinate values in the J2000 coordinate system:
Figure FDA0002518098200000026
wherein Δ i is the latitude of the geosynchronous belt;
Figure FDA0002518098200000027
and
Figure FDA0002518098200000028
wherein M is1Is a transformation matrix of the inertial system to the J2000 coordinate system, M2Is a transformation matrix of the inertial system to the body coordinate system of the observation state,
Figure FDA0002518098200000029
8. the method of claim 1, wherein the exposure time is matched according to the scan speed when the camera is viewing.
9. A space debris observation system, comprising:
the camera is not less than 2 degrees multiplied by 2 degrees in the field of view and is arranged on the satellite along the Y-axis direction of the satellite; and
a controller configured to perform the method of one of claims 1 to 8.
10. The system of claim 9, wherein the satellite is disposed in a morning and evening sun synchronous orbit with the + Y axis reference of the satellite pointing downwind of the orbital plane normal.
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CN113619813A (en) * 2021-09-17 2021-11-09 中国科学院微小卫星创新研究院 High-orbit space debris fast traversal space-based optical observation system and method
CN113619813B (en) * 2021-09-17 2022-08-26 中国科学院微小卫星创新研究院 High-orbit space debris fast traversing space-based optical observation system and method
CN114355962A (en) * 2021-12-09 2022-04-15 北京航空航天大学 Close-range smooth approaching and maintaining control method for fuel optimization under time constraint
CN114355962B (en) * 2021-12-09 2024-05-14 北京航空航天大学 Near-distance smooth approaching and maintaining control method for fuel optimization under time constraint
CN115610704A (en) * 2022-09-27 2023-01-17 哈尔滨工业大学 Orbital transfer method, device and medium capable of realizing grazing flight observation task on orbit
CN115610704B (en) * 2022-09-27 2023-09-29 哈尔滨工业大学 Rail changing method, device and medium capable of realizing glancing observation task on rail
CN115563437A (en) * 2022-10-11 2023-01-03 中国人民解放军63921部队 Three-dimensional sensing method for GEO space debris by sun synchronous orbit observation platform
CN115687847A (en) * 2022-10-11 2023-02-03 中国人民解放军63921部队 Common-scan sensing method for GEO space debris by low-orbit observation platform
CN115683090A (en) * 2022-10-11 2023-02-03 中国人民解放军63921部队 Method for observing GEO band target in full coverage mode through multiple observation platforms on sun synchronous track in morning and evening
CN115583369A (en) * 2022-10-11 2023-01-10 中国人民解放军63921部队 Ubiquitous perception observation method for GEO space debris by low-orbit multi-observation platform

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