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Method for determining orbit of spacecraft by using relative position increment

A technology of relative position vector and relative position, which is applied in the direction of integrated navigator, navigation calculation tool, etc., to achieve the effect of simple orbit determination method, improved autonomy and easy realization

Active Publication Date: 2016-11-09
HARBIN INST OF TECH
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  • Summary
  • Abstract
  • Description
  • Claims
  • Application Information

AI Technical Summary

Problems solved by technology

[0003] The present invention aims to solve the problem that other orbit determination methods need to be used to obtain initial position information in the existing orbit determination method based on measurement epoch difference, and now provides a spacecraft relative position incremental orbit determination method

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  • Method for determining orbit of spacecraft by using relative position increment
  • Method for determining orbit of spacecraft by using relative position increment
  • Method for determining orbit of spacecraft by using relative position increment

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specific Embodiment approach 1

[0012] Specific implementation mode one: refer to image 3 Specifically illustrate the present embodiment, the relative position incremental orbit determination method of the spacecraft described in the present embodiment, the method is:

[0013] First, determine the orbital plane of the spacecraft moving in three-dimensional space;

[0014] Then, transform the three-dimensional relative position vector into the relative position coordinates of the spacecraft in a two-dimensional plane, use the least square principle to fit the shape of the orbit of the spacecraft in the two-dimensional plane, and use the Lagrangian multiplier method to solve the eigenvalues ​​of the fitted matrix;

[0015] Finally, the orbit type of the spacecraft is judged according to the eigenvalues ​​of the fitting matrix. If the orbit of the spacecraft is an ellipse, the central celestial body is located at one focus of the ellipse, and the central celestial body is determined by combining the Kepler ti...

specific Embodiment approach 2

[0016] Specific implementation mode two: refer to figure 1 This embodiment is described in detail. This embodiment is a further description of the relative position incremental orbit determination method of the spacecraft described in the first embodiment. In this embodiment, the specific method for determining the orbital plane of the spacecraft moving in three-dimensional space for:

[0017] Measure the spacecraft at the epoch time t in the earth-centered inertial coordinate system j with t j-1 The relative position vector Δr(t j-1 ,t j ), and according to the relative position vector Δr(t j-1 ,t j ) to obtain the initial position of the spacecraft r(t 0 ), where j=1,2,3,...;

[0018] According to the relative position vector Δr(t j-1 ,t j ), using the principle of least squares to determine the unit angular momentum vector h;

[0019] Δr x ...

specific Embodiment approach 3

[0029] Specific embodiment three: This embodiment is to further explain the relative position incremental orbit determination method of the spacecraft described in the second specific embodiment. In this embodiment, the three-dimensional relative position vector is converted into the relative position of the spacecraft in the two-dimensional plane The specific method of coordinates is:

[0030] Set reference frame P xyz Taking the initial position of the spacecraft as the origin, the z-axis is perpendicular to the orbital plane, and the x and y-axes are based on the coordinate transformation matrix Q of the Earth's J2000 reference system Xx get;

[0031] Obtain the relative position of the spacecraft in the reference frame P according to the orbital inclination i and the right ascension of the ascending node Ω xyz The vector x in k :

[0032] x k =Q Xx m k (4)

[0033] In the formula, is the vector of the spacecraft position relative to the initial position at time ...

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Abstract

A method for determining orbit of a spacecraft by using relative position increment belongs to the technical field of determination of spacecraft orbit, and aims at solving the problem that a conventional orbit determination method based on measured epoch difference needs to employ other orbit determination method for obtaining initial position information. The method for determining orbit of the spacecraft by using relative position increment comprises: firstly, determining the motional orbit plane of the spacecraft in a three-dimensional space; then, converting the three-dimensional relative position vector into a spacecraft relative position coordinate in a two-dimensional plane, fitting the spacecraft motion orbit shape in the two-dimensional plane by employing a least square method, and determining eigenvalue of the fit matrix by employing lagrangian multiplier; and finally, determining the spacecraft orbit type according to the eigenvalue of the fit matrix, and obtaining the absolute position information of the spacecraft in an inertia reference system according to the position of a central celestial body, so as to determine the initial orbit of the spacecraft. The method is applicable to orbit determination of spacecraft.

Description

technical field [0001] The invention belongs to the technical field of spacecraft orbit determination. Background technique [0002] The task of spacecraft orbit determination is to use a small amount of observation data to quickly determine the orbit elements under the two-body problem of the spacecraft. Due to the difference in the number of stations and the attributes of measurement quantities (such as position, velocity, angle) in the observation data, the orbit determination methods are also different accordingly. Among them, the position observation can be obtained through the global satellite navigation system or ground radar, and the position information obtained by the measurement can be obtained by simple conversion to obtain the position vector or direction vector of the spacecraft in the earth inertial coordinate system. Using three geocentric position vectors can be obtained by The Gibbs method of orbit determination is the Lambert problem of orbit determinatio...

Claims

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Application Information

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IPC IPC(8): G01C21/20G01C21/24
CPCG01C21/20G01C21/24
Inventor 徐国栋宋佳凝姜昆董立珉徐振东张兆祥
Owner HARBIN INST OF TECH
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