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Satellite flight-tracking secular perturbation compensation method based on on-orbit parameter identification and offset

A technology of parameter identification and compensation method, which is applied in the field of micro-satellite formation, can solve the problems of long-term natural stability of unfavorable formation, large control fuel consumption, high control frequency, etc., and achieve the effects of improving flexibility, reducing fuel load, and controlling low fuel consumption

Active Publication Date: 2017-09-15
TSINGHUA UNIV
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  • Application Information

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Problems solved by technology

Due to the error of the current measurement data and the inaccuracy of the relative motion model, the repeated boundary control method used in the on-orbit project has the disadvantages of large control fuel consumption and high control frequency, which is not conducive to the long-term natural stability of the formation

Method used

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  • Satellite flight-tracking secular perturbation compensation method based on on-orbit parameter identification and offset
  • Satellite flight-tracking secular perturbation compensation method based on on-orbit parameter identification and offset
  • Satellite flight-tracking secular perturbation compensation method based on on-orbit parameter identification and offset

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specific Embodiment

[0094] Set the initial conditions as shown in Table 1.

[0095] Table 1 Initial orbital parameters of target star and tracking star

[0096]

a

e

i

Ω

ω

M 0

Goal star Shungen

6961.181km

0.001679

97.622°

262.154°

265.810°

323.848°

Tracking Xing Shun Gen

6961.086km

0.001726

97.622°

262.188°

263.930°

325.152°

[0097] Among them, a is the semi-major axis of the orbit, e is the eccentricity, i is the orbital inclination, Ω is the right ascension of the ascending node, M 0 is the mean anomaly angle at the initial moment.

[0098] Input the initial conditions into STK, and use STK to perform high-precision simulation, the steps are as follows:

[0099] step one:

[0100] Give the on-orbit identification time period [t 0 t f ] is [0,0.3] days, according to the STK data, the relative drift rate H along the track angle is

[0101] H=6.6008×10 -8 (rad / s) (37)

[0102] Step two:

[0103] Given ...

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Abstract

The invention provides a satellite flight-tracking secular perturbation compensation method based on on-orbit parameter identification and offset, and the method comprises the following steps: 1, carrying out the on-orbit identification of an along-track angle relative drift rate; 2, calculating an orbit semi-major axis offset of a tracking satellite at a given control moment based on the along-track angle relative drift rate obtained at step 1; 3, giving an orbit semi-major axis of the tracking satellite at the control moment, and obtaining the velocity increment, needed by the orbit semi-major axis offset control, of the tracking satellite according to an orbit dynamical model. The method is advantageous in that the (1), the method has no requirements for the control time, and improves the flexibility of formation keeping control; (2), the method is low in control burnup, and effectively reduces the fuel load of a satellite formation; (3), the method can achieve the long-time natural maintaining through one control operation, and reduces the control frequency.

Description

technical field [0001] The invention belongs to the technical field of micro-satellite formation, and in particular relates to a long-term perturbation compensation method for satellite follow-up based on on-orbit parameter identification and offset. Background technique [0002] In order to realize space missions, especially long-time follow-up missions, the key to the formation of satellites is to keep the relative configuration of the follow-up satellites participating in the formation accurately. However, the follow-up satellite is affected by various perturbations and control errors during the actual on-orbit operation, and the relative motion state presents a divergent trend. It requires configuration maintenance control to maintain its relative configuration, which brings challenges to long-term on-orbit work. To this end, it is necessary to carry out research on the perturbation law of satellite follow-up, and achieve long-term stability of the relative motion of fol...

Claims

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Application Information

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Patent Type & Authority Applications(China)
IPC IPC(8): G05D1/10
CPCG05D1/104
Inventor 王兆魁蒋超范丽李泰博
Owner TSINGHUA UNIV
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