Method for determining orbit transferring strategy of time-limited spacecraft coplanar rendezvous

A time-limited and strategy-determined technology, applied to spaceflight vehicles, space vehicle guidance devices, aircraft, etc., can solve the problems of poor observation effect at rendezvous time, affecting the calculation accuracy of rendezvous orbit change strategy, and low lighting conditions. , to achieve the effects of reducing fuel consumption, accurate calculation results, and simple iterative calculation process

Inactive Publication Date: 2019-04-09
HARBIN INST OF TECH
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  • Claims
  • Application Information

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Problems solved by technology

[0004] The purpose of the present invention is to solve the problem that the existing spacecraft rendezvous technology has relatively low requirements on illumination conditions, which leads to poor observation effect at

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  • Method for determining orbit transferring strategy of time-limited spacecraft coplanar rendezvous
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  • Method for determining orbit transferring strategy of time-limited spacecraft coplanar rendezvous

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specific Embodiment approach 1

[0030] A method for determining the time-limited spacecraft rendezvous rendezvous and orbit change strategy in this embodiment, the initial orbit of the spacecraft is a high-orbit orbit within hundreds of kilometers below the geosynchronous orbit, and it will always be in the orbit before and after the orbit change. Operate on the equatorial plane, such an area is convenient to carry out on-orbit service to the target spacecraft, and the method is realized by the following steps:

[0031] Step 1. Determine the initial orbital parameters of the target spacecraft and the service spacecraft;

[0032] Step 2, calculating the theoretical minimum time and maximum time of the entire rendezvous process between the target spacecraft and the service spacecraft;

[0033] Step 3: Considering the constraints of the solar illumination angle, calculate the time spent in the entire rendezvous process between the target spacecraft and the serving spacecraft;

[0034] Step 4. According to the ...

specific Embodiment approach 2

[0042] The difference from Embodiment 1 is that in this embodiment, a time-limited spacecraft rendezvous rendezvous and orbit change strategy determination method, the determination of the initial orbital parameters of the target spacecraft and the serving spacecraft in the first step includes: Orbit half Major axis (a), eccentricity (e), orbital inclination (i), right ascension of ascending node (Ω), argument of perigee (ω) and true anomaly (f); among them,

[0043] Orbital semi-major axis a: Indicates the size of the orbit of the spacecraft, half of the elliptical orbital major axis is determined as the orbital semi-major axis, and the size of the orbital semi-major axis is related to the orbital period;

[0044] Eccentricity e: Indicates the shape of the orbit of the spacecraft. The ratio of the distance between the two focal points of the elliptical orbit and the major axis is determined as the eccentricity. The smaller the eccentricity, the closer the orbit of the spacecra...

specific Embodiment approach 3

[0049] The difference from the first or second embodiment is that in this embodiment, a time-limited spacecraft rendezvous rendezvous and orbit change strategy determination method, in the second step, the calculation of the entire rendezvous process between the target spacecraft and the serving spacecraft The theoretical shortest time and longest time process, specifically:

[0050] Let the angle between the position vector of the serving spacecraft and the position vector of the target spacecraft at the beginning of the process be θ 0 , the angular velocity of the spacecraft in orbit is: Theoretically the shortest time T consumed to complete the entire coplanar rendezvous process min and the maximum time T max They are:

[0051] Among them, μ represents the gravitational constant, usually the value is 398600.4415kg 3 / s 2 ; a represents the semi-major axis of the space orbit; ω ci is the angular velocity of the initial orbit of the serving spacecraft, ω ct The angu...

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Abstract

The invention provides a method for determining an orbit transferring strategy of time-limited spacecraft coplanar rendezvous, and belongs to the field of the time-limited spacecraft coplanar rendezvous. The problem that the accuracy of the rendezvous algorithm calculation is affected caused by the poor observation effect at rendezvous time due to the fact that an existing spacecraft rendezvous technology has relatively low requirements on lighting conditions is solved. The method comprises the steps that the theoretical limiting time of the entire rendezvous process between a target spacecraft and a service spacecraft is calculated; then the two speed increments and the used time of orbit transferring in the service spacecraft orbit transferring process are calculated; initial values of the movement time of an initial orbit and initial values of the movement time of a target orbit for the optimized variable service spacecraft are calculated; integral is carried out on the rendezvous process to obtain position vectors of the service spacecraft and the target spacecraft at the rendezvous time and the angle of the position vectors; when the angle reaches 0 degree, at the moment, themovement time of the initial orbit and the movement time of the target orbit are used as the orbit transferring parameters and the orbit transferring strategy. According to the method, the calculationresults are accurate, and the iterative calculation process is simple.

Description

technical field [0001] The present invention relates to the design of a coplanar rendezvous strategy for spacecraft, in particular to a strategy that provides a feasible limited time interval for the on-orbit service mission planning of a coplanar circular orbit spacecraft, and has high illumination and distance requirements at the end of the rendezvous process. The design of the track change strategy. Background technique [0002] In the coplanar rendezvous problem of geosynchronous orbiting spacecraft, the orbital inclination of the target spacecraft is less than 0.05°. And in the process of co-planar rendezvous, the service spacecraft needs to observe or detect the target spacecraft at a closer distance at the end of the process, so as to complete the on-orbit service process or provide the necessary target spacecraft information for the subsequent on-orbit service process . Therefore, in the coplanar rendezvous process, at the end of the process, it is necessary to ens...

Claims

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Application Information

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IPC IPC(8): B64G1/24
CPCB64G1/242
Inventor 马广富赵广栋郭延宁高泽天邓武东
Owner HARBIN INST OF TECH
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