Increasing the Accuracy of Orbital Elements for a Satellite in a Low Earth Orbit under the Influence of Atmospheric Drag Using Adams-Bashforth Method

Rasha H Ibrahim,Abdul-Rahman H Saleh

doi:10.24996/ijs.2021.si.2.9

Abstract

The perturbed equation of motion can be solved by using many numerical methods. Most of these solutions were inaccurate; the fourth order Adams-Bashforth method is a good numerical integration method, which was used in this research to study the variation of orbital elements under atmospheric drag influence. A satellite in a Low Earth Orbit (LEO), with altitude form perigee = 200 km, was selected during 1300 revolutions (84.23 days) and ASat / MSat value of 5.1 m2/ 900 kg. The equations of converting state vectors into orbital elements were applied. Also, various orbital elements were evaluated and analyzed. The results showed that, for the semi-major axis, eccentricity and inclination have a secular falling discrepancy, Longitude of Ascending Node is periodic, Argument of Perigee has a secular increasing variation, while true anomaly grows linearly from 0 to 360°. Furthermore, all orbital elements, excluding Longitude of Ascending Node, Argument of Perigee, and true anomaly, were more affected by drag than other orbital elements, through their falling as the time passes. The results illustrate a high correlation as compared with literature reviews in this field.

Highlights

The main condition in orbit’s determination is to find state vectors and orbital elements
In the presence of perturbations, the ideal path of a satellite will vary from the hypothetical two-body problem [3]. They are classified into two categories: gravitational and non-gravitational. Gravitational perturbations include those of the non-spherical Earth and the third body attraction, whereas non-gravitational perturbations involve those of the atmospheric drag and solar radiation pressure
Mishra et al elaborated the method of determining a precise ephemeris for an orbiting satellite due to atmospheric drag

Summary

Introduction

The main condition in orbit’s determination is to find state vectors and orbital elements. In the presence of perturbations, the ideal path of a satellite will vary from the hypothetical two-body problem [3] They are classified into two categories: gravitational and non-gravitational. The process of orbit determination is based on the predication of the satellite’s drift, which in turn is used to estimate the future orbit [4, 5, 6] This process occurs by a numerically integrated perturbed equation of motion, using one of the integration methods, such as. Mishra et al elaborated the method of determining a precise ephemeris for an orbiting satellite due to atmospheric drag. This method includes guessing of state vectors from a sequence of data [9]. Thangavel et al studied the effects of atmospheric drag and J2 in a closeproximity operation for LEO [11]

Methods

Conclusions