Journal of Aeronautics & Aerospace Engineering
Open Access
ISSN: 2168-9792
+44-77-2385-9429
ISSN: 2168-9792
+44-77-2385-9429
Aravind Gundakaram, Abhirath Sangala, Aditya Sai Ellendula, Prachi Kansal, Suchir Reddy Punuru, Lanii Lakshitaa, Nethra Naveen, and Sanjitha Jaggumantri
All the authors are students from Mahindra University, Hyderabad, India, pursuing their bachelorâ??s degrees in either Mathematics or Computer Science and Artificial Intelligence.
Scientific Tracks Abstracts: J Aeronaut Aerospace Eng
In this paper, we develop a high precision satellite orbit determination model for satellites orbiting the Earth. Solving this model entails numerically integrating the differential equation of motion governing a two-body system, for which we employ Fehlberg’s formulation of the Runge-Kutta class of numerical integrators with adaptive stepsize control. Various perturbing forces are also accounted for in the mathematical model, such as the acceleration due to the geopotential of the Earth, third-body gravitational effects, solar radiation pressure and atmospheric drag. For applications requiring high precision modelling, we also account for Earth radiation pressure, the perturbative effects of solid-Earth tides and ocean tides, and also make adjustments to the total acceleration for relativistic effects. We have provided explicit expressions to calculate the force exerted by each perturbation that contributes to the total acceleration of the satellite. In situations where calculating certain terms in these expressions poses practical challenges, we have provided recurrence relations to assist with implementation. The implementation of this model yields a satellite orbit propagator, which we call the Satellite Orbit Determiner (SED). We have discussed the architecture of SED and the methodology it employs, and have presented the numerical results obtained from it. These results are compared with the widely used High Precision Orbit Propagator (HPOP). Currently, SED has only been implemented for the two-body problem, but future advancements will enable it to handle the three-body problem as well.
Aravind Gundakaram is a student of Mahindra University pursuing his bachelor’s degree in Computation and Mathematics. His understanding of the mathematical model together with the implementation skills of his team allow them to come up with high precision mathematical models and also effective implementations for the said models. The current satellite orbit propagator presented in the said paper, namely the Satellite Ephemeris Determiner (SED) is the result of over two years of consistent work from the entire team, and guidance from several people working in the industry. The mathematical model in this paper has been built of after a thorough review of existing orbit propagator models and the results obtained from them. Moreover, the very high cost of the popular High Precision Orbit Propagator (HPOP) served as further motivation for the team to develop their own propagator SED, that is equivalent to HPOP. In the future, the team aims to make further advancements in the field and enhance SED to handle the three-body problem as well, so that it can outperform HPOP.