For this assignment, you will write a Matlab function that takes a reduced set of orbit elements as an input argument and outputs the simulated trajectory for a satellite with those orbit elements in an Earth Centered Earth Fixed (ECEF) reference frame. You will call the function you have written to propagate a satellites position for each of several different orbits. You will use the data that you generate with another function that I will provide on the course web site to plot the ground tracks of the simulated spacecraft trajectories. To get started, write a Matlab function called: function [lat, lon, h] = circsatpos(a, i, Omega, t) that takes semimajor axis ( a in km), inclination ( i in radians), RAAN (Omega in radians), and time ( t in seconds) and computes the Earth Centered Earth Fixed position coordinates latitude ( lat in radians), longitude ( lon in radians), and altitude ( h in kilometers) for a satellite in a circular orbit ( e = 0) about the spherical Earth ( R e = 6378 km). It is assumed that the ECEF and Earth Centered Inertial frames are aligned at t = 0. Additionally, the satellite starts out at the position of the ascending node as specified by the value of Omega. You will call your function repeatedly in a programming loop to compute the satellite position at each minute (60 seconds) for a full solar day (86400 seconds) to build up vectors that are 1440 elements long for each of the coordinates lat, lon, and h.

A. International Space Station orbit, a = 6778 km, i = 51 . 6 degrees, Omega = radians