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State vectors R & V from Orbital Elements

In this example we will show how to compute state vectors R and V in the geo-centric equatorial frame of reference using orbital elements

```clear all;
clc;
% Six  orbital elements are:
h    = 82000;      % [km^2/s] Specific angular momentum
i    = 50;         % [deg] Inclination
RAAN = 60;         % [deg] Right ascension (RA) of the ascending node
e    = 0.2;        % Eccentricity
omega= 90;         % [deg] Argument of perigee
theta= 35;         % [deg] True anomaly
mu = 398600;       % Earth’s gravitational parameter [km^3/s^2]
% Components of the state vector of a body relative to its perifocal
% reference
rx = h^2/mu*(1/(1 + e*cosd(theta)))*[cosd(theta);sind(theta);0];
vx = mu/h*[-sind(theta); (e +cosd(theta));0];
% Direction cosine matrix
QXx = [cosd(omega), sind(omega),0;-sind(omega),cosd(omega),0;0,0,1]*...
[1,0,0;0,cosd(i),sind(i);0,-sind(i),cosd(i)]*...
[cosd(RAAN), sind(RAAN),0;-sind(RAAN),cosd(RAAN),0;0,0,1];
% Transformation Matrix
QxX = inv(QXx);
% Geocentric equatorial position vector R
R = QxX*rx;
% Geocentric equatorial velocity vector V
V = QxX*vx;
fprintf('R = %4.2f*i +  %4.2f*j + %4.2f*k [km]\n',R);
fprintf('V = %4.4f*i +  %4.4f*j + %4.4f*k [km/s]\n',V);```
```R = -10766.25*i +  -3383.89*j + 9095.35*k [km]
V = -0.9250*i +  -5.1864*j + -2.1358*k [km/s]```