True Anomaly (ν):
The next step is to calculate the satellites True Anomaly ν, which is the angle between Perihelion (Perigee) and the actual position of the satellite measured with respect to the center of the ECI coordinate frame (The center of Earth), as shown in the following figure.
It is determined out from the satellites current Mean Anomaly M_t and the Orbital Eccentricity ecc.
ν = M_t + 2 ecc sin(M_t) + (5/4)(ecc²)(sin(2M_t))

Argument of Latitude (Lat):
The Argument of Latitude Lat represents the angle between the line of nodes and the current position of the satellite with respect to the center of the ECI coordinate frame (The center of Earth). It is calculated as the sum of the Argument of Perigee ω and the True Anomaly ν.

Lat = ω + ν

Position Vector in ECI Coordinate Frame (x_ECI, y_ECI, z_ECI):
Based on the Argument of Latitude Lat, the Right Ascension of Ascending Node RAAN and the Inclination of the satellites orbit Inc, the position of the satellite can be determined as a unit vector in the ECI coordinate frame, by using the following equation.

xECI = cos(Lat)cos(RAAN) - sin(Lat)sin(RAAN)cos(Inc) yECI = cos(Lat)sin(RAAN) + sin(Lat)cos(RAAN)cos(Inc) zECI = sin(Lat)sin(Inc)

Orbital Radius (r_SatPos):
To calculate the orbital position of the satellite in kilometers, the position vector given in the last equation (x_ECI, y_ECI, z_ECI) is multiplied with the radius of the satellites orbital position r_SatPos (in Kilometers) as shown in the next equation.

r_SatPos = A_km [1 - ecc²]/[1 + ecc cos(ν)];