Eccentricity vector (3D)
[eccv] = CL_op_eccVector3d(pos, vel [, mu])
Computes the eccentricity vector from the inertial position and velocity vectors.
The eccentricity vector points to the periapsis of the orbit, and its norm is equal to the orbit's eccentricity.
If pos or vel is [], the result is [].
Notes:
- The origin of the implicit frame is the center of the central body (coordinales = [0;0;0]).
- This function can be used for elliptical or hyperbolic orbits.
Satellite's position vector [m] (3xN or 3x1)
Satellite's velocity vector [m/s] (3xN or 3x1)
(optional) Gravitational constant. Default: %CL_mu. [m^3/s^2] (1x1). (1x1)
Eccentricity vector (3xN)
CNES - DCT/SB