Most readily available Keplerian orbital element conversion utilities do not address circular or parabolic orbits. This set of routines will address the complete spectrum of orbits from a circular equatorial orbit to a retrograde hyperbolic orbit without error. All functionality is vectorized for computational efficiency.
Example Function Call One:
>> [r_ECI v_ECEF] = orb2rv(p,e,i,O,o,nu);
p = semilatus rectum (km) [1 x N]
e = Eccentricity [1 x N]
i = Inclination (rad) [1 x N]
O = Right Ascension of the Ascending Node (rad) [1 x N]
o = Argument of Perigee (rad) [1 x N]
nu = True Anomaly (rad) [1 x N]
Example Function Call Two:
>>[a,e,i,O,o,nu] = rv2orb(r_ECI,v_ECI)
r = Position State Vector in km (ECI) [3 x N]
v = Velocity State Vector in km/s (ECI) [3 x N]
For those orbits which are equatorial or circular, the following full form function calls are necessary:
Convert state vector to full set of orbital elements:
>>[a,e,i,O,o,nu,truLon,argLat,lonPer,p] = rv2orb(r_ECI,v_ECI);
Convert full set of orbital elements back to a state vector:
>>[r_ECI,v_ECI] = orb2rv(p,e,i,O,o,nu,truLon,argLat,lonPer);
truLon = True Longitude (rad) [1 x N]
argLat = Argument of Latitude (rad) [1 x N]
Possibly mu is 3.986004418*10^14; Standard Gravitational Parameter for Earth
Good, but getConst() is not a function in Octave, so have to replace it with the numerical value of mu.
oops, forgot the comment, here it goes:
was about to write eci conversion functions myself, glad i've found these ones first!
Updated outputs for orbital parameters when state vector is circular and equatorial.
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