# Convert between state vectors and orbital elements

1. Apr 5, 2013

### napst3r

Hi there Ive been struggling to convert the following two vectors into orbital elements,
the vectors are as follows:

R = 0i - 7950j + 0k km
V = 5i + 0j + 5k km/sec

I would like someone to help with the calculations of the following COEs:
- semimajor axis (a)
- eccentricity vector
- eccentricity (e)
- inclination (i)
- RAAN (Ω)
- argument of perigee (ω)
- true anomaly (v)

I have tried to understand the following link but im not getting the answers i should be getting
http://www.cdeagle.com/omnum/pdf/csystems.pdf

Thanks