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Convert between state vectors and orbital elements

  1. Apr 5, 2013 #1
    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
     
  2. jcsd
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