calculating classical orbital elements

08032016, 01:13 PM
Post: #1




calculating classical orbital elements
This HP Prime program (demo_eci2orb1) demonstrates a numerical method that can be used to convert Earthcenterinertial (ECI) position and velocity vectors to classical orbital elements.
This program and its support routines were ported from the Orbital Mechanics with MATLAB software suite. The PDF documentation for this suite is included in the attached zip file. Information about classical orbital elements can be found in the appendix. I plan to eventually port many of the MATLAB routines described in the PDF to the HP Prime. Here's the screen output created by this program. demo_eci2orb1 eci position vector (km) r_x = −5864.79273288 r_y = −1781.73078828 r_z = −2156.29990858 eci velocity vector (km/sec) v_x = 0.492973713131 v_y = −7.31828662424 v_z = 8.19193017342 classical orbital elements semimajor axis = 223705.147138 km eccentricity = 0.971280418487 inclination = 51.9434659015 degrees argument of perigee = 347.329965385 degrees raan = 212.885279071 degrees true anomaly = 347.743436481 degrees 

08032016, 05:51 PM
Post: #2




RE: calculating classical orbital elements
Thanks for all the updates about your programs cdeaglejr!


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