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calculating classical orbital elements - Printable Version

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calculating classical orbital elements - cdeaglejr - 08-03-2016 01:13 PM

This HP Prime program (demo_eci2orb1) demonstrates a numerical method that can be used to convert Earth-center-inertial (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


RE: calculating classical orbital elements - Klaas Kuperus - 08-03-2016 05:51 PM

Thanks for all the updates about your programs cdeaglejr!