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This HP Prime program calculates the altitude relative to a triaxial ellipsoidal planet or other celestial body. The algorithm is based on the numerical method described in “Geodetic Altitude to a Triaxial Ellipsoidal Planet”, by Charles C. H. Tang, The Journal of the Astronautical Sciences, Vol. 36, No. 3, July-September 1988, pp. 279-283.

The user can provide classical orbital elements of a spacecraft starting at line 56 of the source code. Here is an example orbit.

// classical orbital elements

oev(1) := 8000.0;
oev(2) := 0.015;
oev(3) := 28.5 * dtr;
oev(4) := 120.0 * dtr;
oev(5) := 45.0 * dtr;
oev(6) := 30.0 * dtr;

In this array oev(1) is the semimajor axis in kilometers, oev(2) is the non-dimensional orbital eccentricity, oev(3) is the orbital inclination, oev(4) is the argument of perigee, oev(5) is the right ascension of the ascending node and oev(6) is the true anomaly. All angular orbital elements should be provided in degrees. They are converted to radians by the multiplication of dtr which is the degrees-to-radians conversion factor.

Can you select the proper set of orbital elements to verify the equatorial and polar altitudes?
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