To compute an elliptical spacecraft orbit in MATLAB, we need to follow these steps:
mu
of the central body around which the spacecraft is orbiting.a
and the eccentricity e
of the spacecraft's orbit.b
of the spacecraft's orbit using the following formula: b = a * sqrt(1 - e^2)
.i
, the longitude of ascending node Omega
, and the argument of periapsis omega
of the spacecraft's orbit.M
of the spacecraft's orbit.E
of the spacecraft's orbit using the following iterative formula: E = M + e * sin(E)
.nu
of the spacecraft's orbit using the following formula: nu = 2 * atan(sqrt((1 + e) / (1 - e)) * tan(E / 2))
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Here's the MATLAB code that implements these steps:
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Note that this code assumes a circular central body, which is a good approximation for orbits around Earth. If the central body is not spherical or has significant oblateness, more complex calculations may be necessary.
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