2024年4月26日发(作者：资颐和)
matlab 椭圆轨道
English Answer:
Orbital Elements of an Elliptical Orbit.
An elliptical orbit is a Keplerian orbit that has a
non-zero eccentricity. It is characterized by the following
orbital elements:
Semi-major axis (a)。
Eccentricity (e)。
Inclination (i)。
Longitude of the ascending node (Ω)。
Argument of periapsis (ω)。
True anomaly (ν)。
The semi-major axis is the mean of the perihelion and
aphelion distances. The eccentricity is a measure of how
elongated the orbit is, with a value of 0 indicating a
circular orbit and a value of 1 indicating a parabolic
orbit. The inclination is the angle between the orbital
plane and the reference plane. The longitude of the
ascending node is the angle between the direction of the
vernal equinox and the direction of the ascending node. The
argument of periapsis is the angle between the direction of
the ascending node and the direction of periapsis. The true
anomaly is the angle between the direction of periapsis and
the current position of the orbiting body.
Calculating the Orbital Elements.
The orbital elements of an elliptical orbit can be
calculated using the following equations:
a = (r_p + r_a) / 2。
e = (r_a r_p) / (r_a + r_p)。
2024年4月26日发(作者：资颐和)
matlab 椭圆轨道
English Answer:
Orbital Elements of an Elliptical Orbit.
An elliptical orbit is a Keplerian orbit that has a
non-zero eccentricity. It is characterized by the following
orbital elements:
Semi-major axis (a)。
Eccentricity (e)。
Inclination (i)。
Longitude of the ascending node (Ω)。
Argument of periapsis (ω)。
True anomaly (ν)。
The semi-major axis is the mean of the perihelion and
aphelion distances. The eccentricity is a measure of how
elongated the orbit is, with a value of 0 indicating a
circular orbit and a value of 1 indicating a parabolic
orbit. The inclination is the angle between the orbital
plane and the reference plane. The longitude of the
ascending node is the angle between the direction of the
vernal equinox and the direction of the ascending node. The
argument of periapsis is the angle between the direction of
the ascending node and the direction of periapsis. The true
anomaly is the angle between the direction of periapsis and
the current position of the orbiting body.
Calculating the Orbital Elements.
The orbital elements of an elliptical orbit can be
calculated using the following equations:
a = (r_p + r_a) / 2。
e = (r_a r_p) / (r_a + r_p)。