keplerian_elements

Struct Orbit

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pub struct Orbit {
    pub eccentricity: f32,
    pub semi_major_axis: f32,
    pub inclination: f32,
    pub longitude_of_ascending_node: f32,
    pub argument_of_periapsis: f32,
}

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§eccentricity: f32§semi_major_axis: f32§inclination: f32§longitude_of_ascending_node: f32§argument_of_periapsis: f32

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impl Orbit

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pub fn state_vectors_to_orbit( state_vectors: StateVectors, central_body_mass: f32, ) -> Self

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pub fn standard_gravitational_parameter(mass: f32) -> f32

https://en.wikipedia.org/wiki/Standard_gravitational_parameter

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pub fn period(&self, mass: f32) -> f32

https://en.wikipedia.org/wiki/Orbital_period

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pub fn mean_motion(&self, mass: f32) -> f32

https://en.wikipedia.org/wiki/Mean_motion

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pub fn mean_anomaly(&self, mass: f32, time: f32) -> f32

https://en.wikipedia.org/wiki/Mean_anomaly

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pub fn estimate_eccentric_anomaly( &self, mass: f32, time: f32, tolerance: f32, ) -> f32

https://en.wikipedia.org/wiki/Eccentric_anomaly

Eccentric Anomaly (EA) is given by the equation: M = EA - e*sin(EA) where M is the mean anomaly e is the eccentricity

To estimate the Eccentric Anomaly we use Newton’s method https://en.wikipedia.org/wiki/Newton%27s_method

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pub fn state_vectors_at_epoch( &self, mass: f32, time: f32, tolerance: f32, ) -> StateVectors

Auto Trait Implementations§

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impl Freeze for Orbit

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impl RefUnwindSafe for Orbit

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impl Send for Orbit

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impl Sync for Orbit

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impl Unpin for Orbit

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impl UnwindSafe for Orbit

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impl<T> Any for T
where T: 'static + ?Sized,

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