Orbit and State Reference¶

In ANISE, an Orbit represents the full state of an object at a specific time within a specific coordinate frame.

The `Orbit` Struct¶

An Orbit consists of:

State Vector: Position and Velocity (\(\mathbf{r}, \mathbf{v}\)).
Epoch: The time at which the state is defined (using hifitime).
Frame: The reference frame in which the vectors are expressed.

Initialization¶

You can initialize an Orbit from various representations, including with the following initializers. Refer to https://docs.rs/anise/latest/anise/astro/orbit/type.Orbit.html for the exhaustive list, which includes Keplerian altitudes, Keplerian apses radii, Keplerian mean anomaly, etc.

Cartesian¶

Orbit::cartesian(x, y, z, vx, vy, vz, epoch, frame) Directly sets the position and velocity in km and km/s.

Keplerian¶

Orbit::keplerian(sma, ecc, inc, raan, aop, true_anomaly, epoch, frame) Calculates the state from classical orbital elements. - Note: Requires the frame to have a defined gravitational parameter (\(\mu\)).

Geodetic / LatLongAlt¶

Orbit::try_latlongalt(lat, lon, alt, epoch, frame) Defines a state relative to a body's surface. Useful for ground stations.

Common Operations¶

`transform_to`¶

almanac.transform_to(orbit, target_frame, aberration) Returns a new Orbit expressed in the target_frame. This accounts for both translation and rotation of the frames.

Orbital Elements¶

You can extract orbital elements from an Orbit object:

sma_km(), ecc(), inc_deg(), raan_deg(), aop_deg(), ta_deg()
Other Anomalies: ma_deg() (Mean Anomaly), ea_deg() (Eccentric Anomaly).
Other Parameters: periapsis_km(), apoapsis_km(), hmag() (Specific Angular Momentum).
Python: Accessible as methods (e.g., orbit.sma_km()).

Propagation (Two-Body)¶

ANISE provides basic two-body propagation for quick estimates:

at_epoch(new_epoch): Propagates the orbit to a new time using Keplerian motion.
Warning: For high-fidelity propagation including J2, solar radiation pressure, etc., use a dedicated propagator like those found in the Nyx library.