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orbitalElements

Orbital elements of satellites in scenario

    Description

    elements = orbitalElements(sat) returns the orbital elements of the specified satellite sat.

    Input Arguments

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    Satellite, specified as a row vector of Satellite objects.

    Output Arguments

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    Orbital elements of input sat, returned as a structure. The fields of the structure depend on the orbit propagator chosen using the OrbitPropagator property of the satelliteScenario object.

    Two Body Keplerian

    The two-body-keplerian orbit propagator has these fields:

    • SemiMajorAxis

    • Eccentricity

    • Inclination

    • RightAscensionOfAscendingNode

    • ArgumentOfPeriapsis

    • TrueAnomaly

    • Period

    SGP4 and SDP4

    The sgp4 and sdp4 orbit propagators have these fields:

    • Eccentricity

    • Inclination

    • RightAscensionOfAscendingNode

    • ArgumentOfPeriapsis

    • MeanAnomaly

    • MeanMotion

    • Epoch

    • BStar

    • Period

    The orbital elements represent the mean values at Epoch.

    Ephemeris

    The ephemeris propagator has these fields:

    • EphemerisStartTime

    • EphemerisStopTime

    • PositionTimeTable

    • VelocityTimeTable

    Introduced in R2021a