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Orbital elements of satellites in scenario


    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.

    For more information regarding orbital elements, see Orbital Elements.

    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.


    The ephemeris propagator has these fields:

    • EphemerisStartTime

    • EphemerisStopTime

    • PositionTimeTable

    • VelocityTimeTable

    Introduced in R2021a