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orbitalElements

Orbital elements of satellites in scenario

    Description

    example

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

    Examples

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    Create a satellite scenario object.

    sc = satelliteScenario;

    Add a satellite to the scenario.

    tleFile = "eccentricOrbitSatellite.tle";
    sat1 = satellite(sc,tleFile);

    Retrieve the orbital elements of sat1.

    elements1 = orbitalElements(sat1)
    elements1 = struct with fields:
                           MeanMotion: 0.0083
                         Eccentricity: 0.7415
                          Inclination: 60.0000
        RightAscensionOfAscendingNode: 30.0000
                  ArgumentOfPeriapsis: 280
                          MeanAnomaly: 289.4697
                               Period: 43200
                                Epoch: 05-May-2020 13:51:55
                                BStar: 0
    
    

    Add a satellite from Keplerian elements to the scenario.

    semiMajorAxis = 6878137;                               % meters
    eccentricity = 0;
    inclination = 20;                                      % degrees
    rightAscensionOfAscendingNode = 0;                     % degrees
    argumentOfPeriapsis = 0;                               % degrees
    trueAnomaly = 0;                                       % degrees
    sat2 = satellite(sc,semiMajorAxis,eccentricity, ...
           inclination,rightAscensionOfAscendingNode, ...
           argumentOfPeriapsis,trueAnomaly, ...
           "OrbitPropagator","two-body-keplerian", ...
           "Name","Sat2");

    Retrieve the orbital elements of sat2.

    elements2 = orbitalElements(sat2)
    elements2 = struct with fields:
                        SemiMajorAxis: 6878137
                         Eccentricity: 0
                          Inclination: 20
        RightAscensionOfAscendingNode: 0
                  ArgumentOfPeriapsis: 0
                          TrueAnomaly: 0
                               Period: 5.6770e+03
    
    

    Input Arguments

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    Satellite object, specified as a scalar.

    Output Arguments

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

    For more information on orbital elements, see Orbital Elements.

    The orbital elements are defined in the Geocentric Celestial Reference Frame (GCRF).

    FieldDescription
    SemiMajorAxisSemimajor axis, in meters
    EccentricityEccentricity
    InclinationInclination angle, in degrees
    RightAscensionOfAscendingNodeRight ascension of Ascending node, in degrees
    ArgumentOfPeriapsisArgument of periapsis, in degrees
    TrueAnomalyTrue anomaly at reference time, in degrees
    PeriodOrbital period, in seconds

    Data Types: struct

    The orbital elements represent general perturbation mean elements.

    FieldDescription
    MeanMotionGeneral perturbation mean motion, in degrees per second
    EccentricityMean eccentricity
    InclinationMean inclination angle, in degrees
    RightAscensionOfAscendingNodeMean right ascension of Ascending node, in degrees
    ArgumentOfPeriapsisMean argument of periapsis, in degrees
    MeanAnomalyMean mean anomaly at the reference time, in degrees
    EpochEpoch
    BStarDrag term, in per EarthRadius
    PeriodMean orbital period, in seconds

    Data Types: struct

    FieldDescription
    EphemerisStartTimeEphemeris start time
    EphemerisStopTimeEphemeris stop time
    PositionTimeTablePosition timetable
    VelocityTimeTableVelocity timetable

    Data Types: struct

    The orbital elements are derived from the SEM almanac file or RINEX GPS navigation message, and defined in the Earth-Centered-Earth-Fixed (ECEF) frame.

    FieldDescription
    PRNPseudorandom noise number
    SatelliteHealthSatellite health
    GPSWeekNumberGPS week number, with time of ephemeris.
    GPSTimeOfApplicabilityTime of Applicability, in seconds
    SemiMajorAxisSemimajor axis, in meters
    EccentricityEccentricity
    InclinationInclination angle at reference time, in degrees
    GeographicLongitudeOfOrbitalPlaneLongitude of ascending node of orbit plane at weekly epoch, in degrees
    RateOfRightAscensionReference rate of right ascension, in degrees per second
    ArgumentOfPerigeeArgument of perigee, in degrees
    MeanAnomalyMean anomaly at the reference time, in degrees
    PeriodOrbital period, in seconds

    Data Types: struct

    The orbital elements are derived from the RINEX Galileo navigation message, and defined in the Earth-Centered-Earth-Fixed (ECEF) frame.

    FieldDescription
    SatelliteIDSatellite system number
    GALWeekNumberGAL week number
    TimeofEphemerisTime of ephemeris, in seconds
    SatelliteHealthSatellite health
    SemiMajorAxisSemimajor axis, in meters
    EccentricityEccentricity
    InclinationInclination angle at reference time, in degrees
    GeographicLongitudeOfOrbitalPlaneLongitude of ascending node of orbit plane at weekly epoch, in degrees
    RateOfRightAscensionReference rate of right ascension, in degrees per second
    ArgumentOfPerigeeArgument of perigee, in degrees
    MeanAnomalyMean anomaly at the reference time, in degrees
    PeriodOrbital period, in seconds

    Data Types: struct

    Data Types: struct

    Version History

    Introduced in R2021a