Main Content

orbitalElements

Orbital elements of satellites in scenario

    Description

    example

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

    Examples

    collapse all

    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: 1.4544e-04
                         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

    collapse all

    Satellite, specified as a row vector of Satellite objects.

    Output Arguments

    collapse all

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

    Two-Body Keplerian

    The two-body-keplerian orbit propagator returns these fields.

    • SemiMajorAxis, in meters

    • Eccentricity

    • Inclination, in degrees

    • RightAscensionOfAscendingNode, in degrees

    • ArgumentOfPeriapsis, in degrees

    • TrueAnomaly, in degrees

    • Period, in seconds

    SGP4 and SDP4

    The sgp4 and sdp4 orbit propagators returns these fields.

    • MeanMotion, in degrees/second

    • Eccentricity

    • Inclination, in degrees

    • RightAscensionOfAscendingNode, in degrees

    • ArgumentOfPeriapsis, in degrees

    • MeanAnomaly, in degrees

    • Epoch

    • BStar, in 1/EarthRadius

    • Period, in seconds

    The orbital elements represent the mean values at Epoch.

    Ephemeris

    The ephemeris propagator returns these fields.

    • EphemerisStartTime

    • EphemerisStopTime

    • PositionTimeTable

    • VelocityTimeTable

    GPS

    The gps propagator returns these fields.

    • PRN

    • SatelliteHealth

    • GPSWeekNumber

    • GPSTimeOfApplicability, in seconds

    • SemiMajorAxis, in meters

    • Eccentricity

    • Inclination, in degrees

    • GeographicLongitudeOfOrbitalPlane, in degrees

    • RateOfRightAscension, in degrees/second

    • ArgumentOfPerigee, in degrees

    • MeanAnomaly, in degrees

    • Period, in seconds

    The orbital elements are derived from the SEM almanac file and defined in the Earth-Centered-Earth-Fixed (ECEF) frame.

    Version History

    Introduced in R2021a