orbitalElements

R2026b

Orbital elements of satellites in scenario

Syntax

elements = orbitalElements(sat)

Description

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

Examples

Retrieve Orbital Elements of Satellite

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

sat — Satellite

scalar

Satellite object, specified as a scalar.

ReferenceCelestialBody — Reference celestial body for orbital elements

CentralBody property of satellite (default) | "Earth" | "Moon"

Since R2026b

Reference celestial body for orbital elements, specified as "Earth" or "Moon". The default value is the CentralBody property of the satellite. When ReferenceCelestialBody equals "Moon", the elements are computed using the gravitational parameter of the Moon.

Dependencies

When ReferenceCelestialBody equals "Moon", the OrbitPropagator property of the satellite must equal "two-body-keplerian", "ephemeris", or "numerical". Otherwise, the function returns an error.

In addition, the Moon must have been added to the satellite scenario as a central body. Otherwise, the function returns an error.

Data Types: string

Output Arguments

elements — Orbital elements

structure

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.

Two-Body Keplerian — Two-Body Keplerian orbit propagator fields

structure

The orbital elements are defined in the International Celestial Reference Frame (ICRF).

Field	Description
SemiMajorAxis	Semimajor axis, in meters
Eccentricity	Eccentricity
Inclination	Inclination angle, in degrees. The inclination is defined with respect to ICRF as the angle between the orbital plane and the ICRF xy plane.
RightAscensionOfAscendingNode	Right ascension of Ascending node, in degrees
ArgumentOfPeriapsis	Argument of periapsis, in degrees
TrueAnomaly	True anomaly at reference time, in degrees
Period	Orbital period, in seconds. Returns Inf for parabolic and hyperbolic trajectories (eccentricity >= 1).

Data Types: struct

SGP4 and SDP4 — SGP4 and SDP4 orbit propagator fields

structure

The orbital elements represent general perturbation mean elements.

Field	Description
MeanMotion	General perturbation mean motion, in degrees per second
Eccentricity	Mean eccentricity
Inclination	Mean inclination angle, in degrees
RightAscensionOfAscendingNode	Mean right ascension of Ascending node, in degrees
ArgumentOfPeriapsis	Mean argument of periapsis, in degrees
MeanAnomaly	Mean mean anomaly at the reference time, in degrees
Epoch	Epoch
BStar	Drag term, in per EarthRadius
Period	Mean orbital period, in seconds

Data Types: struct

Ephemeris — Ephemeris orbit propagator fields

structure

Field	Description
EphemerisStartTime	Ephemeris start time
EphemerisStopTime	Ephemeris stop time
PositionTimeTable	Position timetable
VelocityTimeTable	Velocity timetable

Data Types: struct

GPS — GPS orbit propagator fields

structure

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.

Field	Description
PRN	Pseudorandom noise number
SatelliteHealth	Satellite health
GPSWeekNumber	GPS week number, with time of ephemeris.
GPSTimeOfApplicability	Time of Applicability, in seconds
SemiMajorAxis	Semimajor axis, in meters
Eccentricity	Eccentricity
Inclination	Inclination angle at reference time, in degrees
GeographicLongitudeOfOrbitalPlane	Longitude of ascending node of orbit plane at weekly epoch, in degrees
RateOfRightAscension	Reference rate of right ascension, in degrees per second
ArgumentOfPerigee	Argument of perigee, in degrees
MeanAnomaly	Mean anomaly at the reference time, in degrees
Period	Orbital period, in seconds

Data Types: struct

Galileo — Galileo orbit propagator fields

structure

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

Field	Description
SatelliteID	Satellite system number
GALWeekNumber	GAL week number
TimeofEphemeris	Time of ephemeris, in seconds
SatelliteHealth	Satellite health
SemiMajorAxis	Semimajor axis, in meters
Eccentricity	Eccentricity
Inclination	Inclination angle at reference time, in degrees
GeographicLongitudeOfOrbitalPlane	Longitude of ascending node of orbit plane at weekly epoch, in degrees
RateOfRightAscension	Reference rate of right ascension, in degrees per second
ArgumentOfPerigee	Argument of perigee, in degrees
MeanAnomaly	Mean anomaly at the reference time, in degrees
Period	Orbital period, in seconds

Data Types: struct

Data Types: struct

Version History

Introduced in R2021a

R2026b: ReferenceCelestialBody name‑value argument added

You can now compute orbital elements with respect to the Moon by specifying the ReferenceCelestialBody name‑value argument as "Moon". When ReferenceCelestialBody equals "Moon", elements are computed using the Moon's gravitational parameter. This requires the OrbitPropagator to be "two-body-keplerian" or "numerical".

Version History

Introduced in R2021a

R2026b: Parabolic and hyperbolic trajectory support

The orbitalElements function now returns orbital elements for satellites on parabolic (eccentricity = 1) and hyperbolic (eccentricity > 1) trajectories. The Period field returns Inf for these orbit types.