SGP4 Add Elements

The Add Elements window allows you to import or manually define your own TLE.

The Start Time and Stop Time fields display the Analysis time period over which the satellite will be propagated.

Switching Method

Defines when to switch between TLE sets, when multiple TLEs are used.

Epoch - Switch between TLE sets based on the epoch of the TLE.
Mid-Point - Switch between TLE sets based on the mid-point between two TLE epochs.
TCA - Switch between TLE sets based on the time of closest approach. Calculated by propagating each TLE of a pair over the time period between their respective epochs, and then determining the point at which the two paths are nearest each other. This option provides the smoothest possible fit between two TLE sets but may not represent the best estimated state of the vehicle.
Override - Switch between TLE sets at the exact time specified.

Source

There are three types of sources that you can use to define the TLE.

Import from AGI Server

Import the elements for the selected SSC Number from the AGI Server.

If you are updating satellites from an online server, you should disable the storing of access intervals. Stored access intervals can become outdated when the latest satellite ephemerides are applied to the scenario.

Import from File

Import the elements for the selected SSC Number from a TLE file. Use the button to browse for a *.tce or *.tle file.

Enter Elements

Manually enter the orbital elements.

Element	Description
Orbit Epoch	The UTC epoch for the elements. The format is YYDDD.DDDDDDDD. The grayed-out field that appears beneath the Orbit Epoch displays the date in the current scenario date units, for your reference.
Mean Motion	A measure of the osculating period of the orbit, expressed as an angular rate.
Eccentricity	Describes the shape of the ellipse. A value of 0 represents a circular orbit; values greater than 0 but less than 1 represent ellipses of varying degrees of ellipticality.
Inclination	The angle between the angular momentum vector (perpendicular to the plane of the orbit) and the inertial Z-axis.
Argument of Perigee	The angle from the ascending node to the eccentricity vector (lowest point of orbit) measured in the direction of the satellite's motion. The eccentricity vector points from the center of the Earth to perigee with a magnitude equal to the eccentricity of the orbit.
RAAN	The angle from the inertial X-axis to the ascending node. The ascending node is the point where the satellite passes through the inertial equator moving from south to north. Right ascension is measured as a right-handed rotation about the inertial Z-axis.
Mean Anomaly	The angle from the eccentricity vector to a position vector where the satellite would be if it were always moving at its average angular rate.
Mean Motion Dot	The first time derivative of mean motion.
Motion Dot Dot	The second time derivative of mean motion.
Classification	A one-letter classification indicator. U - Unclassified C - Classified S - Secret
Bstar	The drag term for the satellite.
Rev. Number	The pass number of the object at the orbit epoch shown.
Int'l Designator	The international designation of the satellite