SGP4 Add Elements
The Add Elements window enables you to import or manually define your own GP element data.
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 GP element when using two or more TLEs.
- Epoch - Switch between GP elements based on the epoch of the TLE.
- Mid-Point - Switch between GP elements based on the midpoint between two TLE epochs.
- TCA - Switch between GP element based on the time of closest approach. STK calculates TCA 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 GP elements but may not represent the best estimated state of the vehicle.
- Override - Switch between GP elements at the exact time specified.
Source
There are three types of sources that you can use to define the GP element set.
Import from AGI Server
Select this to 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 Stored access intervals can become outdated when the latest satellite ephemerides are applied to the scenario.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
Select this to import the elements for the selected SSC Number from a GP data file. Click 
to browse for a TCE/TLE, TXT, GZ, CSV, XML, or OMM file.
Enter Elements
Manually enter the orbital elements.
| Element | Description |
|---|---|
| Orbit Epoch | Enter 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 | Specify 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 eccentricity. |
| Inclination | Specify the angle between the angular momentum vector, perpendicular to the plane of the orbit, and the inertial Z axis. |
| Argument of Perigee | Specify 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 | Specify 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 | Specify 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 | Specify the first time derivative of mean motion. |
| Motion Dot Dot | Specify the second time derivative of mean motion. |
| Classification | Enter a one-letter classification indicator.
|
| Bstar | Specify the drag term for the satellite. |
| Rev. Number | Enter the pass number of the object at the orbit epoch shown. |
| Int'l Designator | Enter the international designation of the satellite |