
Orbital elements are used to describe the motion of satellites within an orbit. 

Get the latest International Space Station orbital elements as it grows over the next several years. 
Orbital Elements
Information regarding the orbit trajectory of the International Space Station is provided here courtesy of the Johnson Space Center's Flight Design and Dynamics Division  the same people who establish and track U.S. spacecraft trajectories from Mission Control. This data can be used in a variety of programs that require accurate knowledge of the orbit of the satellite. Such applications might include ground track plotting programs, visual sighting programs and programs that predict past or future spacecraft trajectory information.
The trajectory data, which describes the motion of a satellite, can take on a variety of formats. We provide this data as either Keplerian osculating or mean element sets, which include the TWOLINE and AMSAT formats. In general, it takes at least six parameters to uniquely define an orbit and a satellite's position within the orbit. The mean element set format also contains the mean orbital elements, plus additional information such as the element set number, orbit number and drag characteristics. Both of these formats are common and most applications requiring trajectory data will be able to accept data of this kind.
The six orbital elements used to completely describe the motion of a satellite within an orbit are summarized below:
Orbital Elements: 
Semimajor axis 
a 
Defines the size of the orbit. 
Eccentricity 
e 
Defines the shape of the orbit. 
Inclination 
i 
Defines the orientation of the orbit with respect to the Earth's equator.< 
Argument of Perigee 

Defines where the low point, perigee, of the orbit is with respect to the Earth's surface. 
Right Ascension of the Ascending Node 

Defines the location of the ascending and descending orbit locations with respect to the Earth's equatorial plane. 
True/Mean Anomaly 

Defines where the satellite is within the orbit with respect to perigee. 
If maneuvers that change the size of the orbit are planned, a detailed table of maneuver parameters is included plus new orbital elements and mean element sets after each maneuver. These maneuver parameters include the LVLH maneuver delta V components at the impulsive ignition time, the M50 delta V components, the maneuver magnitude and the resulting mean apogee and perigee altitudes. In general, new data is provided after each maneuver or 48 hours, whichever occurs first.
Additionally, a satellite's motion can be described using a 3D inertial coordinate system or frame of reference. This data set is called a Cartesian state vector and provides the position and velocity of the satellite with respect to a threeaxis Cartesian coordinate system. For informational purposes, the following technical definitions are provided to define the frames of reference of the displayed Cartesian state vectors.
Definitions : 
M50

= 
Inertial mean of year 1950 frame of reference. 
Epoch 
= 
beginning of Besselian year 1950. 
X axis 
= 
mean vernal equinox of epoch. 
Z axis 
= 
earth mean rotation axis of epoch. 
Y axis 
= 
completes right handed Cartesian Earthcentered system. 

J2K 
= 
Inertial mean of year 2000 frame of reference. 
Epoch 
= 
Julian date 2451545.0 tdT. 
X axis 
= 
mean vernal equinox of epoch. 
Z axis 
= 
earth mean rotation axis of epoch. 
Y axis 
= 
completes right handed Cartesian Earthcentered system. 

LVLH 
= 
Local Vertical Local Horizontal rotating, instantaneously inertial, frame of reference. 
Epoch 
= 
Vector epoch time. 
X axis 
= 
completes the right handed orthogonal system. 
Z axis 
= 
lies along the radius vector, positive toward the center of the earth. 
Y axis 
= 
lies along the instantaneous angular momentum vector, negative in the direction of the angular momentum. 
Please direct any comments to:
DM33/ Orbit Analysis and Navigation at email: william.h.tracy@nasa.gov 