Satgen318 In Orbit Pt2 - Keplerian Elements by GM4IHJ 29 April 95 In order to track a satellite from our ground station , we need to be able to describe its orbit. Johannes Kepler showed how this can be done , if we measure certain parameters of the orbit at a given EPOCH time ,and can then calculate how the satellites position will have changed at a later time. We call the necessary set of orbital parameters the "KEPLERIAN ELEMENTS". These elements describe the orbital ellipse , where the centre of the earth is one focus of the ellipse, and the MAJOR AXIS of the ellipse runs >From the orbit perigee low point nearest the earth , up through the earth centre to the apogee orbit highest point above the earth. This major axis is crossed at its centre by the minor axis of the ellipse, and the ratio between the lengths of these two axis is known as the ECCENTRICITY of the ellipse , where eccentricity e = square root of (1- (major/minor) (^2)) and distance centre of earth to apogee = 1/2( major axis) x( 1+ e) and distance centre of earth to perigee = 1/2( major axis) x( 1- e) Please note that in practice most keplerian elements quote NUMBER OF ORBITS PER DAY rather than Major axis . Noting that this is much more informative and is derived directly from the length of the major axis. To position the satellite orbit plane with respect to the earths equatorial plane Kepler included INCLINATION, the angle between the orbit plane and the equator. While the equally important alignment of the orbit plane with the stars, is given as the angle between the stellar reference point in the sky known as the first point of aries, and the south to north ascending point of the orbit as it overflies the equator. This angle is called the RIGHT ASCENSION OF THE ASCENDING NODE. WE also use this equator crossing point to locate the orbit perigee low point by describing the angle around the orbit from the equator crossing to the perigee point as the ARGUMENT OF PERIGEE. Last but not least we need to say how far the satellite had gone around the orbit from perigee at the epoch time the parameters were measured. We call this angle the orbit MEAN ANOMALY. So a set of Keplerian Elements for satellite Dove Oscar 17 is Epoch Time year day of year.decimal day 95086.25250159 Inclination of orbit with respect to equator 98.5834 Right Ascension of Ascending Node 173.9260 Eccentricity 0.0012275 Arguhent of Perigee 78.6O08 Mean Anomaly 281.6448 Number of orbits per day (Mean Motion) 14.30074631 Please note that NASA elements also include other figures , but these are for satellite identity or for other forms of tracking calculation . Users should also be aware that this form of orbital elements applies only to artificial satellites of the earth . Different data formats are used when tracking bodies which do not rotate around the earth Eg Sun ,Moon, etc