Trajectory Predictions for High Eccentricity Orbits fo Space Debris Objects

Since 1957 the number of space debris has been increasing and it cause threat of collision. To calculate precisely space debris orbit we used several perturbations in our force model: geopotential, luni-solar effects, solar radiation pressure and influence of Earth's atmosphere. For satellites with altitude of perigee higher than 1000 km perturbations from the atmosphere is negligible. However for objects which reaches its lower parts is one of the most important perturbation. For the last perturbation we used NRMLMSISE-00 empirical model to calculate precise parameters for the atmosphere. For large amount of objects using numerical integration there appears to be a problem with time of calculations. For this reason, we decided to use analytical model, which is much faster and more convenient. Due to highly elliptical orbit we had to exchange the eccentricity function by the Hansen coefficients.