Abstract: The lifetime of the low orbit earth satellite is mainly dominated by atmospheric drag. The satellite loses altitude in the dense atmosphere and eventually falls. However, the lifetime of space objects of HEO type, which goes in a large-eccentricity ellipse with very low perigee altitude and very high apogee altitude, is determined mainly by third-body （Sun and Moon） perturbation, as well as its initial state, which, closely related to RAAN Ω and Argument of Perigee ω, determines the long periodic evolution of the eccentricity. Here examples are taken for Geosynchronous Transfer Orbit debris and the orbit lifetime are estimated for double stars TC-land TC-2, proving that two GTO debris with the same initial elements but 3 hour＇s difference in entering expected orbits will have distinct destinies, the same conclusion is also valid for TC-1 and TC-2. All the mentioned distinction could be resulted in by third-body perturbation and a difference in initial elements （Ω and ω）. This paper would end up with a detailed computation, which accords with previous analysis.