This article has multiple issues. Unsourced material may be challenged and removed. Since, for station-keeping, satellites intended for this orbit typically carry highly efficient but low thrust engines, total mass delivered to GSO is generally maximized if the launch vehicle provides only the Delta-v required to fundamentals of astrodynamics and applications pdf download at high thrust—i.
Earth’s atmosphere and overcome gravitational losses—and the satellite provides the Delta-v required to turn the resulting intermediate orbit, which is the GTO, into the useful GSO. The period of a standard geosynchronous transfer orbit is about 10. This method however takes much longer to achieve due to the low thrust injected into the orbit. The typical launch vehicle injects the satellite to a supersynchronous orbit having the apogee above 42,164 km. The satellite’s low thrust engines are thrusted continuously around the geostationary transfer orbits in an inertial direction. This inertial direction is set to be in the velocity vector at apogee but with an outer plane direction.
The outer plane direction removes the initial inclination set by the initial transfer orbit while the inner plane direction raises simultaneously the perigee and lowers the apogee of the intermediate geostationary transfer orbit. In case of using the Hohmann transfer orbit, only a few days are required to reach the geosynchronous orbit. By using low thrust engines or electrical propulsion, months are required until the satellite reaches its final orbit. The inclination and eccentricity must both be reduced to zero to obtain a geostationary orbit.
V required for a plane change is proportional to the instantaneous velocity, the inclination and eccentricity are usually changed together in a single manoeuvre at apogee where velocity is lowest. V for the two maneuvers. V in a combined maneuver will always be less than in two maneuvers. Even at apogee, the fuel needed to reduce inclination to zero can be significant, giving equatorial launch sites a substantial advantage over those at higher latitudes. Although some launchers can take their payloads all the way to geostationary orbit, most end their missions by releasing their payloads into GTO. The spacecraft and its operator are then responsible for the manoeuvre into the final geostationary orbit. The five-hour coast to first apogee can be longer than the battery lifetime of the launcher or spacecraft, and the manoeuvre is sometimes performed at a later apogee or split among multiple apogees.
The solar power available on the spacecraft supports the mission after launcher separation. Also, many launchers now carry several satellites in each launch to reduce overall costs, and this practice simplifies the mission when the payloads may be destined for different orbital positions. Because of this practice, launcher capacity is usually quoted as spacecraft mass to GTO, and this number will be higher than the payload that could be delivered directly into GEO. 185 km x 35,786 km at 27. If the manoeuvre from GTO to GEO is to be performed with a single impulse, as with a single solid rocket motor, apogee must occur at an equatorial crossing and at synchronous orbit altitude. This implies an argument of perigee of either 0 or 180 degrees. If the GTO inclination is zero, as with Sea Launch, then this does not apply.
It also would not apply to an impractical GTO inclined at 63. The preceding discussion has primarily focused on the case where the transfer between LEO and GEO is done with a single intermediate transfer orbit. More complicated trajectories are sometimes used. Proton even offers to perform a supersynchronous apogee maneuver up to fifteen hours after launch. Space Mission Design and Analysis, 2nd Edition. Published jointly by Microcosm, Inc.
Issue 5 Revision 1, 2011 July, pp. 2013 July 27 Appendix F. This page was last edited on 8 November 2017, at 21:32. The diagram shows a Hohmann transfer orbit to bring a spacecraft from a lower circular orbit into a higher one.