Optimal Design Comparison of Hybrid Rocket for Small Satellite Launchers

In the present paper, a three stage hybrid rocket is considered as a small satellite launcher. The same engine is used in different numbers in each stage: 6,3 and 1 in the first, second and third stage, respectively. This design choice aims at an overall reduction of the launcher cost. Liquid oxygen and a paraffin-based fuel are employed as propellants. The feasibility of a ground launch from 45° North latitude is evaluated and compared to a similar three-stage launcher with airborne launch. In the present work, an electric turbo pump feed system is used. The optimization procedure exploits of a direct method for engine design parameters, whereas an indirect method optimizes the ascent trajectory once the engine design is given. Constant power operation is assumed for the electrical feed system. The initial mass of the launcher is given (5000 kg) and the payload mass is maximized for a given insertion orbit. The initial thrust is fixed in order to have an initial acceleration equal to 1.4 g. The results show that the proposed small satellite launcher concept is able to deliver payload mass in the range 50-100 kg into the desired orbit.