This paper proposes the use of a three-stage hybrid rocket as a small-satellite launcher. Each stage is made with the same engine in different numbers (6, 3, and 1 in the first, second, and third stage) as a cost-reduction strategy. Liquid oxygen and a paraffin-based fuel are the propellants. A pressurized blowdown feed system is assumed to keep low costs and system simplicity. Airborne launch is considered at a given altitude and speed. The velocity angle on the horizon at first stage ignition is set free. The design optimization strategy employs a direct method for the engine design parameters and an indirect method to optimize the ascent trajectory given the engine design. The launcher initial mass is considered to be given and the payload mass is maximized for a given insert orbit. Two different design strategies are compared. In the first ease the acceleration at first stage ignition is fixed, while in the second case the initial thrust is optimized. The latter case, whose payload results to be larger, requires a constraint on the maximum acceleration (6 g). Results show that a 5-ton hybrid- rocket launcher can be viable for launch of satellites in the 50-100 kg range.
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