In this paper, design and optimization approaches are applied to a hybrid-powered small-satellite launcher, namely a deterministic and a robust-based method. A unique hybrid rocket engine is considered, employed in different numbers in each stage of the three-stage launcher: six, three, and one, respectively, in the first, second, and third stage. Liquid oxygen and paraffin-based fuel are considered as propellant combination. A blowdown feed system is assumed to ensure system simplicity and reduce the overall launcher cost. An airborne launch is proposed at a given altitude and speed, whereas the velocity angle on the horizon at first stage ignition is set free. The optimization procedure of the engine design parameters employs a direct method and an evolutionary algorithm in the deterministic and robust-based approach, respectively. An indirect method is used to optimize the ascent trajectory, given the engine design, in both the approaches. In the robust-based approach, uncertainties in the fuel regression rate are taken into account. Small-launcher initial mass is given and payload mass is maximized for a given insertion orbit. A mission-specific objective function is used in the robust-based optimization process to grant the actual achievement of mission goals, despite the uncertainties in the design. Two different design strategies are compared. In the first case the acceleration at first-stage ignition is fixed, whereas in the second case the initial thrust is optimized. Results show that a 5-ton hybrid-rocket launcher can be viable for launch of satellites in the 50–100 kg range and that a small payload reduction, with respect to deterministic optimized solutions, is sufficient to ensure mission accomplishment and also grant the required design robustness.