Development and testing of a small hybrid rocket motor for space applications

In recent years there has been a significant renewed interest in hybrid propulsion for its unique features, in a world becoming more careful about safety, costs and environmental impact. Since the solid grain is made only of inert fuel and the propellants are physically separated by distance and phase, fabrication, transportation, storage, and handling can be done in complete safety and there is no possibility of an explosion. Moreover, the operational reliability is higher compared to solid and liquid rockets. Hybrid rocket motors can also be throttled controlling only the flow of the liquid oxidizer and can be stopped and restarted multiple times with an appropriate ignition system. There is also the possibility to find several green combinations of oxidizer and fuel within the wide propellant choice. All the inherent advantages of hybrid propulsion can lead to a significant reduction in the total operational costs. Unfortunately, this technology presents also several disadvantages, due to the characteristic diffusive flame mechanism, like low regression rate, low combustion efficiency, low volume loading, and mixture ratio shift. However, most of these drawbacks can be solved through a correct design process and choosing convenient approaches. The final ambition of this research is to help with the understanding of hybrid propulsion and its design. To achieve this goal, a set of simple explicit equations to describe hybrid rocket behavior and motor sizing have been derived, in order to have a better sensibility on the trends, possibilities, and limits of this promising but tricky kind of propulsion system. The equations refer to the length, diameter, volume loading and length to diameter ratio of the combustion chamber, the performance penalties and thrust variations incurred with time and throttling and the mission envelope of hybrid rocket motors. Two alternative configurations of hybrid rocket motors have been designed, numerically validated and tested, in order to demonstrate that the technology has reached a high maturity level and that it can be implemented in actual space missions. The first approach is based on swirl injection of HTP burning with a HDPE solid grain, while the second one uses an axial injector and a propellant combination of HTP and a paraffin-based fuel. In both cases, the HTP is first forced to pass through a catalytic reactor, where it decomposes into gaseous oxygen and water vapor at a temperature of around 1000 K. The issues that arise designing the flight weight version of the same hybrid rocket motor have been also discussed and some techniques have been proposed in order to correctly size thermal protections that can efficiently withstand the thermal loads.