Numerical geometry optimization of a hybrid rocket engine using finite volume method

Abstract

Hybrid rockets are recognized for their safety, affordability, ease of thrust control, and low-risk fuel transportation. However, their low regression rate results in a reduced specific impulse, limiting their performance. This drawback can be partially mitigated by using paraffin as the fuel, owing to its high regression rate and favorable combustion properties. The primary objective of rocket engine design is to achieve maximum thrust under the prevailing operational conditions. Specific impulse, in general, is influenced by various factors such as the oxidizer/fuel (O/F) ratio, oxidizer injector diameter, mass flow rates of oxidizer and fuel, lengths of the pre-combustion and post-combustion chambers, fuel dimensions, nozzle geometry (including the divergent and convergent sections), ambient and combustion chamber pressures, as well as the chemical composition of the propellants. In this study, an optimum hybrid rocket engine design was developed using a liquid oxygen-paraffin propellant pair. The preliminary engine design was conducted following methods described in the literature, and parameter optimization was performed using the Ansys Fluent software, which employs the finite volume numerical method. Key performance metrics, including thrust, specific impulse, Mach numbers at the nozzle throat and outlet, as well as temperature, pressure, and density distributions within the engine, were analyzed. The results indicated that the optimum engine parameters deviate from those predicted by traditional design equations. While the targeted thrust value based on literature-derived equations was 230 N, the numerically optimized design achieved a thrust of 258.6 N, representing an 11% improvement.