Improving Thrust Performance in Hybrid Rocket Engines through Fuel Grain Geometry Modification

This project focuses on optimizing hybrid rocket engines by investigating the effects of fuel grain and boundary geometries on thrust efficiency. Hybrid rocket engines, which use solid fuel and liquid oxidizer, offer advantages in safety and ease of handling compared to purely solid or liquid systems. However, their thrust performance remains lower than that of solid rocket engines due to combustion inefficiencies and flow instabilities. Previous literature demonstrates that helical geometries in the solid grain fuel ports increase fuel regression rate over a classic cylindrical geometry. Additionally, the fuel regression of cylindrical ports in solid engines increases with the addition of a “finned” section near the head end of the fuel grains. No publications explore adding such fins to a helical port. Additionally, previous literature identifies that modifying boundary geometries at the combustion chamber’s inlet and outlet significantly impacts stability and thrust. Studies demonstrate that adjusting outflow boundary geometries, such as varying chamber aspect ratio and smoothing transitions between the combustion chamber and nozzle, reduces pressure fluctuations and mitigates instability. Hybrid rocket engines are especially susceptible to instabilities at inlet and outlet boundaries, where sudden flow direction changes lead to turbulent eddies that disrupt combustion uniformity. Poorly designed boundary interfaces cause recirculation zones, choked flow, and uneven pressure distributions, all of which reduce thrust. Our group improves the thrust performance of a hybrid rocket first by implementing a novel fuel grain geometry utilizing a finned, helical geometry; and secondly by refining the inlet and outlet boundary geometries of hybrid fuel grains. Preliminary testing indicates that helical fuel grains outperform cylindrical fuel grains, consistent with prior published works. Adding “fins” along the helical cross-section further increases the rocket performance beyond that of a helical fuel grain with a circular cross-section. To improve the boundary geometries, we investigate the effects of a smooth, funnel-shaped transition between the fuel port and engine casing. By implementing doubly-tangent boundary transitions, we hypothesize that we minimize flow separation, reduce recirculating turbulence, and promote a more stable combustion process. To validate this hypothesis, we design, manufacture, and hot-fire a hybrid engine system with varying boundary transition geometries. Our methodology includes both computational modeling and experimental validation. First, we use simplified computational models in ANSYS Fluent to predict fluid flow properties such as skin friction and turbulence, which correlate with thrust performance. Given the limitations of fully resolving combustion flowfields in a research setting, we use appropriate approximations to guide design iterations. Second, we manufacture test articles using a refined process. Initially, we used 3D-printed ABS plastic fuel grains, which proved challenging to integrate and prone to inconsistencies, but allowed for complex internal geometries through additive manufacturing. To investigate the boundary transitions, we switch to paraffin wax fuel grains, which offer higher combustion efficiency, safer handling, and simpler manufacturing. By casting paraffin wax around 3D-printed negative molds of the desired fuel port geometries, we produce more uniform and repeatable test samples. The modular design allows for rapid iteration and testing of different inlet and outlet geometries. The experimental campaign evaluates thrust, specific impulse, and combustion stability for different boundary geometries. We first establish a baseline using cylindrical fuel grains before testing three variations of the proposed funnel-shaped boundary transition. The transition region’s length is systematically varied to determine its impact on flow characteristics and thrust performance. The oxidizer, liquid nitrous oxide (NOS), is atomized using a brass conical spiral nozzle, ensuring uniform injection into the combustion chamber. To prevent backflow, a one-way check valve is integrated into the forward closure. Our experimental findings are compared against computational predictions to refine our understanding of flow behavior in hybrid rocket engines. By iterating through design modifications based on test results, we develop a set of design guidelines for optimizing hybrid rocket fuel grain and boundary geometries. Ultimately, this research bridges the gap between theoretical combustion stability principles and practical hybrid rocket engine performance improvements. The results of this study contribute to the broader effort of improving hybrid propulsion systems, providing valuable insights into how boundary condition modifications enhance thrust performance and combustion stability.