F. S. Mechentel, A. M. Coates, K. Haefner, P. Challa, Prof. B. Cantwell Hybrid rocket motors represent a game-changing technology in the space industry. The hybrid concept (Figure 1) is a type of chemical propulsion utilizing a fuel and an oxidizer in two different phases, typically a liquid or gaseous oxidizer and a solid fuel. Hybrid motors have multiple advantages over their solid or liquid propellant counterparts. They are mechanically simple and can be throttled, shut down, and restarted which makes them safer and more cost-efficient than conventional systems. However, small- scale hybrids can be subject to low combustion efficiency limiting their potential of being a leading candidate for propulsion solutions. Introduction The goal of this study is to advance hybrid rocket technology for in-space applications requiring small-scale thrusters. The challenges for the design of such compact chemical propulsion systems include volume, mass, and safety constraints as well as efficiency requirements. This involves determining appropriate propellant combinations and the usefulness and efficiency of post combustion chambers to improve propellant mixing. Four tasks have been proposed to further advance the research in small-scale hybrid rocket propulsion. • Task 1: A zero-dimensional combustion model has been developed in order to quantify the effects of incomplete reaction and heat losses on c* efficiency. This model suggests that insufficient mixing is a crucial limitation and passive mixing devices have the potential to significantly increase the overall propulsive efficiency. (Completed) • Task 2: Computational Fluid Dynamics (CFD) will be used as a design tool to improve combustion chamber designs. These simulations will give unique insights into efficient chamber and post combustion chamber mixing areas and will reduce the number and cost of experimental studies that would be required to achieve these results. (Ongoing) • Task 3: An optimization design code will be developed in order to identify the most important parameters affecting the overall design from a system’s point of view. (Spring 2016) • Task 4: An experimental setup using gaseous oxygen will be designed and built in order to validate the computational results and obtain valuable regression rate data that will be used in preliminary designs and the optimization design code. (Ongoing) Objectives Small-Scale Hybrid Rocket Combustion Chamber Design for Improved Efficiency Figure 1: Schematic of a typical hybrid rocket motor The characteristic velocity (c*) is an interesting parameter to work with since it translates the ability to extract thermodynamic energy from a chemical system. It essentially determines the efficiency of the combustion process independently of nozzle characteristics. This model uses steady conservation equations to determine the equilibrium conditions given a certain reaction efficiency or amount of heat losses. Results are shown in Figure 3. Task 1: Zero-Dimensional Combustion Model Computational models provide useful guidance for combustion chamber design. ANSYS Fluent has been chosen as the CFD solver for its ability to include chemistry models. Qualitative results can be drawn from the study of flow patterns resulting from specific chamber geometries. Research in the area ([1]) has been used as a baseline for this task. Task 2: CFD Simulations A new hybrid rocket motor (3” diameter) is currently being designed and will be built and tested in the Summer 2016. This setup has the capability of running gaseous oxygen and various fuels such as PMMA (Poly(methyl methacrylate), HDPE (High Density Polyethylene), HTPB (Hydroxyl- terminated polybutadiene), and Paraffin. The chamber and post combustion chamber lengths can be varied, and the motor will use a gaseous O 2 /CH 4 torch as a reliable ignition source. Clear PMMA will be used to have visual access to the fuel port and possibly optically measure the fuel regression rate. Ultrasound sensing is a non intrusive regression rate measurement technique ([2]) that is currently being studied. A small-scale portable setup has been used to determine the feasibility of these measurement techniques for possible implementation on the actual test stand. Task 4: Experimental Efforts This project uses a variety of tools including combustion modeling, system optimization, computational fluid dynamics and experimental testing to • Identify the main parameters limiting small-scale hybrid rocket performance • Suggest design guidelines to overcome these losses • Provide qualitative (modeling) and quantitative (testing) data to improve hybrid motor designs Conclusion For further information please contact Flora S. Mechentel ([email protected]) Stanford Propulsion and Space Exploration Lab (SpaSe) Further Information This project is currently funded under a NASA-CalTech Jet Propulsion Laboratory (JPL) 2015-2016 Strategic University Research Partnership (SURP). The authors would like to thank Dr. R. Mitchell, G. Zilliac, B. Nakazono, D. Vaughan, Dr. A. Karp and Dr. B. Evans for their insight and valuable comments. Acknowledgments [1]: M. Lazzarin et. al. “Computational Fluid Dynamics Simulation of Hybrid Rockets of Different Scales”, Journal of Propulsion and Power, 2015 [2]: F. Cauty, “The Ultrasound Waves : a Measurement Tool for Energetic Material Characterization”, 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 2004-4057 Bibliography ∗ = ሶ η ∗ = Figure 2: 0-D combustion model setup. Left: c* and c* efficiency equations. Right: Conservation of energy in the modeled chamber Figure 3: c* efficiency variation for the incomplete reaction and heat loss problems for O 2 and various fuels Figure 6: Solidworks drawings of the small-scale hybrid rocket motor design (02/2016) Figure 5: Left: O 2 /PMMA portable setup used for proof of concepts (optical regression rate measurements and ultrasound sensing). Right: ultrasound sensor on a clear PMMA fuel grain April 2016 Figure 4: Current ANSYS Fluent axisymmetric simulations. Left: choked turbulent pipe flow (k-ε model). Right: CH 4 /O 2 reaction in a pipe