F. S. Mechentel, A. M. Coates, K. Haefner, P. Challa, Prof. B. Cantwell

Hybrid rocket motors represent a game-changing technology in the spaceindustry. The hybrid concept (Figure 1) is a type of chemical propulsionutilizing a fuel and an oxidizer in two different phases, typically a liquid orgaseous oxidizer and a solid fuel. Hybrid motors have multiple advantagesover their solid or liquid propellant counterparts. They are mechanicallysimple and can be throttled, shut down, and restarted which makes themsafer and more cost-efficient than conventional systems. However, small-scale hybrids can be subject to low combustion efficiency limiting theirpotential of being a leading candidate for propulsion solutions.

Introduction

The goal of this study is to advance hybrid rocket technology for in-spaceapplications requiring small-scale thrusters. The challenges for the design ofsuch compact chemical propulsion systems include volume, mass, and safetyconstraints as well as efficiency requirements. This involves determiningappropriate propellant combinations and the usefulness and efficiency ofpost combustion chambers to improve propellant mixing. Four tasks havebeen proposed to further advance the research in small-scale hybrid rocketpropulsion.

• Task 1: A zero-dimensional combustion model has been developed inorder to quantify the effects of incomplete reaction and heat losses on c*efficiency. This model suggests that insufficient mixing is a cruciallimitation and passive mixing devices have the potential to significantlyincrease the overall propulsive efficiency. (Completed)

• Task 2: Computational Fluid Dynamics (CFD) will be used as a design toolto improve combustion chamber designs. These simulations will giveunique insights into efficient chamber and post combustion chambermixing areas and will reduce the number and cost of experimental studiesthat would be required to achieve these results. (Ongoing)

• Task 3: An optimization design code will be developed in order to identifythe most important parameters affecting the overall design from asystem’s point of view. (Spring 2016)

• Task 4: An experimental setup using gaseous oxygen will be designed andbuilt in order to validate the computational results and obtain valuableregression rate data that will be used in preliminary designs and theoptimization 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 toextract thermodynamic energy from a chemical system. It essentially determines the efficiency of thecombustion process independently of nozzle characteristics. This model uses steady conservationequations to determine the equilibrium conditions given a certain reaction efficiency or amount of heatlosses. Results are shown in Figure 3.

Task 1: Zero-Dimensional Combustion Model

Computational models provide useful guidance for combustion chamber design. ANSYS Fluent has beenchosen as the CFD solver for its ability to include chemistry models. Qualitative results can be drawnfrom 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 willbe built and tested in the Summer 2016. This setup has the capability ofrunning gaseous oxygen and various fuels such as PMMA (Poly(methylmethacrylate), HDPE (High Density Polyethylene), HTPB (Hydroxyl-terminated polybutadiene), and Paraffin. The chamber and post combustionchamber lengths can be varied, and the motor will use a gaseous O2/CH4

torch as a reliable ignition source.

Clear PMMA will be used to have visual access to the fuel port and possiblyoptically measure the fuel regression rate. Ultrasound sensing is a nonintrusive regression rate measurement technique ([2]) that is currently beingstudied. A small-scale portable setup has been used to determine thefeasibility of these measurement techniques for possible implementation onthe actual test stand.

Task 4: Experimental Efforts

This project uses a variety of tools including combustion modeling, systemoptimization, 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 (floram@stanford.edu)Stanford Propulsion and Space Exploration Lab (SpaSe)

Further Information

This project is currently funded under a NASA-CalTech Jet Propulsion Laboratory (JPL) 2015-2016Strategic 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

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η𝑐∗ =𝑃𝑡𝑚𝑒𝑎𝑠𝑢𝑟𝑒𝑑

𝑃𝑡𝑖𝑑𝑒𝑎𝑙

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 O2 and various fuels

Figure 6: Solidworks drawings of the small-scalehybrid rocket motor design (02/2016)

Figure 5: Left: O2/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: CH4/O2 reaction in a pipe