Development of a hybrid rocket propulsion system for small satellites with torch plume simulation and temperature measurements

This thesis integrates a variety of topics, from testing small-scale hybrid rocket engines to applying optical temperature measurement techniques to generating computational fluid dynamics (CFD) models. While these may seem like disparate topics, they combine in this thesis to advance crucial areas of hybrid rocket engine development and testing. Hybrid Rocket Testing This project started out by taking the initial steps toward creating a modular, small-scale hybrid rocket engine, called the Primary Propulsion Thruster (PPT), that could serve as a small satellite propulsion system in future years. The primary goal was generating an engine using a solid polymethyl methacrylate (PMMA) fuel grain and a gaseous nitrous oxide oxidizer with a c* combustion efficiency of at least 92% when operating at or near the optimal oxidizer-to-fuel (O/F) ratio of 3.8. Achieving the desired combustion efficiency in an engine with the short, wide design oftentimes required for small satellites necessitated adding a variety of mixers at the aft end of the combustion chamber during successive tests. All of the goals set out for this project were achieved by shortening the fuel grain to 8.53 cm, adding a 6.35 mm G-10 thermoset diaphragm at the end of the cylindrical fuel grain and filling the engine's converging section with a solid PMMA fuel grain. This layout surpassed original project goals by providing an average 92% efficiency, achieving a high volumetric loading fraction, completing a record four-minute-long hybrid rocket engine burn and restarting three times after the initial burn using the same fuel grain. Furthermore, all of these goals were achieved in an engine with a short, wide shape while retaining the cost effectiveness, safety and environmental friendliness associated with hybrid rocket engines. Sodium Line Reversal The success of increasing the PPT combustion efficiency by adding a diaphragm inspired a related project exploring the effect of diaphragm placement on the mixing and combustion efficiencies in the engine. At the outset, a non-intrusive technique was sought to measure both efficiencies outside the combustion chamber, where researchers have free access to the flow. A fitting technique was identified in Baumann's two-wavelength modified sodium line reversal (SLR) technique, an optical temperature-measurement technique that can be used to measure the flow temperature just downstream of the engine nozzle throat. A multiple-line-of-sight instrument was designed to simultaneously apply Bauman's method at multiple locations in the plume to provide the temperature measurements required to more clearly elucidate changes in the combustion and mixing efficiencies during each test. Before applying Baumann's two-wavelength modified SLR technique and the multiple-line-of-sight instrument to a hybrid rocket engine plume, the technique was applied in a lab environment using an oxyacetylene torch as a surrogate for the hybrid rocket engine plume. Subsequent construction and testing of a single-line-of-sight, high-precision optical setup revealed that positioning a saltwater atomizer far behind the oxyacetylene torch handle, applying a very low flow rate of salt water and keeping the atomizer nozzle diameter small provided consistent sodium addition to the flame. Subsequent testing showed that the technique could be used to successfully calculate the plume temperature. Computational Fluid Dynamics The oxyacetylene torch handle, nozzle and combusting plume were concurrently simulated in ANSYS Academic Research Fluent, 2019 R1 to gain insight into the torch plume temperature (for comparison to the results of the SLR experiment) as well as to create maps of a plethora of other plume properties. The experiment's flow properties were simulated using a volume mesh containing over six million hexcore and non-conformal cells. Turbulence was modeled using the SST k-ω model while combustion was modeled using a 15-species, 25-step reaction mechanism implemented through CHEMKIN. The simulation resulted in the first computational mapping of the flow field in an oxyacetylene torch handle, nozzle and combusting plume, giving insight into a flow frequently used in laboratory and scientific experiments. The property maps provided in this thesis are of practical rather than merely academic value to the research community. Researchers using oxyacetylene torch flows similar or identical to the one used in this experiment can apply the data maps provided in this thesis to produce educated guesses of the modeled properties needed to complete calculations in their own experiments. As a result, the CFD property maps provided in this thesis have the potential to facilitate higher fidelity calculations in a plethora of other projects. Overall Project Synthesis The three sections of this thesis merge to provide a guide for the development of a non-toxic, environmentally friendly, throttleable and restartable hybrid-rocket-based in-space propulsion system delivering performance equivalent to existing engines, increases in system safety and decreases in overall system costs. Blending • the sodium line reversal technique discussed in Chapter 5 with • the oxyacetylene torch simulation results provided in Chapter 6 and • the multi-channel instrument designed and discussed in Appendix A facilitates studies of the combustion and mixing efficiencies in any region containing high-temperature combusting gases, for example small-scale hybrid rocket engine plumes. Researchers can use these tools to further enhance the efficiency of the already successful and innovative design for the short and wide hybrid-rocket-based Primary Propulsion Thruster developed in Chapter 4, which has the potential to radically alter the type of propulsion engine used on small satellites. For example, using these engines on small satellites significantly decreases the likelihood of engine explosion during launch, thereby easing the acceptance of small satellites as secondary payloads on a much broader range of launches. This will decrease the system cost and more-readily fulfill the high demand for small satellite launches. Furthermore, reducing requirements for tanks and plumbing by half means that small satellites designs can be used on a much wider array of missions due to the increase in storage space and decrease in the propulsion system mass.

Keywords: Aeronautics and Astronautics.
Advisor: Cantwell Brian
Alonso Juan José
Senesky Debbie