Browne, Elizabeth C., authorMarchese, Anthony, advisorWindom, Bret, advisorWatson, Ted, committee member2020-09-072020-09-072020https://hdl.handle.net/10217/212046Rocket engines create extreme conditions for any material to withstand. The combustion temperatures in rocket engines are substantially greater than the melting points of metals, and wall temperatures must be maintained well below the melting point to ensure structural integrity. This requirement necessitates a robust cooling system for the combustion chamber and nozzle to endure the mandated burn times. A liquid rocket engine utilizing ethane/ethylene as the fuel and nitrous oxide as the oxidizer, which is currently under development by Pioneer Astronautics, required a detailed analysis of thrust chamber cooling options. Due to the impracticality of experimentally validating the performance of each design parameter, this thesis employed computational methods to investigate two common cooling systems for rocket engines - ablative and regenerative - to determine their effectiveness at 130 and 200 chamber pressures, as prescribed by Pioneer Astronautics. Additional 1000-psi chamber pressure models were investigated for prediction validation. An analytical model was developed and utilized to elucidate the behavior of both cooling methods, while regenerative cooling was additionally analyzed using numerical modeling, coupling finite element analysis (FEA) and computational fluid dynamics (CFD) software. Simulations were created of the fluid dynamics and heat transfer within the rocket engine and coolant channels for numerous regenerative designs. The designs examined included a single-channel model utilizing only the liquid ethylene/ethane fuel as the coolant, and a dual-channel model using both the fuel and the nitrous oxide as coolants in separate sets of channels. In the single-channel regenerative cooling design, both the analytical and numerical models exhibited insufficient cooling capacity with coolant temperatures of 3-11 K above the critical temperature of 292.5 K. However, the dual-channel model provided the supplemental thermal energy absorption necessary to maintain engine wall and coolant temperatures within the allowable limits. From a design and manufacturing standpoint, ablative cooling is far simpler to implement than regenerative cooling. Although, material erosion at the throat reduces engine performance over time. Integrating ablative cooling in the combustion chamber and nozzle bell with dual-channel regenerative cooling near the throat has the potential to provide the requisite heat removal to ensure sustained material strength while maintaining all reactants in a condensed liquid phase.born digitalmasters thesesengCopyright and other restrictions may apply. User is responsible for compliance with all applicable laws. For information about copyright law, please see https://libguides.colostate.edu/copyright.CFDregenerative coolingliquid rocket engineablative coolingText