Thesis (M.S.)--University of Missouri--Rolla, 2007. / Vita. The entire thesis text is included in file. Title from title screen of thesis/dissertation PDF file (viewed October 24, 2007) Includes bibliographical references (p. 95-98).

Thesis (Ph. D.)--Duke University, 2008. / Includes bibliographical references.

A premixed gaseous rocket injector was designed and successfully operated over a limited range of fuel-rich operating conditions for the purpose of soot modeling for ethylene and oxygen mixtures. The injector had the advantage of delivering a homogenous mixture to the combustion chamber, lower soot production, and higher performance potential by removing the fuel atomization process which affects the combustion process and is inherent for non-premixed injectors. The premixed injector was operated at oxygen-fuel ratios from 1.0 to 1.8 with a mass flow of 0.024 kg/sec achieving a chamber pressure of 76 psi without propensity of flashback for 0.032[gamma] injector orifices. Increased mass flow rates of 0.027 kg/sec were achieved by increasing the injector orifice diameters to 0.0625[gamma] which produced a chamber pressure of 127 psi and a characteristic exhaust velocity efficiency of 90.1 %. Flashback was eventually observed at an oxygen-to-fuel ratio of 1.2 where the pressure drop was across the injector was less than 388.6 kPa and the bulk mixture velocity through the injector orifices was approximately 90 m/s. Maintaining bulk velocity sufficiently above this value should prevent flashback from occurring, but will likely need to be characterized for additional orifice diameters and pressure differentials. / Funded by: SEinc307.

Thesis (M.S. in Engineering Science (Mechanical Engineering))--Naval Postgraduate School, December 2003. / Thesis advisor(s): Ashok Gopinath, Christopher Brophy. Includes bibliographical references (p. 71-72). Also available online.

No description available.

Thesis (M.S.)--Virginia Polytechnic Institute and State University, 1993. / Vita. Abstract. Includes bibliographical references (leaves 121-122). Also available via the Internet.

Thesis (M.S. in Astronautical Engineering, Aeronautical and Astronautical Engineer) Naval Postgraduate School, Dec. 2001. / Thesis advisors: Christopher Brophy, Jose Sinibaldi, Ashok Gopinath. "December 2001." Includes bibliographical references (p. 73). Also available in print.

The understanding and prediction of transient phenomena inside Liquid Rocket Engines (LREs) have been very difficult because of the many challenges posed by the conditions inside the combustion chamber. This is especially true for injectors involving liquid oxygen LOX and gaseous hydrogen GH₂. A wide range of length scales needs to be captured from high-pressure flame thicknesses of a few microns to the length of the chamber of the order of a meter. A wide range of time scales needs to be captured, again from the very small timescales involved in hydrogen chemistry to low-frequency longitudinal acoustics in the chamber. A wide range of densities needs to be captured, from the cryogenic liquid oxygen to the very hot and light combustion products. A wide range of flow speeds needs to be captured, from the incompressible liquid oxygen jet to the supersonic nozzle. Whether one desires to study these issues numerically or experimentally, they combine to make simulations and measurements very difficult whereas reliable and accurate data are required to understand the complex physics at stake. This thesis focuses on the numerical simulations of flows relevant to LRE applications using Large Eddy Simulations (LES). It identifies the required features to tackle such complex flows, implements and develops state-of-the-art solutions and apply them to a variety of increasingly difficult problems. More precisely, a multi-species real gas framework is developed inside a conservative, compressible solver that uses a state-of-the-art hybrid scheme to capture at the same time the large density gradients and the turbulent structures that can be found in a high-pressure liquid rocket engine. Particular care is applied to the implementation of the real gas framework with detailed derivations of thermodynamic properties, a modular implementation of select equations of state in the solver. and a new efficient iterative method. Several verification cases are performed to evaluate this implementation and the conservative properties of the solver. It is then validated against laboratory-scaled flows relevant to rocket engines, from a gas-gas reacting injector to a liquid-gas injector under non-reacting and reacting conditions. All the injectors considered contain a single shear coaxial element and the reacting cases only deal with H₂-O₂ systems. A gaseous oyxgen-gaseous hydrogen (GOX-GH₂) shear coaxial injector, typical of a staged combustion engine, is first investigated. Available experimental data is limited to the wall heat flux but extensive comparisons are conducted between three-dimensional and axisymmetric solutions generated by this solver as well as by other state-of-the-art solvers through a NASA validation campaign. It is found that the unsteady and three-dimensional character of LES is critical in capturing physical flow features, even on a relatively coarse grid and using a 7-step mechanism instead of a 21-step mechanism. The predictions of the wall heat flux, the only available data, are not very good and highlight the importance of grid resolution and near-wall models for LES. To perform more quantitative comparisons, a new experimental setup is investigated under both non-reacting and reacting conditions. The main difference with the previous setup, and in fact with most of the other laboratory rigs from the literature, is the presence of a strong co-flow to mimic the surrounding flow of other injecting elements. For the non-reacting case, agreement with the experimental high-speed visualization is very good, both qualitatively and quantitatively but for the reacting case, only poor agreement is obtained, with the numerical flame significantly shorter than the observed one. In both cases, the role of the co-flow and inlet conditions are investigated and highlighted. A validated LES solver should be able to go beyond some experimental constraints and help define the next direction of investigation. For the non-reacting case, a new scaling law is suggested after a review of the existing literature and a new numerical experiment agrees with the prediction of this scaling law. A slightly modified version of this non-reacting setup is also used to investigate and validate the Linear-Eddy Model (LEM), an advanced sub-grid closure model, in real gas flows for the first time. Finally, the structure of the trans-critical flame observed in the reacting case hints at the need for such more advanced turbulent combustion model for this class of flow.

NASAâ s VAriable Specific Impulse Magnetoplasma Rocket Engine (VASIMR) will provide a highly efficient propulsion source that can dramatically reduce Martian transit times, provide for more abort contingencies, and protect astronauts from space radiation with its highly radiation-absorbent hydrogen fuel. The VASIMR is still in its developmental infancy and requires many years of research before its initial operational capability. Much is still unknown about the complex plasma interactions in the exhaust. A Langmuir probe was designed, constructed, and operated to determine current density radial profiles and total particle flux at various stages in the exhaust of VASIMR. The Langmuir probe results proved that the exhaustâ s radial profile is Gaussian, experimentally validated predictions of magnetic field line dragging, and verified the ionization efficiency of VASIMR.

Thesis (Aeronautical and Astronautical Engineer)--Naval Postgraduate School, June 2003. / Thesis advisor(s): I. Michael Ross, Steve Matousek. Includes bibliographical references (p. 149-150). Also available online.