A. James Clark School of Engineering

Permanent URI for this communityhttp://hdl.handle.net/1903/1654

The collections in this community comprise faculty research works, as well as graduate theses and dissertations.

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Now showing 1 - 7 of 7
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    Massively Parallel Large Eddy Simulation of Rotating Turbomachinery for Variable Speed Gas Turbine Engine Operation †
    (MDPI, 2020-02-06) Jain, Nishan; Bravo, Luis; Kim, Dokyun; Murugan, Muthuvel; Ghoshal, Anindya; Ham, Frank; Flatau, Alison
    Gas turbine engines are required to operate at both design and off-design conditions that can lead to strongly unsteady flow-fields and aerodynamic losses severely impacting performance. Addressing this problem requires effective use of computational fluid dynamics tools and emerging models that resolve the large scale fields in detail while accurately modeling the under-resolved scale dynamics. The objective of the current study is to conduct massively parallel large eddy simulations (LES) of rotating turbomachinery that handle the near-wall dynamics using accurate wall models at relevant operating conditions. The finite volume compressible CharLES solver was employed to conduct the simulations over moving grids generated through Voronoi-based unstructured cells. A grid sensitivity analysis was carried out first to establish reliable parameters and assess the quality of the results. LES simulations were then conducted to understand the impact of blade tip clearance and operating conditions on the stage performance. Variations in tip clearance of 3% and 16% chord were considered in the analysis. Other design points included operation at 100% rotor speed and off-design conditions at 75% and 50% of the rotor speed. The simulation results showed that the adiabatic efficiency improves dramatically with reduction in tip gap due to the decrease in tip leakage flow and the resulting flow structures. The analysis also showed that the internal flow becomes highly unsteady, undergoing massive separation, as the rotor speed deviates from the design point. This study demonstrates the capability of the framework to simulate highly turbulent unsteady flows in a rotating turbomachinery environment. The results provide much needed insight and massive data to investigate novel design concepts for the US Army Future Vertical Lift program.
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    NEUTRON SHIELDING DESIGN FOR CENTRIFUGALLY CONFINED SPACE PROPULSION SYSTEM
    (2023) Parsons, Jennifer; Sedwick, Raymond J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
    This thesis presents a preliminary neutron shielding design for the HTS coils of a centrifugally confined fusion space propulsion system, which is a promising technology for future space travel. The design process involved a comprehensive study of neutron transport, material selection, and shielding optimization using MCNP and MATLAB simulations. First, the neutron attenuating properties of reflector, moderator, and absorber candidate materials were compared in MCNP. The thickness and composition of the shield were optimized from the resulting MCNP data. Next, two overall reactor and shielding geometry models were developed in MATLAB to estimate the total mass of the HTS shielding for both coils. The first model assumed a point neutron source and uniform thickness across the surface area of the shield. The second model improved upon the first by considering a source distribution and the varying distance between the source and surface of the shield. Both D-T and D-D fuel cases were run with the model and the resulting mass estimates were used to compare the specific mass to the state-of-the-art technology.
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    Development and Validation of an NPSS Model of a Small Turbojet Engine
    (2017) Vannoy, Stephen; Cadou, Christopher P.; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
    Recent studies have shown that integrated gas turbine engine (GT)/solid oxide fuel cell (SOFC) systems for combined propulsion and power on aircraft offer a promising method for more efficient onboard electrical power generation. However, it appears that nobody has actually attempted to construct a hybrid GT/SOFC prototype for combined propulsion and electrical power generation. This thesis contributes to this ambition by developing an experimentally validated thermodynamic model of a small gas turbine (~230 N thrust) platform for a bench-scale GT/SOFC system. The thermodynamic model is implemented in a NASA-developed software environment called Numerical Propulsion System Simulation (NPSS). An indoor test facility was constructed to measure the engine’s performance parameters: thrust, air flow rate, fuel flow rate, engine speed (RPM), and all axial stage stagnation temperatures and pressures. The NPSS model predictions are compared to the measured performance parameters for steady state engine operation.
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    Development and Application of Theoretical Models for Rotating Detonation Engine Flowfields
    (2016) Fievisohn, Robert; Yu, Kenneth; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
    As turbine and rocket engine technology matures, performance increases between successive generations of engine development are becoming smaller. One means of accomplishing significant gains in thermodynamic performance and power density is to use detonation-based heat release instead of deflagration. This work is focused on developing and applying theoretical models to aid in the design and understanding of Rotating Detonation Engines (RDEs). In an RDE, a detonation wave travels circumferentially along the bottom of an annular chamber where continuous injection of fresh reactants sustains the detonation wave. RDEs are currently being designed, tested, and studied as a viable option for developing a new generation of turbine and rocket engines that make use of detonation heat release. One of the main challenges in the development of RDEs is to understand the complex flowfield inside the annular chamber. While simplified models are desirable for obtaining timely performance estimates for design analysis, one-dimensional models may not be adequate as they do not provide flow structure information. In this work, a two-dimensional physics-based model is developed, which is capable of modeling the curved oblique shock wave, exit swirl, counter-flow, detonation inclination, and varying pressure along the inflow boundary. This is accomplished by using a combination of shock-expansion theory, Chapman-Jouguet detonation theory, the Method of Characteristics (MOC), and other compressible flow equations to create a shock-fitted numerical algorithm and generate an RDE flowfield. This novel approach provides a numerically efficient model that can provide performance estimates as well as details of the large-scale flow structures in seconds on a personal computer. Results from this model are validated against high-fidelity numerical simulations that may require a high-performance computing framework to provide similar performance estimates. This work provides a designer a new tool to conduct large-scale parametric studies to optimize a design space before conducting computationally-intensive, high-fidelity simulations that may be used to examine additional effects. The work presented in this thesis not only bridges the gap between simple one-dimensional models and high-fidelity full numerical simulations, but it also provides an effective tool for understanding and exploring RDE flow processes.
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    Simulation of Dual-Mode Scramjet Under Thermally Choked vs. Supersonic Combustion Mode
    (2014) Butcher, Cameron; Yu, Kenneth H; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
    Effects of combustion mode and cavity flame-holder on dual-mode scramjet performance were investigated using a two-dimensional computational framework developed from commercial finite element software. The objectives were to simulate the experimental data from a laboratory model scramjet with mixing enhancement device, provide better understanding of the physical processes, and to analyze the quantitative effects on the potential performance. The isolator flow field was modeled separately to match the experimentally obtained pressure rise during the Mach 2.1 isolator entry condition. The combustor heat release distribution was systematically adjusted to reproduce the wall pressure distributions from the experiments. Case studies were conducted with and without the presence of the wall cavity for scramjet operation under both thermally-choked and supersonic-combustion mode. The combustion mode affected potential tradeoffs between thrust increase and higher thermal protection need. The presence of the cavity dampened the extent of the tradeoffs by reducing the temperature change.
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    SIMULATION AND MODELING OF AN ACOUSTICALLY FORCED MODEL ROCKET INJECTOR
    (2010) Gers, David; Yu, Ken; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
    A numerical and experimental study was performed to assess the capability of the Loci-CHEM CFD solver in simulating dynamic interaction between hydrogen-oxygen turbulent diffusion flames and periodic pressure waves. Previous experimental studies involving a single-element shear-coaxial model injector revealed an unusual flame-acoustic interaction mechanism affecting combustion instability characteristics. To directly compare the simulation and experiments, various models in the present solver were examined and additional experiments conducted. A customized mesh and corresponding boundary conditions were designed and developed, closely approximating the experimental setup. Full 3-D simulations were conducted using a hybrid RANS/LES framework with appropriate chemistry and turbulence models. The results were compared for both reacting and non-reacting flows that were excited at various forcing frequencies representing both resonant and non-resonant behaviors. Although a good qualitative agreement was obtained for the most part, there was a significant discrepancy in simulating the flame-acoustic interaction behavior observed under non-resonant forcing conditions.
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    Axisymmetric Inlet Design for Combined Cycle Engines
    (2005-05-03) Colville, Jesse; Lewis, Mark J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
    Performance considerations for a turbine-based combined-cycle engine inlet are presented using the inlet of the Lockheed SR-71 as a baseline. A numerical model is developed using the axisymmetric method of characteristics to perform full inviscid flow analysis, including any internal shock reflections. Self-starting characteristics are quantified based upon the Kantrowitz limit. The original SR-71 inlet is analyzed throughout the designed self-starting regime, beginning at Mach 1.7 and ending with the shock-on-lip condition at Mach 3.2. The characteristics model is validated using computational fluid dynamics. A series of modifications are then considered for their ability to extend the range of the inlet into the hypersonic flight regime. Self-starting characteristics of these new designs are also characterized; results indicate that two new designs can maintain self-starting capability into the Mach 6-7 range. Full external and internal flow properties of the new designs are determined using the characteristics model. Mach number, total pressure ratio, temperature, pressure and mass flow properties (and their levels of distortion) are quantified at the inlet exit plane for all cases considered.