The desire to gather acoustic data of a propeller in simulated hover without environmental effects is met with the challenge of recirculating flow in a closed anechoic chamber. Flow recirculation in an anechoic chamber is not ideal because the effects of it being ingested by the rotor cause unsteady loadings on the propeller. This unsteady loading causes vortex noise from the turbulence introduced by the recirculation of the flow. The aim is to study the acoustic signature of a propeller separate from external factors and engine noise. Two different propellers with three different screen combinations were tested in a closed anechoic chamber. The results of these tests showed that the screens are able to reduce the vortex and broadband noise levels, with the reduction being more pronounced when the dual screen was installed.

Scramjet engines are key for sustained hypersonic flights. Analytic models play a critical role in the preliminary design of a scramjet engine configuration. The objective of this research is to develop and validate a quasi-1D model for the scramjet engine encompassing inlet, isolator and combustor, to evaluate the impact of flight conditions and design parameters on the engine functionality. The model is developed assuming isentropic flow in the inlet with a single turn; modified Fanno-flow equations in the isolator that account for the area change of the core flow; and the combustor is modeled using Heiser-Pratt equations accounting for the fuel mixing efficiency. The isolator and combustor models are validated against experimental results. The model accounts for twelve parameters allowing for a decent range of possible configurations. Finally, the model was applied to five sets of parametric studies to evaluate the effect of multiple parameters on the engine functionality.

Advanced aerospace materials require extensive testing and characterization to anticipate and ensure their integrity under hostile environments. Characterization methods utilizing synchrotron X-Ray diffraction and spectroscopy can decrease the time required to determine an emerging material's readiness for application through intrinsic information on the material response and failure mechanisms. In this study, thermal barrier coating samples applicable to turbine blades of jet engines were studied using Raman and Photoluminescence spectroscopy as well as Synchrotron X-ray diffraction while Kevlar based fiber composites applicable to ballistic resistant armor were studied using Raman spectroscopy to investigate the mechanical state and corresponding damage and failure mechanisms. Piezospectroscopic studies on the stress state of the thermally grown oxide (TGO) within the thermal barrier coatings, on a hollow cylindrical specimen, provided results that indicate variations within the TGO. Comparison of measured photo-luminescence spectra of the specimen before and after long duration thermal aging showcases the development of the system and the initiation of micro-damage. Raman spectroscopy performed on Kevlar ballistic composites with nano-scale additives, presented insight into the additives' role in load transfer and damage propagation through a comparison of the shift in optical spectra to that of the pristine fibers. The results presented herein utilize changes in the measured emission from these non-destructive testing techniques to link the phenomena with material response. Techniques to optimize imaging and spectral collection are addressed as well. The findings will advance the use of the techniques in the development of aerospace materials, providing a more complete understanding of land and aircraft turbine blade coatings, and fiber composite response to complex loading.

Planar Laser Induced Fluorescence (PLIF) with acetone seeding was applied to measure the scalar fields of an axisymmetric freejet and an inclined jet-in-crossflow as applicable to film cooling. From the scalar fields, jet-mixing and trajectory characteristics were obtained. In order to validate the technique, the canonical example of a nonreacting freejet of Reynolds Numbers 900-9000 was investigated. Desired structural characteristics were observed and showed strong agreement with computational modeling. After validating the technique with the axisymmetric jet, the jet-in-crossflow was tested with various velocity ratios and jet injection angles. Results indicated the degree of wall separation for different injection angles and demonstrate both the time-averaged trajectories as well as select near-wall concentration results for varying jet momentum fluxes. Consistent with literature findings, the orthogonal jet trajectory for varying blowing ratios collapsed when scaled by the jet-to-freestream velocity ratio and hole diameter, rd. Similar collapsing was demonstrated in the case of a non-orthogonal jet. Computational Fluid Dynamic (CFD) simulations using the OpenFOAM software was used to compare predictions with select experimental cases, and yielded reasonable agreement. Insight into the importance and structure of the counter rotating vortex pair and general flow field turbulence was highlighted by cross validation between CFD and experimental results.

Ocean energy research has grown in popularity in the past decade and has produced various designs for wave energy extraction. This thesis focuses on the performance analysis of a uni-directional impulse turbine for wave energy conversion. Uni-directional impulse turbines can produce uni-directional rotation in bi-directional flow, which makes it ideal for wave energy extraction as the motion of ocean waves are inherently bi-directional. This impulse turbine is currently in use in four of the world’s Oscillating Wave Columns (OWC). Current research to date has documented the performance of the turbine but little research has been completed to understand the flow physics in the turbine channel. An analytical model and computational fluid dynamic simulations are used with reference to experimental results found in the literature to develop accurate models of the turbine performance. To carry out the numerical computations various turbulence models are employed and compared. The comparisons indicate that a low Reynolds number Yang-shih K-Epsilon turbulence model is the most computationally efficient while providing accurate results. Additionally, analyses of the losses in the turbine are isolated and documented. Results indicate that large separation regions occur on the turbine blades which drastically affect the torque created by the turbine, the location of flow separation is documented and compared among various flow regimes. The model and simulations show good agreement with the experimental results and the two proposed solutions enhance the performance of the turbine showing an approximate 10% increase in efficiency based on simulation results.

Flow boiling heat transfer allows for the dissipation of large amounts of heat. In this work, the effect of heat flux and pressure on flow boiling of liquid refrigerant R-134a is studied in a vertical thin channel. The experimental setup mimics a refrigeration cycle and specifically looks at the effect of pressure and wall heat flux on the departure size and bubble generation rate. The experimental setup consists of a closed loop which includes a vertical narrow rectangular channel and two synchronized high speed cameras for optical measurements at either sides of the channel. The setup is built to employ an accurate measurement technique to define wall temperatures of the representative flow boiling process. Instead of using thermocouples on the surface channel, the thermochromic liquid crystallography (TLC) technique is used to determine non-invasively the heater surface temperature at high temporal and spatial resolution. The TLC interval range is 30-50°C. The TLC is attached to a Fecralloy heating section. The high speed Prosilica cameras simultaneously capture, colored TLC images as well as bubble nucleation and departure at very high frame rates. Experiments on subcooled flow boiling heat transfer have been conducted with refrigerant R-134a under a mass flux range of 484.838 kg/m2 s to 1212.1 kg/m2 s. With the low mass flux, the wall heat flux ranged from 167.2 to 672.1 kW/m2 , the inlet subcooling ranged from 0.35°C to 16.55 °C, the system pressure ranged from 621 kPa to 1034 kPa. At high mass flux, the wall heat flux ranged from 329.8 kW/m2 to 744 kW/m2 , the iv inlet subcooling from 0.16°C to 17.21 °C, and the system pressure from 621 kPa to 1034 kPa. A parametric study was done by maintaining various input parameters constant. From the high speed images, bubble parameters such as size and frequency are calculated. Temperature contours are utilized to determine the surface wall temperature at specific points. Sequential wall temperatures are traced over a short period of time to understand the cooling effects. The bubble propagation and coalescence are also visualized. Results show that bubble size and frequency increased with heat flux at any particular pressure. At higher pressure, the trend would be for the bubble size to decrease; however, the inlet subcooling and heat flux also affect bubble size. The bubble frequency is also seen to be affected by the inlet subcooling and the heat flux. Even though the inlet subcooling is maintained approximately constant, any slight decrease in subcooling increased bubble growth rate. Another trend that is observed is that at higher the heat flux, the bubble generation frequency is faster; however no specific trend is observed for wall superheat. With an increase in heat flux, the wall superheats are expected to increase; however, the localized nature of the nucleation activity sites is seen to affect the results. The variables are non-dimensionalized to note trends in parameters. In summary, the data analysis demonstrates that both heat flux and pressure significantly influence the bubble generation rate, size, propagation and coalescence.

The combustion of ultra-lean fuel to air mixtures provides an efficient way to convert the chemical energy of hydrocarbons into useful power. Conventional burning techniques of a mixture have defined flammability limits beyond which a flame cannot self-propagate due to heat losses. Matrix stabilized porous medium combustion is an advanced technique in which a solid porous matrix within the combustion chamber accumulates heat from the hot gaseous products and preheats incoming reactants. This heat recirculation extends the standard flammability limits and allows the burning of ultra-lean fuel mixtures, conserving energy resources, or the burning of gases of low calorific value, utilizing otherwise wasted resources. The heat generated by the porous burner can be harvested with thermoelectric devices for a reliable method of generating electricity for portable electronic devices by the burning of otherwise noncombustible mixtures. The design of the porous media burner, its assembly and testing are presented. Highly porous (~80% porosity) alumina foam was used as the central media and alumina honeycomb structure was used as an inlet for fuel and an outlet for products of the methane-air combustion. The upstream and downstream honeycomb structures were designed with pore sizes smaller than the flame quenching distance, preventing the flame from propagating outside of the central section. Experimental results include measurements from thermocouples distributed throughout the burner and on each side of the thermoelectric module along with associated current, voltage and power outputs. Measurements of the burner with catalytic coating were obtained for stoichiometric and lean mixtures and compared to the results obtained from the catalytically inert matrix, showing the effect on overall efficiency for the combustion of fuel-lean mixtures

With the pursuit of more advanced and environmentally-friendly technologies of today’s society, the airline industry has been pushed further to investigate solutions that will reduce airport noise and congestion, cut down on emissions, and improve the overall performance of aircraft. These items directly influence airport size (runway length), flight patterns in the community surrounding the airport, cruise speed, and many other aircraft design considerations which are setting the requirements for next generation aircraft. Leading the research in this movement is NASA, which has set specific goals for the next generation regional airliners and has categorized the designs that meet the criteria as Cruise Efficient Short Takeoff and Land (CESTOL) aircraft. With circulation control (CC) technology addressing most of the next generation requirements listed above, it has recently been gaining more interest, thus the basis of this research. CC is an active flow control method that uses a thin sheet of high momentum jet flow ejected over a curved trailing edge surface and in turn utilizes Coanda effect to increase the airfoil’s circulation, augmenting lift, drag, and pitching moment. The technology has been around for more than 75 years, but is now gaining more momentum for further development due to its significant payoffs in both performance and system complexity. The goal of this research was to explore the design of the CC flap shape and how it influences the local flow field of the system, in attempt to improve the performance of existing CC flap configurations and provide insight into the aerodynamic characteristics of the geometric parameters that make up the CC flap. Multiple dual radius flaps and alternative flap geometry, prescribed radius, flaps were developed by varying specific flap parameters from a baseline dual radius flap configuration that had been previously developed and researched. The aerodynamics of the various flap geometries were analyzed at three different flight conditions using two-dimensional CFD. The flight conditions examined include two low airspeed cases with blown flaps at 60° and 90° of deflection, and a transonic cruise case with no blowing and 0° of flap deflection. Results showed that the shorter flaps of both flap configurations augmented greater lift for the low airspeed cases, with the dual radius flaps producing more lift than the corresponding length prescribed radius. The large lift generation of these flaps was accompanied by significant drag and negative pitching moments. The incremental lift per drag and moment produced was best achieved by the longer flap lengths, with the prescribed radius flaps out-performing each corresponding dual radius. Longer flap configurations also upheld the better cruise performance with the least amount of low airspeed flow, drag, and required angle of attack for a given cruise lift coefficient. The prescribed radius flaps also presented a favorable trait of keeping a more continuous skin friction distribution over the flap when the flaps were deflected, where all dual radius configurations experienced a distinct fluctuation at the location where the surface curvature changes between its two radii. The prescribed radius flaps displayed a similar behavior when the flaps were not deflected, during the cruise conditions analyzed. Performance trends for the different flap configurations, at all three flight conditions, are presented at the end of each respective section to provide guidance into the design of CC geometry. The results of the presented research show promise in modifying geometric surface parameters to yield improved aerodynamics and performance.

When simulating complex flows, there are some physical situations that exhibit large fluctuations in particle density such as: planetary reentry, ablation due to arcing, rocket exhaust plumes, etc. When simulating these events, a high level of physical accuracy can be achieved with kinetic methods otherwise known as particle methods. However, this high level of physical accuracy requires large amounts of computation time. If the simulated flow is in collisional equilibrium, then less computationally intensive continuum methods, otherwise known as fluid methods, can be utilized. Hybrid Particle-Continuum (HPC) codes attempt to blend particle and fluid solutions in order to reduce computation time for transitional flows that exhibit both continuum and rarefied flow in a single domain. This thesis details the development of an HPC code in OpenFoam for Cal Poly's Aerospace Engineering department. The primary benchmark for the solver, named hybridFoam, was to simulate a 1D sod-shock simulation. This primary goal was achieved and a collection of test simulations were conducted to map out the solvers current capabilities and identify where future development efforts should focus.

A new parametric model for the growth rate of turbulent mixing layers is proposed. A database of experimental and numerical mixing layer studies was extracted from the literature to support this effort. The domain of the model was limited to planar, spatial, nonreacting, free shear layers that were not affected by artificial mixing enhancement techniques. The model is split into two parts which were each tuned to optimally fit the database; equations for an incompressible growth rate were derived from the error function velocity profile, and a function for a compressibility factor was generalized from existing theory on the convective Mach number. The compressible model is supported by a detailed evaluation of the currently accepted models and practices, including error analysis of the convective Mach number derivation and a critical analysis of Slessor’s re-normalization technique which affected his 1998 compressibility parameter. Analysis of the database suggested that a distinction should be made between thickness definitions that are based on the velocity profile and those based on the density profile. Additionally, the accumulation of different normalization approaches throughout the literature was shown to have introduced non-physical variance in the trends. Resolution of this issue through a consistent normalization process has greatly improved the normality and scatter of the data and the goodness-of-fit of the models, resulting in R2 = 0.9856 for the incompressible model and R2 = 0.9004 for the compressible model.