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Mora Sánchez, Pablo AInvestigation of the Noise Radiation from Heated Supersonic Jets
PhD, University of Cincinnati, 2016, Engineering and Applied Science: Aerospace Engineering
This work focuses in the investigation of crackle and Mach wave radiation in heated supersonic jets. The skewness and kurtosis of the acoustic pressure signal and its time derivative were adopted as metrics for identifying crackling jets and quantifying levels of crackle. Cold and heated jets from supersonic nozzles with different geometric parameters and scales are analyzed to draw conclusions on noise sources and propagation. In order to complement the investigation, results are also presented for the mixing noise, broadband shock-associated noise and screech. Chapter 4 focuses on the impact of jet operating condition on the skewness and kurtosis levels of a jet issuing from a converging-diverging conical nozzle, with a 1.5 design Mach number. An increase in convective Mach number, achieved by increasing jet temperature, proved to be related to elevated values of OASPL, skewness, and kurtosis, in both the near and far fields. Intense levels of the dP/dt high-order statistics appear to be generated at different locations in the shear layer of the jet and strengthen away from the jet by non-linear propagation effects. Chapter 5 studies how adding chevrons to a converging-diverging nozzle impacts Mach wave radiation and crackle. The chevrons decreased OASPL in the downstream angles but increased the broadband shock-associated noise. Pressure skewness, dP/dt skewness and kurtosis were all reduced by the chevrons in the near field and far field, and thus they effectively mitigated crackle and Mach wave radiation; however, chevrons showed no evidence of changing the convective Mach number. The evolution of noise signals was analyzed in the near-field to the far-field to identify the strengthening of skewness through nonlinear propagation effects. Chapter 6 investigates a jet exhausting over a plate at different stand-off distances, to simulate jets exhausting over airframe surfaces and jet-ground interaction during take-off and landing operations. Far-field acoustics were measured at the reflected direction, sideline, and shielded azimuthal directions. At the sideline, the plate attached to the nozzle lip diminished broadband shock-associated noise, and mitigated screech for the cold case. When the plate was moved away from the nozzle, screech tones were intensified at the under-expanded condition. Crackle levels were significantly intensified in the sideline, within a range of stand-off positions. Chapter 7 analyzes the impact of nozzle scale and nozzle internal contours on the levels of crackle. Three scaled converging-diverging nozzles, with jet exit diameters of 0.542 in, 0.813 in, and 1.085 in were investigated. Far-field arrays were setup at a constant non-dimensionalized radial distance of 40 nozzle exit diameters. The pressure skewness and kurtosis plots collapsed for all three scaled nozzles when the pressure signals were not filtered. The dP/dt statistics collapsed when the signals were downsampled proportional to the nozzle exit diameters. Baseline nozzle results were also compared to a smooth contoured nozzle designed by the Method of Characteristics. This nozzle almost had no broadband shock-associated noise, but contained the same skewness and kurtosis levels, concluding that crackle is not linked to the shock-cell structures in the jet.

Committee:

Ephraim Gutmark, Ph.D. D.Sc. (Committee Chair); Kailas Kailasanath, Ph.D. (Committee Member); Jeffrey Kastner, Ph.D. (Committee Member); Mark Turner, Sc.D. (Committee Member)

Subjects:

Aerospace Materials

Keywords:

Supersonic Jet Noise;Crackle;Mach Wave Radiation;Supersonic Heated Jets

Heeb, Nicholas S.Azimuthally Varying Noise Reduction Techniques Applied to Supersonic Jets
PhD, University of Cincinnati, 2015, Engineering and Applied Science: Aerospace Engineering
An experimental investigation into the effect of azimuthal variance of chevrons and fluidically enhanced chevrons applied to supersonic jets is presented. Flow field measurements of streamwise and cross-stream particle imaging velocimetry were employed to determine the causes of noise reduction, which was demonstrated through acoustic measurements. Results were obtained in the over- and under- expanded regimes, and at the design condition, though emphasis was placed on the overexpanded regime due to practical application. Surveys of chevron geometry, number, and arrangement were undertaken in an effort to reduce noise and/or incurred performance penalties. Penetration was found to be positively correlated with noise reduction in the overexpanded regime, and negatively correlated in underexpanded operation due to increased effective penetration and high frequency penalty, respectively. The effect of arrangement indicated the beveled configuration achieved optimal abatement in the ideally and underexpanded regimes due to superior BSAN reduction. The symmetric configuration achieved optimal overexpanded noise reduction due to LSS suppression from improved vortex persistence. Increases in chevron number generally improved reduction of all noise components for lower penetration configurations. Higher penetration configurations reached levels of saturation in the four chevron range, with the potential to introduce secondary shock structures and generate additional noise with higher number. Alternation of penetration generated limited benefit, with slight reduction of the high frequency penalty caused by increased shock spacing. The combination of alternating penetration with beveled and clustered configurations achieved comparable noise reduction to the standard counterparts. Analysis of the entire data set indicated initial improvements with projected area that saturated after a given level and either plateaued or degraded with additional increases. Optimal reductions were 3-7dB depending on operating condition and observation angle. The fluidic enhancement of the low penetration chevrons indicated significant improvement in the overexpanded regime, with detrimental effect at higher conditions. Improvements were generally due to shock noise and turbulent mixing noise reductions caused by decreased shock strength and LSS growth inhibition. Investigation of azimuthal configurations indicated further improvements were achieved by the clustered configuration due to additional BSAN reductions caused by drastic modification of the shock cell structure due to elliptification of the jet cross section.

Committee:

Ephraim Gutmark, Ph.D. D.Sc. (Committee Chair); James Bridges, Ph.D. (Committee Member); Kailas Kailasanmath, Ph.D. (Committee Member); Steve Martens, Ph.D. (Committee Member); Shaaban Abdallah, Ph.D. (Committee Member); Paul Orkwis, Ph.D. (Committee Member)

Subjects:

Aerospace Materials

Keywords:

Supersonic Jet;Aeroacoustics;Jet Noise;Chevrons;Fluidic Injection;Compressible Flow

MURUGAPPAN, SHANMUGAMINNOVATIVE TECHNIQUES TO IMPROVE MIXING AND PENETRATION IN SCRAMJET COMBUSTORS
PhD, University of Cincinnati, 2005, Engineering : Aerospace Engineering
Scramjet combustors are characterized by an extremely short residence time for the completion of fuel atomization, mixing and combustion. It is therefore desired to develop fuel injection schemes that will accelerate the mixing process by improving penetration, achieving small-dispersed fuel droplets, and enhancing mixing. To date, scramjet-fueling schemes require multiple fuel-injection locations in conjunction with flame-holding cavities. This thesis involves the design, development and testing of two innovative fuel injection strategies that could be used to precisely enhance mixing and control the fuel distribution in a scramjet combustor. The first approach employs a high frequency-forcing injector where the amplitude and frequency of excitation was chosen to affect both jet spread and penetration depth, independently, by excitation at different jet instability modes. Another novel fuel injection technique is to employ a vortex-breakdown-controlled injector, which includes a swirling annular jet combined with a control non-swirling jet. The vortex is stabilized by injection of a continuous jet near the core of the vortex. When the control jet introduces additional momentum along the vortex mean flow, the breakdown is delayed. This has been used to control both the intensity of vortex breakdown and penetration of the jet, which could be effectively applied to enhance mixing and penetration.

Committee:

Dr. Ephraim Gutmark (Advisor)

Subjects:

Engineering, Aerospace

Keywords:

Supersonic; Mixing; Scramjets; Fuel Injection

Xiao, RuiyangThe Freezing of Highly Sub-cooled H2O/D2O Droplets
Master of Science, The Ohio State University, 2008, Environmental Science
The condensation of H2O and D2O in a supersonic Laval nozzle was investigated at different stagnation condition by using Pressure Trace Measurements (PTM) and Fourier Transformation Infrared (FTIR) spectroscopy. PTM determined several key properties highly related to nucleation such as the temperature and pressure corresponding to the onset of condensation, Ton, pon as well as the temperature and pressure corresponding to the maximum nucleation rate TJmax and Jmax. Moreover, the results from PTM provide important information for the FTIR study. The FTIR spectra of D2O and H2O nanodroplets in N2 carrier gas were measured in our nozzle. The observed spectra of D2O droplets had some clear peaks, and the shapes of the spectra changed as a function of flow rates and position in the nozzle. The broad peak of D2O between 2400 cm-1 and 2600 cm-1 was due to ν1, ν3, and overtone of ν2 in the liquid phase, and its peak area was correlated to the product of the weight fraction of condensate (g) and the density of the flowing mixture (ρ), values derived from PTM. There is good correlation between the peak area and g*ρ (p<0.001). From our FTIR H2O nanodroplets study, the same trends regarding spectral changes and flow rate were observed. Moreover, the first observation of cubic ice in our supersonic nozzle was made by FTIR spectroscopy at a location x =6 cm from the throat. The peak in the spectra was located at a frequency of 3250 cm-1. This result is consistent with previous FTIR and electron diffraction scattering studies of H2O nanodroplets done by Buch (Buch V., Bauerecker S., Devlin J. P., Buck U., and Kazimirski J. K. 2004. Int. Rev. Phys. Chem. 23. 375-433) and Huang (Huang J. F. and Bartell L. S. 1995, J. Phys. Chem. 99. 3924-3931), respectively. To determine the freezing rate from liquid phase to cubic ice, however, requires further optimization of the experimental setup and more quantitative study

Committee:

Barbara Wyslouzil, PhD (Advisor); Heather Allen, PhD (Advisor); Linda Weavers, PhD (Committee Member)

Subjects:

Chemistry

Keywords:

sub-cooled droplets; FTIR; supersonic nozzle

Clark, Kylen D.A Numerical Comparison of Symmetric and Asymmetric Supersonic Wind Tunnels
MS, University of Cincinnati, 2015, Engineering and Applied Science: Aerospace Engineering
Supersonic wind tunnels are a vital aspect to the aerospace industry. Both the design and testing processes of different aerospace components often include and depend upon utilization of supersonic test facilities. Engine inlets, wing shapes, and body aerodynamics, to name a few, are aspects of aircraft that are frequently subjected to supersonic conditions in use, and thus often require supersonic wind tunnel testing. There is a need for reliable and repeatable supersonic test facilities in order to help create these vital components. The option of building and using asymmetric supersonic converging-diverging nozzles may be appealing due in part to lower construction costs. There is a need, however, to investigate the differences, if any, in the flow characteristics and performance of asymmetric type supersonic wind tunnels in comparison to symmetric due to the fact that asymmetric configurations of CD nozzle are not as common. A computational fluid dynamics (CFD) study has been conducted on an existing University of Michigan (UM) asymmetric supersonic wind tunnel geometry in order to study the effects of asymmetry on supersonic wind tunnel performance. Simulations were made on both the existing asymmetrical tunnel geometry and two axisymmetric reflections (of differing aspect ratio) of that original tunnel geometry. The Reynolds Averaged Navier Stokes equations are solved via NASA’s OVERFLOW code to model flow through these configurations. In this way, information has been gleaned on the effects of asymmetry on supersonic wind tunnel performance. Shock boundary layer interactions are paid particular attention since the test section integrity is greatly dependent upon these interactions. Boundary layer and overall flow characteristics are studied. The RANS study presented in this document shows that the UM asymmetric wind tunnel/nozzle configuration is not as well suited to producing uniform test section flow as that of a symmetric configuration, specifically one that has been scaled to have equal aspect ratio. Comparisons of numerous parameters, such as flow angles, pressure (both static and stagnation), entropy, boundary layers and displacement thickness, vorticity, etc. paint a picture that shows the symmetric equal aspect ratio configuration to be decidedly better at producing desirable test section flow. It has been shown that virtually all parameters of interest are both more consistent and have lower deviation from ideal conditions for the symmetric equal area configuration.

Committee:

Paul Orkwis, Ph.D. (Committee Chair); Shaaban Abdallah, Ph.D. (Committee Member); Mark Turner, Sc.D. (Committee Member)

Subjects:

Aerospace Materials

Keywords:

Computational Fluid Dynamics;Wind tunnels;Asymmetric Supersonic Nozzles;Test section flow;flow uniformity;Converging Diverging Nozzles

Pathak, HarshadNucleation and Droplet Growth During Co-condensation of Nonane and D2O in a Supersonic Nozzle
Doctor of Philosophy, The Ohio State University, 2013, Chemical and Biomolecular Engineering
Raw natural gas consists mainly of methane and has impurities like water vapor, higher alkanes, H2S etc. Dehydration of natural gas is important to prevent hydrate formation in pipelines carrying natural gas over long distances. Traditionally, dehydration is done using chemical methods like pressure swing absorption and glycol dehydration. An alternate method of dehydration is by using a mechanical process of supersonic separation. In this method, raw natural gas is cooled down by adiabatic expansion resulting in condensation of water vapor and higher alkanes. The goal of this work is to understand the nucleation and droplet growth when droplet sizes are of the order of nm and timescales are of the order of microseconds when water and alkanes, two substances which are immiscible, condense together. We use supersonic nozzles in this work where cooling rates are of the order of 105-106 K/s. The supersonic velocities of the flow enable measurements on a resolution of the order of microseconds. Pressure trace measurement (PTM) is our basic experimental technique and it characterizes the flow by measuring the pressure profile inside the supersonic nozzle as the vapor-gas mixture expands and vapor condenses inside the nozzle. These experiments give the initial estimate of temperature, density, velocity and mass fraction of the condensate. We use Fourier transform infrared spectroscopy (FTIR) to get the composition of the condensed liquid/vapor. To determine the amount of nonane condensed, we fit the measured spectrum of nonane to a linear combination of a well-characterized vapor and liquid spectrum. For D2O analysis, we calculate D2O vapor concentration by analyzing the vibrational-rotational spectrum of O-D stretch region. The size and number of droplets is characterized using small angle x-ray scattering (SAXS) that are performed in Argonne National Laboratory. The nucleation rates for pure D2O and nonane agree with previous measurements done by other researchers. The subsequent process of growth of the droplets can be sensitive to droplet temperatures Td. For pure nonane droplets, we observe that Td is not important enough to alter the growth rates unlike pure D2O. The growth of D2O droplets is further affected by coagulation once condensation has slowed down. We also observe that when nonane and D2O both are condensing, the presence of nonane inhibits D2O condensation even when D2O dominates the nucleation process. Prediction of the droplet structure of composite nonane-D2O droplets is challenging because the SAXS spectra of these droplets does not fit to standard shapes like spheres or core-shell structures. The small size of these droplets makes it possible to study them through molecular dynamics simulations. Our collaborators conduct simulations of these droplets and calculate the scattering behavior for those shapes. The SAXS spectra are fit to scattering from shapes derived from both density functional theory (DFT) calculations and molecular dynamics (MD) simulations. Although the lens-on-sphere structures derived from MD simulations fits the scattering spectra better than all other structures which we tested, the overall composition from this structure predicts that the amount of D2O condensed is 30-40% less than that measured from FTIR.

Committee:

Barbara Wyslouzil (Advisor); Isamu Kusaka (Committee Member); Bhavik Bakshi (Committee Member)

Subjects:

Chemical Engineering; Chemistry; Energy; Engineering; Experiments; Physics

Keywords:

nonane; D2O; supersonic nozzle; condensation; droplet growth; FTIR; SAXS; non-equilibrium, nucleation, droplet structure, guinier analysis, aerosol spectroscopy, natural gas dehydration

Cuppoletti, Daniel RSupersonic Jet Noise Reduction with Novel Fluidic Injection Techniques
PhD, University of Cincinnati, 2013, Engineering and Applied Science: Aerospace Engineering
Supersonic jets provide unique challenges in the aeroacoustic field due to very high jet velocities, shock associated noise components, flow dependence on jet expansion, and stringent performance requirements. Current noise suppression technology for commercial and military jet engines revolves around using chevrons or mechanical vortex generators to increase mixing near the nozzle exit, subsequently reducing peak turbulence levels in the mixing region. Passive noise control methods such as mechanical chevrons cause thrust loss throughout the flight envelope and performance can vary with the engine operating condition. Development of active noise control methods have the potential of improved performance throughout the flight envelope and the benefit of being deactivated when noise control is unnecessary. Fluidic injection of air into a supersonic jet is studied as an active control method with an emphasis on understanding the physics of the problem and identifying the controlling parameters. An experimental investigation with computational collaboration was conducted to understand the effect of nozzle design on supersonic jet noise and to develop various fluidic injection techniques to control noise from a supersonic jet with a design Mach number of 1.56. The jet was studied at overexpanded, ideally expanded, and underexpanded conditions to evaluate the effects throughout the operational envelope. As a passive noise control method, the internal contour of a realistic nozzle was modified to investigate the effect on acoustics and performance. Thrust was improved up to 10% with no acoustic penalties through nozzle design, however it was found that the shock noise components were highly sensitive to the shock structure in the jet. Steady fluidic injection was used to generate vorticity at the trailing edge of the nozzle showing that noise reduction is achieved through vorticity generation, modification of the shock structure, and interference with the screech feedback mechanism by decoupling the phase relationship between jet turbulence and shock spacing. Reduction of shock noise was found to be optimum at an intermediate injection pressure due to shock weakening from the fluidic injectors and injector interactions with the jet shock-expansion structure. Large-scale mixing noise reduction was shown to depend on the vorticity strength and circulation. Unprecedented reduction of OASPL up to -8.5 dB were achieved at the peak noise direction through strong jet mixing and rapid collapse of the potential core. Pulsed fluidic injection was investigated to understand the acoustic benefits and drawbacks of unsteady injection. Valve frequency response up to 500 Hz was achieved but noise reduction dropped off above 100 Hz due to poor flow response as verified by hot-wire and dynamic pressure measurements. At low pulse frequencies it was found that moderate noise reduction could be achieved with less flow than steady injection, but in general the mixing noise reduction scaled with the time integrated mass flow injection. It was discovered that the different components of supersonic jet noise had different characteristic response times to unsteady injection. Analysis of high speed shadowgraph images and acoustic spectra was used to identify time response of the jet during the unsteady injection cycle.

Committee:

Ephraim Gutmark, Ph.D., D.Sc. (Committee Chair); Steve Martens, Ph.D. (Committee Member); Awatef Hamed, Ph.D. (Committee Member); Jeffrey Kastner, Ph.D. (Committee Member); David Munday, Ph.D. (Committee Member); Mark Turner, Sc.D. (Committee Member)

Subjects:

Aerospace Materials

Keywords:

Jet Noise;Aeroacoustics;Fluid Dynamics;Particle Image Velocimetry;Acoustics;Supersonic

Bunnag, ShaneBleed Rate Model Based on Prandtl-Meyer Expansion for a Bleed Hole Normal to a Supersonic Freestream
MS, University of Cincinnati, 2010, Engineering and Applied Science: Aerospace Engineering
The presented work shows that Prandtl-Meyer expansion can be used as a foundation to predict bleed rate for a single bleed hole oriented normal to a supersonic freestream. A CFD study was used to explore flowfield phenomena that can be used in conjunction with Prandtl-Meyer expansion theory to improve model accuracy. Of these phenomena, the shear layer and barrier shock were the best defined and their geometric placement within the bleed hole were the basis for the bleed rate model. Coefficients of variation of the root mean square error between data and predictions were between 0.10 and 0.15 for all but the highest of freestream Mach numbers evaluated. Development of an analytical bleed rate model and recommendations for follow-on activity are presented.

Committee:

Awatef Hamed, PhD (Committee Chair); Prem Khosla, PhD (Committee Member); John Slater, PhD (Committee Member)

Subjects:

Aerospace Materials

Keywords:

boundary layer bleed;Prandtl-Meyer expansion;bleed rate model;shockwave boundary layer interaction;first principles;supersonic inlet

Rizzetta, Donald PasqualeAsymptotic solution for two-dimensional viscous supersonic and hypersonic flows past compression and expansion corners /
Doctor of Philosophy, The Ohio State University, 1976, Graduate School

Committee:

Not Provided (Other)

Subjects:

Engineering

Keywords:

Supersonic compressors;Asymptotes;Aerodynamics

Sinha, SomnathExperimental And Modeling Study Of Condensation In Supersonic Nozzles
Doctor of Philosophy, The Ohio State University, 2008, Chemical Engineering

The formation of aerosols by condensation from the vapor phase proceeds through nucleation, growth and aging stages. Of these, nucleation is the least well understood, hardest to predict and hardest to measure. Most of the experimental work on nucleation has been carried out using materials, for example water, that interact via complex intermolecular potentials. Recently, approaches based on density functional theory, molecular dynamics and Monte Carlo simulations, are becoming increasingly important in advancing nucleation theory. Most of these studies focus on noble gas condensation, usually Argon, because the intermolecular potentials for these species are relatively simple.

The primary goal of this work is to develop a new cryogenic supersonic nozzle apparatus to investigate Argon nucleation. The experimental data on Ar nucleation measured with the new setup is compared to the data available in the literature and the predictions of Classical Nucleation Theory (CNT). The results presented here resolve the prior inconsistency between the data from nucleation pulse chambers and supersonic nozzles at high pressures and temperatures. Comparing the data with the widely used CNT indicates that the predictions of this theory are about 12 orders of magnitude below the measured nucleation rates. Combining the current results with those of the cryogenic nucleation pulse chamber yields the size of the critical cluster n*. The experimental values of n* are 30 - 55% smaller than predicted by the Gibbs-Thomson equation and this may be the primary reason behind the failure of CNT. Since n* is not predicted correctly, correcting CNT by multiplying with a temperature dependent factor based on results from one device would not help in predicting nucleation rates for other experimental devices, where onset occurs at significantly different supersaturations.

In order to understand the entire condensation process, it is also important to understand growth and aging. In most devices, nucleation and growth occur simultaneously. Thus, in order to test the available growth laws it is necessary to develop models that incorporate both nucleation and growth. The second goal of this work is to modify and use a steady state 1-D model to examine the formation and growth of H2O/D2O droplets in a supersonic nozzle. The model incorporates a nucleation rate expression and one of five different growth models and predicts both the properties of the flow, i.e. pressure, temperature, density, etc, as well as the properties of the aerosol, i.e. the number density, average size, etc.. The flow properties are compared to the complimentary pressure trace measurements and the aerosol properties are compared to those derived from small angle x- ray scattering experiments. Contrary to expectations, the physically more realistic nonisothermal growth rate expressions over predict the number densities and, therefore, underpredict the average aerosol size. Surprisingly, under certain conditions, the isothermal growth rate expressions provide results closer to the experimental data. The failure of the nonisothermal growth expressions may be due to an overestimation of the mean droplet temperatures. Subtle effects related to changes in the boundary layer, not accounted for in a 1-D model, may also play a role in how nucleation is terminated.

Committee:

Barbara Wyslouzil, PhD (Advisor); Igor Adamovich, PhD (Committee Member); Isamu Kusaka, PhD (Committee Member)

Subjects:

Chemical Engineering; Chemistry; Physics

Keywords:

Nucleation; Condensation; Aerosols; Supersonic Nozzles

Hamza, AbdulhamidStructure and Exciton Coupling in Jet-Cooled Bichromophores
Doctor of Philosophy (Ph.D.), Bowling Green State University, 2008, Photochemical Sciences

The spectroscopy and exciton coupling in a series of conformationally flexible bichromophoric molecules have been investigated in a supersonic jet. The investigated molecules are: 1,2-diphenylethane (DPE), 1,2-bis(4-methylphenyl)ethane (M2DPE), 5,6,11,12-tetrahydrodibenzo[a,e]cyclooctene (THDC), cis and trans isomers of 1,2-diphenylcyclopropane (DPCP), 2-phenylindane(2PI), and 2-(4-fluorophenyl)indane (2FPI). Resonant enhanced two-photon ionization (R2PI) spectra of these compounds and many of their deuterated isotopomers have been recorded for the first time. The observed spectral features have been assigned. The experimental results are compared with the predictions of the dipole-dipole based Forster theory, and the supramolecular model of bichromophoric molecules with identical chromophores.

Analysis of the experimental data is facilitated by the spectral analysis of the single chromophore analogues of the investigated bichromophoric molecules. These include ethylbenzene-d0, α-ethylbenzene-d1, cis and trans isomers of 1-methyl-2-phenylcyclopropane, and 2-methylindane.

The molecular structures and other properties of the investigated molecules in their ground and exited singlet states have been computed at various levels of theory. The calculations predict energetic preference of localized electronic excitation of most of the investigated molecules, suggesting weak inter-chromophore interactions in the excited states. Transition density surfaces indicate that the lowest excited singlet states of most of the investigated molecules have Lb ππ* character.

Two conformers of DPE, M2DPE, THDC, and 2-methylindan have been observed. The rest of the investigated molecules are found to exist in a single conformation. The S2←S0 transitions of both anti and gauche conformers of DPE, the anti conformer of M2DPE, the chair conformer of THDC, and the trans isomer of DPCP are forbidden. However, the S2←S0 transitions of the asymmetrically deuterated isotopomers are observed, which allows for the estimation of the exciton splittings in these bichromophores. The exciton splittings are ca. three times higher in the more compact C2 symmetric gauche conformers of DPE and M2DPE than in the C2h symmetric anti conformers. The experimental exciton splittings of DPE conformers are higher than those of the corresponding M2DPE conformers. Dipole-dipole based Forster theory predicts higher splittings in M2DPE conformers, and incorrect ordering of the excited states of the anti conformers of DPE and M2DPE The exciton splitting in the C2 symmetric twist-boat conformer of THDC is seven times higher than that of the C2h symmetric chair conformer. The exciton splittings in THDC conformers are substantially larger than those of the structurally similar conformers of DPE and M2DPE.

Experimental evidence is provided in support of the conjugated nature of cyclopropyl ring. The trans isomer of DPCP has C2 symmetry. The spectrum of trans-DPCP is devoid of and like transitions. The cis isomer of DPCP is asymmetric. The S1←S0 and S2←S0 transitions of cis-DPCP are localized on the bisected and perpendicular phenyl rings, respectively. The S2←S0 minimum is calculated to lie just 60 cm-1 below the transition state along the phenyl torsional coordinates; hence, S2←S0 vibronic transitions are not observed in the spectrum of cis-DPCP. The S1←S0 and S2←S0 transitions of 2PI and 2FPI are found to be completely localized on the ortho-xylyl and phenyl (4-fluorophenyl) chromophores, respectively.

Committee:

John R. Cable (Advisor); Douglas C. Neckers (Committee Member); Sheila J. Roberts (Committee Member); Deanne L. Snavely (Committee Member)

Subjects:

Chemistry

Keywords:

spectroscopy; exciton coupling; supersonic jet; dimer; excited states; deuterated; 1,2-diphenylethane; 1,2-bis(4-methylphenyl)ethane; 5,6,11,12-tetrahydrodibenzo[a,e]cyclooctene; 1,2-diphenylcyclopropane; 2-phenylindane; 2-(4-fluorophenyl)indane

Benyo, Theresa L.Analytical and Computational Investigations of a Magnetohydrodynamic (MHD) Energy-Bypass System for Supersonic Turbojet Engines to Enable Hypersonic Flight
PHD, Kent State University, 2013, College of Arts and Sciences / Department of Physics
Historically, the National Aeronautics and Space Administration (NASA) has used rocket-powered vehicles as launch vehicles for access to space. A familiar example is the Space Shuttle launch system. These vehicles carry both fuel and oxidizer onboard. If an external oxidizer (such as the Earth's atmosphere) is utilized, the need to carry an onboard oxidizer is eliminated, and future launch vehicles could carry a larger payload into orbit at a fraction of the total fuel expenditure. For this reason, NASA is currently researching the use of air-breathing engines to power the first stage of two-stage-to-orbit hypersonic launch systems. Removing the need to carry an onboard oxidizer leads also to reductions in total vehicle weight at liftoff. This in turn reduces the total mass of propellant required, and thus decreases the cost of carrying a specific payload into orbit or beyond. However, achieving hypersonic flight with air-breathing jet engines has several technical challenges. These challenges, such as the mode transition from supersonic to hypersonic engine operation, are under study in NASA's Fundamental Aeronautics Program. One propulsion concept that is being explored is a magnetohydrodynamic (MHD) energy- bypass generator coupled with an off-the-shelf turbojet/turbofan. It is anticipated that this engine will be capable of operation from takeoff to Mach 7 in a single flowpath without mode transition. The MHD energy bypass consists of an MHD generator placed directly upstream of the engine, and converts a portion of the enthalpy of the inlet flow through the engine into electrical current. This reduction in flow enthalpy corresponds to a reduced Mach number at the turbojet inlet so that the engine stays within its design constraints. Furthermore, the generated electrical current may then be used to power aircraft systems or an MHD accelerator positioned downstream of the turbojet. The MHD accelerator operates in reverse of the MHD generator, re-accelerating the exhaust flow from the engine by converting electrical current back into flow enthalpy to increase thrust. Though there has been considerable research into the use of MHD generators to produce electricity for industrial power plants, interest in the technology for flight-weight aerospace applications has developed only recently. In this research, electromagnetic fields coupled with weakly ionzed gases to slow hypersonic airflow were investigated within the confines of an MHD energy-bypass system with the goal of showing that it is possible for an air-breathing engine to transition from takeoff to Mach 7 without carrying a rocket propulsion system along with it. The MHD energy-bypass system was modeled for use on a supersonic turbojet engine. The model included all components envisioned for an MHD energy-bypass system; two preionizers, an MHD generator, and an MHD accelerator. A thermodynamic cycle analysis of the hypothesized MHD energy-bypass system on an existing supersonic turbojet engine was completed. In addition, a detailed thermodynamic, plasmadynamic, and electromagnetic analysis was combined to offer a single, comprehensive model to describe more fully the proper plasma flows and magnetic fields required for successful operation of the MHD energy bypass system. The unique contribution of this research involved modeling the current density, temperature, velocity, pressure, electric field, Hall parameter, and electrical power throughout an annular MHD generator and an annular MHD accelerator taking into account an external magnetic field within a moving flow field, collisions of electrons with neutral particles in an ionized flow field, and collisions of ions with neutral particles in an ionized flow field (ion slip). In previous research, the ion slip term has not been considered. Detailed thermodynamic cycle analysis of an annular MHD generator and an annular MHD accelerator revealed that including the ion slip term to the generalized Ohm's Law decreased the needed magnetic fields and conductivity levels as compared to previous research. For the MHD generator, the needed magnetic fields decreased from 5 T to 3 T for all flight speeds studied (Mach 7, 5, and 3). The conductivity levels required for the ionized airflow within the MHD generator at 3 T decreased from 11 mhos/m to 9 mhos/m for a flight speed of Mach 7 and remained the same for Mach 5 and 3. For the MHD accelerator, the needed magnetic fields decreased from 5 T to 3 T for flight speeds of Mach 7 and 5, and decreased from 3 T to 1.5 T for a flight speed of Mach 3. The conductivity levels required for the ionized airflow within the MHD accelerator (at 3 T) decreased from 2.6 mhos/m to 1.1 mhos/m for a flight speed of Mach 7 and remained the same for Mach 5 and 3. The MHD energy-bypass system model showed that it is possible to expand the operating range of a supersonic jet engine from a maximum of Mach 3.5 to a maximum of Mach 7. The inclusion of ion slip within the analysis further showed that it is possible to 'drive&apos; this system with maximum magnetic fields of 3 T and with maximum conductivity levels of 11 mhos/m. These operating parameters better the previous findings of 5 T and 10 mhos/m, and reveal that taking into account collisions between ions and neutral particles within a weakly ionized flow provides a more realistic model with added benefits of lower magnetic fields and conductivity levels especially at the higher Mach numbers.

Committee:

David Allendar, PhD (Committee Co-Chair); Isaiah Blankson, PhD (Committee Co-Chair); John Portman, PhD (Committee Member); Mark Manley, PhD (Committee Member); John West, PhD (Committee Member); Jonathan Maletic, PhD (Committee Member)

Subjects:

Aerospace Engineering; Electromagnetics; Theoretical Physics

Keywords:

magnetohydrodynamics; energy bypass; ion slip; supersonic jet propulsion; hypersonic flight; air-breathing jet engine; weakly ionized gases; plasma flows; thermodynamic cycle analysis

Nieberding, Zachary JAn Investigation of Acoustic Wave Propagation in Mach 2 Flow
MS, University of Cincinnati, 2014, Engineering and Applied Science: Aerospace Engineering
Hypersonic technology is the next advancement to enter the aerospace community; it is defined as the study of flight at speeds Mach 5 and higher where intense aerodynamic heating is prevalent. Hypersonic flight is achieved through use of scramjet engines, which intake air and compress it by means of shock waves and geometry design. The airflow is then directed through an isolator where it is further compressed, it is then delivered to the combustor at supersonic speeds. The combusted airflow and fuel mixture is then accelerated through a nozzle to achieve the hypersonic speeds. Unfortunately, scramjet engines can experience a phenomenon known as an inlet unstart, where the combustor produces pressures large enough to force the incoming airflow out of the inlet of the engine, resulting in a loss of acceleration and power. There have been several government-funded programs that look to prove the concept of the scramjet engine and also tackle this inlet unstart issue. The research conducted in this thesis is a fundamental approach towards controlling the unstart problem: it looks at the basic concept of sending a signal upstream through the boundary layer of a supersonic flow and being able to detect a characterizeable signal. Since conditions within and near the combustor are very harsh, hardware is unable to be installed in that area, so this testing will determine if a signal can be sent and if so, how far upstream can the signal be detected. This experimental approach utilizes several acoustic and mass injection sources to be evaluated over three test series in a Mach 2 continuous flow wind tunnel that will determine the success of the objective. The test series vary in that the conditions of the flow and the test objectives change. The research shows that a characterizeable signal can be transmitted upstream roughly 12 inches through the subsonic boundary layer of a supersonic cross flow. It is also shown that the signal attenuates as the distance between the source and sensors increases. Individual studies including detection sensor and source comparison, material selection, transfer rates, and shadowgraph imagery are also investigated. The acoustic signal is affected by the boundary layer, which is impacted by the shock train and its location. With the capability to characterize an acoustic signal within a scramjet engine to detect the shock train location, any disturbance in the acoustic signals can be linked to shock train displacement that could lead to an inlet unstart. With these results in mind, it is possible that acoustic hardware can be designed to be implemented into the scramjet engine to detect an inlet unstart before it should happen.

Committee:

Ephraim Gutmark, Ph.D. D.Sc. (Committee Chair); Jeffrey M Donbar, Ph.D. (Committee Member); Michael S. Brown, Ph.D. (Committee Member); Paul Orkwis, Ph.D. (Committee Member)

Subjects:

Aerospace Materials

Keywords:

hypersonics;supersonic;scramjet;experimental;signal propagation;inlet unstart

Olson, Lawrence ElroyThe influences of artificially induced turbulence upon boundary-layer transition in supersonic flows /
Doctor of Philosophy, The Ohio State University, 1970, Graduate School

Committee:

Not Provided (Other)

Subjects:

Engineering

Keywords:

Supersonic wind tunnels;Boundary layer

Komar, James JosephInvestigation of fluid dynamic interactions within multiple nozzle arrays /
Doctor of Philosophy, The Ohio State University, 1975, Graduate School

Committee:

Not Provided (Other)

Subjects:

Engineering

Keywords:

Supersonic nozzles

Sevier, AbigailFeasibility Study for Testing the Dynamic Stability of Blunt Bodies with a Magnetic Suspension System in a Supersonic Wind Tunnel
Master of Sciences, Case Western Reserve University, 2017, EMC - Aerospace Engineering
The feasibility of a magnetic suspension and balance system (MSBS) for testing dynamic stability of atmospheric entry capsules in the NASA GRC 225 cm2 Supersonic Wind Tunnel is investigated. The following are examined in the study: largest model size for testing in the MSBS, minimum proximity between wall and model, and analysis techniques using high-speed video images of model movement. Results indicate larger models can be tested in an axisymmetric test section and at locations closer to the nozzle exit resulting from lower boundary-layer blockage. Additionally, models contact the reflected shock from the boundary-layer at a 2.5 cm distance from centerline. Video analysis methods establish a measurement error of 0.6 degrees in pitch or yaw angle. Using these methods, a proof of concept demonstration for a one degree-of-freedom test in pitch simulating blunt body dynamic behavior is compared to ballistic range data for atmospheric entry vehicles.

Committee:

Paul Barnhart, Ph.D. (Advisor); Joseph Prahl, Ph.D. (Committee Member); James T'ien, Ph.D. (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

Magnetic Suspension and Balance System; Blunt Body Dynamic Stability; Supersonic Wind Tunnel; Wind Tunnel Testing; Entry Descent and Landing

Etheridge, Steven J.Effect of Flow Distortion on Fuel Mixing and Combustion in an Upstream-Fueled Cavity Flameholder for a Supersonic Combustor
MS, University of Cincinnati, 2012, Engineering and Applied Science: Aerospace Engineering
Typical studies of scramjet combustion employ as uniform a flowpath as possible. These studies are important to isolate the effects of a given combustor configuration. However, such studies tend to ignore the effects of a shock train created by the vehicle installation and that this shock train changes over the flight envelope. Consequently, the performance of a given configuration is measured without considering the considerable effects of this shock train or how it changes with different flight conditions. This thesis includes experimental and computational studies of the effects of an incident shockwave on the flowfield, fuel distribution and combustion within a cavity flameholder with upstream fuel injection. The effect of the shockwave location (on the upstream fuel jet or over the cavity) and shock angle are controlled by adjusting a shock generator mounted in the tunnel test section. The effect of fuel injection momentum ratio is also examined. Shadowgraphy is used to characterize the flowfield while planar laser induced fluorescence of the NO and OH molecules are used to measure the fuel mixing and combustion, respectively. These experimental data are compared with CFD solutions of the Reynolds Averaged Navier-Stokes equations provided in previous CFD work. The effect of the shock on the cavity shear layer is found to control the fuel distribution within the cavity. The shock on jet impingement forces the shear layer deep within the cavity and results in higher concentrations near the cavity centerline, but low mixing uniformity. The shock on cavity case causes the shear layer to separate upstream of the cavity, mixing uniformity is enhanced by the increased breakup of the fuel plume. Combustion is stronger and more uniform in the shock on cavity case, while it is limited to the edges of the cavity with shock impingement on the jet. The greater mixing afforded in the shock on cavity case reduces the fuel concentration near the centerline and permits stronger burning in the center of the cavity. Small changes in the fuel injection momentum ratio (doubling) do not strongly affect the pattern of fuel distribution in any case. Combustion in the shock on cavity case is reduced by increasing fuel injection momentum because the fuel concentration at the centerline is too high. Small increases in the shock angle did not strongly affect the results.

Committee:

Jongguen Lee, PhD (Committee Chair); San-Mou Jeng, PhD (Committee Member); Paul Orkwis, PhD (Committee Member)

Subjects:

Aerospace Materials

Keywords:

supersonic combustion;shock generator;cavity flameholder;scramjet;

Campos Ramos, Ricardo E.STRUCTURE AND EXCITED-STATE DYNAMICS OF AROMATIC NITRILES IN SUPERSONIC FREE JET
Doctor of Philosophy, University of Akron, 2005, Chemistry
Excitation-energy dependence of fluorescence intensity and lifetime has been measured for 4-dimethylaminobenzonitrile (DMABN), 4-aminobenzonitrile (ABN), 4-diisopropylaminobenzonitrile (DIABN) and 1-cyanonaphthalene (1CNN) in a supersonic free jet. In all cases, the fluorescence yield decreases rather dramatically, whereas the fluorescence lifetime decreases only moderately for ( ) excess vibrational energy exceeding about 1000cm-1. This is confirmed by the normalized fluorescence excitation spectrum with the absorption spectrum of the compound in the vapor phase. The result indicates that the strong decrease in the relative fluorescence yield at higher energies is due mostly to a decrease in the radiative decay rate of the emitting state. Comparison of the experimental results with the Time Dependent Density Functional Theory (DDFT) potential energy curves for excited states strongly suggests that the decrease in the radiative decay rate of the amino-benzonitriles at higher energies is due to the crossing of the singlet state by the lower-lying singlet state of very small radiative decay rate. The threshold energy for the fluorescence “break-off” is in good agreement with the computed energy barrier for the crossing. For 1CNN, on the other hand, the observed fluorescence break-off can be best attributed to the crossing of the singlet state by the triplet state. A concerted experimental (mass-selective spectroscopy) and theoretical (correlated quantum chemistry calculation) study of hydrogen-bonded clusters of 1-cyanonaphthalene (1CNN) with water has been carried out to probe geometries of the conformational isomers. The structures of the two low-energy conformers of 1CNN(H2O) and 1CNN(H2O)2 predicted by MP2/cc-pVDZ calculation, are consistent with ionization-loss (ion-dip) infrared spectra of C-H and O-H stretches of the two conformers, identified by ionization-detected hole-burning spectroscopy. The facile loss of a neutral water molecule from the cluster ion of 1CNN(H2O)2, relative to that of 1CNN(H2O), is in accord with the proposed structures of the complexes. Mass selected resonant two-photon ionization (R2PI) and ion-dip infrared spectroscopies are combined with correlated (MP2) quantum chemistry to probe electronic spectra and ground-state of jet cooled dimer and higher clusters of 1-cyanonaphthalene. The results indicate that the dimer has stacked geometries, consistent with the highly efficient excimer formation that follows photoexcitation of the ground-state clusters.

Committee:

Edward Lim (Advisor)

Subjects:

Chemistry, Physical

Keywords:

STRUCTURE AND DYNAMICS OF AROMATIC NITRILES; SUPERSONIC JET EXPANSION; LASER SPECTROSCOPY LIF MASS HIGH RESOLUTION SPECTROSCOPY

Speth, Rachelle LeaParametric Study of the Effects of the Flapping Mode Excitation on the Near Field Structures of a Mach 1.3 Cold Jet
Master of Science, The Ohio State University, 2012, Aero/Astro Engineering
This effort investigates numerical and physical parameters influencing an ideally-expanded Mach 1.3 jet excited by the m=+/-1 flapping mode. The excitation is imposed by eight Localized Arc Filament Plasma Actuators (LAFPA) placed around the periphery of the circular nozzle exit. The devices are modeled with a proven surface heating approach. The reference case considers the most amplified (jet column mode) frequency corresponding to a Strouhal number of 0.3, based on the diameter of the nozzle and the jet velocity, with an actuator-imposed temperature of 1500K and a duty cycle of 20%. Relative to this reference, the effects of changing frequency, duty cycle and actuator model temperature are explored. In some cases, e.g., actuator temperature, experimental data is not available, but for frequency, there is. The results are analyzed with several different quantitative and qualitative metrics, including time-averaged centerline decay and jet half width as well as phase-averaged coherent structures. Raising the frequency affects the dynamics in several ways. The number of vortical features observed in the phase-averaged data increases and the rate of decay of the centerline velocity is reduced. Furthermore, the alternating vortex ring interactions observed in the reference case are not distinct but are rather replaced by smaller structures, trends which are also observed in experiment. The flow mixes the fastest around the jet column mode (St~0.22). The higher duty cycles exhibits strengthened coherent structures and slightly higher jet growth along the flapping plane, but the overall dynamics remain the same. The response of the jet is relatively insensitive to actuator temperature model within reasonable limits. The latter two studies, with different duty cycles and actuator temperatures, are consistent with previous analyses demonstrating that instability manipulation, rather than heat deposition is the primary mechanism of control.

Committee:

Datta Gaitonde, PhD (Advisor); Mo Samimy, PhD (Committee Member); Mei Zhuang, PhD (Committee Member)

Subjects:

Aerospace Engineering; Engineering; Mechanical Engineering

Keywords:

supersonic jets; Large Eddy Simulation; plasma actuators

Webb, Nathan JosephControl of Supersonic Mixed-Compression Inlets Using Localized Arc Filament Plasma Actuators
Master of Science, The Ohio State University, 2010, Mechanical Engineering
Shock wave/boundary layer interactions (SWBLIs) occur in many supersonic internal flow applications, specifically in mixed compression inlets, as well as in external flows. In this study a nominally Mach 2 mixed compression inlet is modeled by two experimental setups: 1) A compression ramp-generated impinging SWBLI, and 2) a variable angle wedge (VAW) generated impinging SWBLI. The compression ramp and the wedge both serve to generate an oblique shock wave that impinges on the boundary layer on the opposite surface of the wind tunnel. This produces an impinging SWBLI within the test section that replicates the flow found in a mixed compression inlet. A SWBLI can cause flow separation and it is desirable to efficiently prevent this to avoid the many adverse consequences that may result otherwise. The goal of this study is to investigate the ability of localized arc-filament plasma actuators (LAFPAs) to effectively control the interaction. The LAFPAs show significant ability to beneficially affect the SWBLI depending on various operating parameters such as geometry and forcing Strouhal number. This ability apparently stems from a manipulation of instabilities naturally present in the flow. For the compression ramp facility the LAFPAs were most effective when located upstream of the shock foot, forcing with a Strouhal number of 0.03, and operated in-phase. The VAW facility is currently being debugged and will be used for future detailed experiments investigating the control authority of the LAFPAs.

Committee:

Mohammad Samimy, PhD (Advisor); Michael Dunn, PhD (Committee Member); Jeffrey Bons, PhD (Committee Member)

Subjects:

Engineering; Mechanical Engineering

Keywords:

flow control; LAFPA; plasma actuator; SWBLI; supersonic inlet

Soliman, Salah M.Micro-Particles and Gas Dynamics in an Axi-Symmetric Supersonic Nozzle
PhD, University of Cincinnati, 2011, Engineering and Applied Science: Aerospace Engineering
A new biolistic gene gun for micromolecular drug delivery to human skin has been developed and numerically tested. The device generates supersonic flow to accelerate the mediated microparticles to sufficient speeds to breach the outer human skin layer. The device is referred to as CDN-WPI and consists of; a high pressure gas tank, a convergent-divergent nozzle (C-D), and a constant area mixing duct. The mediated microparticles are entrained from a parallel inlet after the exit of the C-D nozzle. The gas from the high pressure tank accelerates through the C-D nozzle to supersonic speeds which in turn accelerate the powder microparticles through the constant area mixing duct to high speeds. A validated numerical procedure is used to study the two-phase flow dynamics inside the biolistic gun using different geometrical configurations. Different driver gas pressures, gas type (helium and air), adding gas swirl, microparticles types and size are considered in this study. The dimensions of the device C-D nozzle, mixing duct length, and the number of particles inlets are the geometrical configurations studied. It is found that using the CDN-WPI device requires reduced driver gas pressure by 50 % compared to the existing devices. The reduction in the gas driver pressure is a result of the elimination of the losses due to boundary layer separation found in all previous devices. The entrainment of the solid microparticles and gas from the parallel inlets precludes flow separation by energizing the boundary layer over the constant area duct walls. As a result, the CDN-WPI is more efficient and safer to use. Further validation is done using semi-empirical particle penetration calculations and the computed flow field which compare very well with the available experimental data. The axi-symmetric model has been used exclusively in all previous numerical solutions of biolistic guns. To check the validity of this assumption, the axi-symmetric results are compared with the results of 3-D model solutions with continuous particles inlet along periphery. The results compare very well which justify the axi-symmetric assumption solutions. We investigated the 3-D two-phase flow field with one, two, and four particle inlets. The 3-D simulations show that for practical and efficient gene gun device, more than one particle inlet is required.

Committee:

Shaaban Abdallah, PhD (Committee Chair); Kelly Cohen, PhD (Committee Member); Milind Jog, PhD (Committee Member); Mark Turner, ScD (Committee Member)

Subjects:

Aerospace Materials

Keywords:

CFD;Supersonic nozzle;3-D numerical study;Biolistic Gene gun;Particles-gas flows

Munday, DavidFlow and Acoustics of Jets from Practical Nozzles for High-Performance Military Aircraft
PhD, University of Cincinnati, 2010, Engineering and Applied Science: Aerospace Engineering

This research project examines supersonic jets from nozzles representative of the practical variable-geometry convergent-divergent nozzles used on high-performance military aircraft. The nozzles employed have conical convergent sections, sharp throats and conical divergent sections. Nozzles with design Mach numbers of 1.3, 1.5, 1.56 and 1.65 are tested and the flow and acoustics examined. Such nozzles are found to produce a double-diamond shock structure consisting of two overlapping sets of shock cells, one cast from the nozzle lip and one cast from the nozzle throat. These nozzles are found to produce no shock-free condition at or near the design condition. As a result they produce shock-associated noise at all supersonic conditions. The shock cell spacing, broad-band shock-associated noise peak frequency and screech frequency all match those of more traditional nearly isentropic convergent-divergent nozzles.

A correlation is proposed which improves upon the Prandtl-Pack relation for shock cell spacing in that it accounts for differences in nozzle design Mach number which the Prandtl-Pack relation does not. This proposed relation reverts to the Prandtl-Pack equation for the case of a design Mach number of 1.0.

Chevrons are applied to the nozzles with design Mach numbers of 1.5 and 1.56. The effective penetration of the chevrons is found to be a function of the jet Mach number. Increasing jet Mach number increases effective penetration of the chevrons and increases the magnitude of all chevron effects. Chevrons on supersonic jets are found to reduce shock cell length, increase mixing and spreading, decrease turbulent kinetic energy at the end of the potential core and increase it near the nozzle. Chevrons corrugate the shear layer but not the shock structures inside the jet which remain axisymmetric. Chevrons thicken the shear layer, reducing the sonic diameter and reducing the diameter of the shock cells. By reducing their diameter they also reduce the shock cell spacing. Chevrons reduce low-frequency mixing noise near the end of the potential core, increase high-frequency noise near the nozzle exit. They eliminate screech and reduce broad-band shock-associated noise and shift it to higher frequencies.

Fluidic injection is applied to the nozzle with design Mach number of 1.5. Fluidic injection corrugates the shear layer, increases mixing and spreading, reduces low frequency mixing noise, increases high frequency noise, reduces broad-band shock-associated noise and shifts its peak to higher frequency.

Committee:

Ephraim Gutmark, PhD, DSc (Committee Chair); Shaaban Abdallah, PhD (Committee Member); Paul Orkwis, PhD (Committee Member); James Bridges, PhD (Committee Member); Kailas Kailasanmath, PhD (Committee Member)

Subjects:

Aerospace Materials

Keywords:

jet noise;chevrons;microjets;fluidics;Prantdl-Pack;supersonic jets

Bhabhe, Ashutosh ShrikantExperimental Study of Condensation and Freezing in a Supersonic Nozzle
Doctor of Philosophy, The Ohio State University, 2012, Chemical and Biomolecular Engineering

Phase transitions like condensation, vaporization, melting and freezing are ubiquitous and important in a host of applications across science and technology. A phase transition is initiated by nucleation, the process where fragments of a new phase begin to form in a supersaturated mother phase. This is followed by the growth of these fragments after they reach a critical size, and, finally, the new stable phase is formed by ageing. Understanding phase transitions, especially the nucleation step, has been an active area of research for over a century with approaches ranging from the development of theory, the refinement of experiments and, more recently, direct in silico simulations. The lack of a robust unified theory that can quantify the process of nucleation for all substances, across a broad range of pressures and temperatures, underscores the challenges involved.

Experimental studies on vapor-liquid nucleation have primarily focused on substances like water, straight chain alcohols and long chain alkanes. Given the complexity of these substances, the uncertainties in key physical properties and a lack of knowledge regarding the intermolecular potentials, meaningful comparisons between theory, experiments and simulations is challenging. A primary goal of this research is, therefore, to experimentally investigate the nucleation behavior of “simple” molecules including argon and nitrogen. Following on the work of Sinha (Dissertation, Ohio State University, 2008), the current work extends the range of experimental argon condensation data in a cryogenic supersonic nozzle (SSN) to lower temperatures and higher supersaturations. Based on the experimental measurements and estimated nucleation rates of 1017±1 cm-3s-1 in our SSN, Classical Nucleation Theory (CNT) predicts nucleation rates that are lower by 11-13 orders of magnitude. In contrast, rates predicted by Mean Field Kinetic Nucleation Theory (MKNT), a recent theory developed within the construct of statistical mechanics, are within 1-2 orders of magnitude of the experimental estimates. The experimental approach used for argon is successfully extended to study the condensation of nitrogen, a challenge given that expansions start at ~85 K, about 15 - 20 K lower than the argon experiments. For nitrogen, MKNT rate predictions again agree better with experiments, by ~11 orders of magnitude, than the predictions of CNT. The possible reasons for the success of MKNT are explored by combining the high rate data measured here with the lower rate data measured by Iland et al. (J. Chem. Phys. 127, 154506, 2007; J. Chem Phys., 130, 114508, 2009) in order to estimate the critical cluster properties, including the cluster size, excess internal energy, and excess entropy, for both argon and nitrogen. Although both CNT and MKNT over-predict the critical cluster size, the predictions of the excess internal energy and entropy predicted by MKNT are in significantly better agreement with experimental estimates of these quantities. In this work, the freezing of supercooled heavy water (D2O) droplets in a supersonic nozzle is studied by applying three in situ position resolved experimental techniques including static pressure measurements, small angle X-ray scattering and Fourier transform infra-red spectroscopy. Combining the information from these three techniques yields the size, phase, composition and number density of the droplets, as well as the flow variables including temperature, density and velocity. Freezing of supercooled D2O droplets in the size range of 3 to 9 nm occurs at droplet temperatures ranging from 222 K to 226 K while the corresponding ice nucleation rates are ~1023 cm-3s-1 assuming the phase transition can occur throughout the volume of the droplet and ~1016 cm-3s-1 assuming nucleation is initiated at the surface. The current D2O ice nucleation data are more consistent with nucleation initiated at the surface of the droplet, but uncertainty in the experiments makes it difficult to definitively state which process dominates. Some of the difficultly may simply be that for the smallest drops the outer 1 nm of the droplet constitutes ~70% of the volume. Using current estimates of the thermophysical properties of the condensed phases of D2O, the theoretical ice nucleation rates reproduce the experimental data trends qualitatively but not quantitatively.

Finally, given the importance of understanding the kinetics of formation of clathrate hydrates, a “proof of concept” study shows that these complex structures form on a microsecond time scale in a supersonic nozzle, when tetrahydrofuran (THF) is the guest molecule.

Committee:

Barbara Wyslouzil, PhD (Advisor); Isamu Kusaka, PhD (Committee Member); James Rathman, PhD (Committee Member); Brian Focht (Other)

Subjects:

Chemical Engineering; Chemistry; Condensation

Keywords:

Nucleation; Condensation; Freezing; Supersonic nozzle; Argon; Nitrogen; Heavy water; Clathrate Hydrates

Liang, Fang-PeiAn inviscid stability analysis of unbounded supersonic mixing layer flows
Doctor of Philosophy, Case Western Reserve University, 1991, Mechanical Engineering
An analysis is performed of the inviscid linear stability of unbounded supersonic mixing layer flows. The general stability characteristics of subsonic and supersonic solutions and the unstable eigen spectrum are discussed. The existence of neutral solutions supersonic to either or both free streams is verified. Two unstable supersonic modes, which can be supersonic to either or both of the free streams, are found and the transition from the supersonic modes to that associated with generalized inflection point is investigated. Systematical study is made of the unstable eigen spectrum using a number of numerical methods, including the method of the principle of argument and the method of multi-domain Chebyshev collocation with the QZ algorithm. The method of the principle of argument is robust and theoretically neat. The Chebyshev collocation method, however, is very expensive when applied to supersonic stability computation because of the presence of the critical point singularity and the slow damping of the eigenfunction at free streams. It is concluded that for unbounded flows there are only a small number of supersonic unstable modes (two modes in our cases) and the exact number for a given supersonic flow is generally not known. To determine the number of unstable modes requires detailed investigation by, e.g. the method of the principle of argument.

Committee:

Eli Reshotko (Advisor)

Keywords:

inviscid unbounded supersonic layer flows

Crowell, Andrew RModel Reduction of Computational Aerothermodynamics for Multi-Discipline Analysis in High Speed Flows
Doctor of Philosophy, The Ohio State University, 2013, Aero/Astro Engineering
This dissertation describes model reduction techniques for the computation of aerodynamic heat flux and pressure loads for multi-disciplinary analysis of hypersonic vehicles. NASA and the Department of Defense have expressed renewed interest in the development of responsive, reusable hypersonic cruise vehicles capable of sustained high-speed flight and access to space. However, an extensive set of technical challenges have obstructed the development of such vehicles. These technical challenges are partially due to both the inability to accurately test scaled vehicles in wind tunnels and to the time intensive nature of high-fidelity computational modeling, particularly for the fluid using Computational Fluid Dynamics (CFD). The aim of this dissertation is to develop efficient and accurate models for the aerodynamic heat flux and pressure loads to replace the need for computationally expensive, high-fidelity CFD during coupled analysis. Furthermore, aerodynamic heating and pressure loads are systematically evaluated for a number of different operating conditions, including: simple two-dimensional flow over flat surfaces up to three-dimensional flows over deformed surfaces with shock-shock interaction and shock-boundary layer interaction. An additional focus of this dissertation is on the implementation and computation of results using the developed aerodynamic heating and pressure models in complex fluid-thermal-structural simulations. Model reduction is achieved using a two-pronged approach. One prong focuses on developing analytical corrections to isothermal, steady-state CFD flow solutions in order to capture flow effects associated with transient spatially-varying surface temperatures and surface pressures (e.g., surface deformation, surface vibration, shock impingements, etc.). The second prong is focused on minimizing the computational expense of computing the steady-state CFD solutions by developing an efficient surrogate CFD model. The developed two-pronged approach is found to exhibit balanced performance in terms of accuracy and computational expense, relative to several existing approaches. This approach enables CFD-based loads to be implemented into long duration fluid-thermal-structural simulations.

Committee:

Jack McNamara (Advisor); Thomas Eason, III (Committee Member); Jeffrey Bons (Committee Member); Mo-How Herman Shen (Committee Member); Mei Zhuang (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

Aerodynamic Heating; Pressure; Fluid-Thermal-Structural Analysis; Reduced Order Modeling; Surrogate Modeling; Computational Fluid Dynamics; Hypersonic; Supersonic; High-Speed Flow

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