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This thesis considers the sound generated by unsteady perturbations interacting with solid aerofoils in background steady flows, in an attempt to further develop analytic models for the noise generated by blades within turboengines. Specifically, high-frequency unsteady gust and sound wave perturbations are considered and asymptotic results are obtained for, primarily, the far-field noise. Previous analytic work has examined high-frequency gust-aerofoil interactions in steady uniform flows using rapid distortion theory, and has focused on aerofoils with simple geometries. We extend this to deal with aerofoils with more realistic geometries (by including camber, thickness, and angle of attack), as well as considering the new topic of sound-aerofoil interactions in steady uniform flows for aerofoils with realistic geometries. The assumption of a steady uniform flow is later relaxed and we investigate the sound generated by high-frequency gust-aerofoil interactions in steady shear flows. Throughout all of the aforementioned work, the key process involves identifying various asymptotic regions around the aerofoil where different sources dominate the generation of sound. Solutions are obtained in each region and matched using the asymptotic matching rule. The dominant regions producing noise are the local, “inner”, regions at the leading and trailing edges of the aerofoil. Approximations for the far-field noise (in the “outer” regions) are the principal results, however one can also extract approximations for the unsteady pressure generated on the surface of the aerofoil. The surface pressure generated by high-frequency gust-aerofoil interaction in uniform flow is found to contain a singularity at the leading-edge stagnation point, thus the final piece of work in this thesis focuses more closely on turbulent interactions with solid body stagnation points in uniform flow, eliminating this singularity.

Acoustic experiments were conducted in an anechoic chamber with a 1/14th scale model of a supersonic aircraft engine inlet using Trailing Edge Blowing (TEB) to reduce the engine fan noise from a turbofan propulsion simulator (TPS). The TPS is 4.1 in. (10.4 cm) in diameter and is powered by compressed air. The supersonic inlet is connected to the TPS and is geometrically and acoustically scaled from a working design. The supersonic inlet is operated in a takeoff or landing operating condition where the inlet core flow is subsonic. TEB is the process of ejecting high pressure air to re-energize the wakes of upstream fan disturbances such as struts or inlet guide vanes (IGV). The elimination of the wakes will provide a uniform flow field at the engine fan face and reduce noise at the blade passing frequency. The TEB was implemented on six non-uniformly spaced support struts in the inlet. Acoustic tests were then performed at 40%, 60% and 88% of the fan design speed (PNC) to measure the reduction in the blade passing tone (BPT) due to TEB from the struts with and without the presence of IGV. The noise reductions without IGV at 40 PNC show the best results with the blade passing tone (BPT) being reduced by an average of 3.1 dB. The first harmonic of the BPT and the overall Sound Pressure Level (SPL) were also reduced by 1 dB. The addition of the IGV in the inlet reduced the effectiveness of the TEB. The addition of IGV changed the reduction in BPT at 40 PNC by 0.5 dB and the overall SPL was unchanged. At 60 PNC the addition of IGV reduced the reduction due to TEB in the BPT from an average of 2 dB to an average of 1 dB. The tests performed at 88 PNC showed negligible effects due to TEB. Aerodynamic experiments performed on the inlet that showed that the wakes of the IGV have a larger velocity defect than the struts, thus making the IGV a greater noise source. / Master of Science

Impinging high-speed planar jets are susceptible to self-excited aeroacoustic feedback mechanisms due to the coupling of the highly unstable shear layer and upstream travelling pressure waves created by the jet impingement. This aeroacoustic feedback mechanism results in intense narrowband acoustic tones and large amplitude oscillations of the jet column which are undesirable for its use as an actuator for coating weight control in the continuous gas-jet wiping line. This thesis experimentally investigates the use of auxiliary high-speed planar jets for the purpose of interrupting and reducing the amplitude of the negative effects of the aeroacoustic feedback mechanism. Testing was performed using a planar multi-slot nozzle jet over a range of impingement distances, velocities and nozzle widths. The amplitudes of the acoustic tones were found to be a function of the ratio of velocities between the main and auxiliary jets with the tones found to be eliminated at sufficiently high-velocity ratios. Larger auxiliary jet widths were found to further reduce the amplitude of the tones in conjunction with the velocity ratio. The reduction of the amplitude of the tones was accompanied by a reduction in the maximum fluctuating pressure at the plate by 75% and an increase in the maximum static pressure by 30% indicating a reduction in the oscillations of the jet column. A proper orthogonal decomposition of particle image velocimetry vector fields revealed that an increase in the auxiliary jet velocity increased the percentage of the kinetic energy of the mean flow field of the jet but decreased the percentage of the kinetic energy of the modes associated with the aeroacoustic feedback mechanism. The vorticity of the modes associated with the aeroacoustic feedback mechanism shows that the coherent structures inherent to the feedback mechanism reduce in size and strength with increasing auxiliary jet velocity. Time-averaged particle image velocimetry vector fields revealed that at jet conditions where the acoustic tones were reduced, the interaction of the auxiliary jets reduced the maximum vorticity of the shear layer by 35% at the jet exit. Smaller amplitude and thicker shear layers are known to result in smaller maximum growth rates of disturbances in shear layers. The reduced growth rate resulted in smaller coherent structures in the jet shear layer which resulted in the smaller jet column oscillations and the elimination of the acoustic tones. / Thesis / Candidate in Philosophy

Noise is a major environmental pollutant, which can inict physical and psychologicalinjury. An important noise contribution stems from aeroacousticsources, which are found in e.g., ventilation ducts, engine exhaust systems andairplane engines.In this thesis, research methods for low Mach number aeroacoustic sourcesin ducts are developed. The basis of the methods is the ability to describe theintrinsic linear properties of the source as an N-port (multiport), where theoutput sound eld is related to the input sound eld and the generated soundeld, all consisting of plane waves. The methods presented are both numericaland experimental. The numerical method treats the passive properties, i.e.,scattering, attenuation and amplication of incident sound, while the experimentalmethod treats the active part (intrinsic sound generation) as well. Themethods are applied in the study of noise generation by a vortex mixer plate,placed in an airow of Mach 0.2, and in the study of acoustic-hydrodynamicinteraction in a T-junction of rectangular ducts.It is found that the accuracy of the experimental methods is signicantly increasedwhen the equations are over-determined, which is achieved by addingadditional microphones to the test rig. In the frequency range studied, themixer plate is found to generate less sound when made exible, without disturbingthe mixing quality.For the numerical method { based on the linearized Navier-Stokes equations,a model of the oscillation of the Reynolds stress (\turbulent damping")due to the acoustic eld is introduced. By comparing with experimental results,it is found that not using this model results in an over-prediction of theamplication at higher frequencies with several factors in magnitude, whileimplementing the model results in a much better agreement. / <p>QC 20121123</p>

Acoustic resonances leading to high unsteady pressure levels may occur in flow past cavities. The resonance involves a coupling between the downstream-propagating instability wave on the shear layer spanning the open face of the cavity, and acoustic waves propagating within and external to the cavity. These elements of the disturbance field are coupled by the scattering processes that occur at the upstream and downstream ends of the cavity. We develop a theoretical prediction method that combines propagation models in the central region of the cavity with scattering models for the end regions. In our analyses of the scattering processes at the cavity ends, the square-corner geometry is treated exactly, by a method employing the Wiener--Hopf technique. The shear layer is approximated as a vortex sheet in the edge scattering analyses, but finite shear-layer thickness is accounted for in analyzing the propagation of the waves along the length of the cavity. The global analysis leads to a prediction for the resonant frequencies which has a form similar to the Rossiter formula, but contains no empirical constants. In addition to prediction of the frequency, our theory determines the temporal growth or decay rate of each mode. Finally, our theory also predicts the influence of secondary feedback loops involving other components of the unsteady field. Comparisons of the predictions with existing experimental data are made.

The supersonic jet plumes that exhaust from the engines of tactical aircraft produce intense noise signatures that expose the Navy personnel working on the deck of aircraft carriers to dangerously high levels of noise that often results in hearing damage. Reducing the noise radiated by these supersonic plumes is of interest to the Department of Defense and is the primary motivation of this research. Fundamentally, jet noise reduction is achieved by manipulating the nozzle boundary condition to produce changes in the turbulence development and decrease the acoustic efficiency of coherent structures. The research presented here focuses on a novel jet noise reduction technique involving a centered thermal non-uniformity that alters the base flow by introducing a temperature-driven centerline velocity deficit into a perfectly expanded Mach 1.5 jet. The results indicate $2 pm 0.5$ dB reductions in peak narrowband spectral sound pressure levels upstream of peak directivity directions for the non-uniform jet compared to a thermally uniform baseline, even for static thrust matched conditions. This reduction is hypothesized to be related to perturbations induced by the thermal non-uniformity that convect inside the irrotational core and reduce the correlation length scales of turbulence at locations far downstream. This hypothesis was evaluated by studying the coherent turbulence via its convective hydrodynamic footprint in the near-field. An indirect investigation of the near-field using a far-field-informed model of the wavenumber-frequency spectra indicate a reduction in the energy contained in the tail of the wavenumber spectra amplitude, suggesting a reduction in the size of large scale structures. A direct evaluation of the spatio-temporal behavior of the near-field was performed using temporally resolved schlieren images. Space-time correlations of the frequency-filtered near-field identified high frequency acoustic waves radiated by compactly coherent turbulent structures and low frequency Mach waves produced by large scale instabilities. In the thermally non-uniform case these features and their sources were found to be decorrelated at downstream regions. These results provide strong evidence that a centered thermal non-uniformity reduces the radiated noise compared to a uniform baseline by shortening the correlation length scales of coherent structures in regions far from the nozzle exhaust. / Doctor of Philosophy / A more complete understanding of the intense noise sources present in supersonic jet plumes is of value to both government and industry, and is a necessary step towards optimizing noise reduction techniques. Tactical aircraft that operate on the deck of aircraft carriers expose Navy personnel to dangerously high levels of noise that often results in permanent hearing damage. Supersonic jet noise reduction is also of relevance to the recent efforts to revitalize supersonic air transport over land. For supersonic air transport to become a reality, the noise produced by these future aircraft during takeoff and landing must meet the increasingly stringent community noise requirements. Fundamental jet noise research is needed to guide the design of future engine architectures for these aircraft to ensure their commercial success. The research presented herein examines a novel noise reduction technique that involves a centered thermal non-uniformity consisting of a heated jet plume with a spot of locally cooler, slower moving air concentrated along the centerline of a Mach 1.5 jet. This temperature driven velocity deficit is shown to reduce the radiated noise by up to 2.5 dB at peak frequencies and at angles just outside of the peak directivity direction. The cause of the noise reduction is hypothesized be related to a reduction in the size of the coherent structures that radiate a majority of the noise produced by turbulent jets. This hypothesis is evaluated by examining the 'footprint' of the coherent structures in the ambient field directly outside of the jet shear layer in an area called the near-field. An indirect investigation of the near-field using a far-field informed analytic model suggests a reduction in the size of large scale structures. A direct evaluation of the space time structure of the near-field was performed using temporally resolved schlieren images. Statistical processing of the density gradient provided by the schlieren images revealed acoustically intense structures known as Mach waves and high frequency acoustic waves. These features and their sources, large scale instabilities and compactly coherent turbulence, were found to be decorrelated by the introduction of the thermal non-uniformity. These results provide strong evidence that the centered thermal non-uniformity produces a noise benefit by reducing the size of the turbulent structures.

The leading-edge tubercle is a type of airfoil modification that inspired by the humpback whale. It was found that the aerodynamic performance of the wing would increase compared to the wing without tubercles. In the past several years, a lot of numerical and experimental studies have been accomplished to explore this leading-edge modification. Besides the aerodynamic performance change, this research explores the aeroacoustics behavior of airfoils with leading-edge tubercles. A numerical study based on Computational Fluid Dynamics (CFD) is established, and simulations using Star CCM are accomplished based on reasonable set-ups. The airfoil chosen to create the wing is NACA 4412 which is an asymmetric airfoil. Two different tubercle wavelengths were used: 20 mm and 25 mm. The baseline airfoil is the wing that made of the same airfoil but without any modifications. For wings with leading-edge tubercles, the wavelength of the tubercles is the only changing parameter. It was found that the wings with leading-edge tubercles can reduce the noise generation, and the best noise reduction is achieved for a value of 2.525 dB (Decibel) at Point Receiver 10 for the wing that has 25 mm wavelength leading-edge tubercles. However, the wavelength of tubercles does not affect the aeroacoustics performance in an obvious way.