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A new noise metric has been developed that may be used for optimization problems involving aerodynamic noise from a clean wing. The modeling approach uses a classical trailing edge noise theory as the starting point. The final form of the noise metric includes characteristic velocity and length scales that are obtained from three-dimensional, steady, RANS simulations with a two- equation k-omega turbulence model. The noise metric is not the absolute value of the noise intensity, but an accurate relative noise measure as shown in the validation studies. One of the unique features of the new noise metric is the modeling of the length scale, which is directly related to the turbulent structure of the flow at the trailing edge. The proposed noise metric model has been formulated so that it can capture the effect of different design variables on the clean wing airframe noise such as the aircraft speed, lift coefficient, and wing geometry. It can also capture three-dimensional effects which become important at high lift coefficients, since the characteristic velocity and the length scales are allowed to vary along the span of the wing. Noise metric validation was performed with seven test cases that were selected from a two-dimensional NACA 0012 experimental database. The agreement between the experiment and the predictions obtained with the new noise metric was very good at various speeds, angles of attack, and Reynolds Number, which showed that the noise metric is capable of capturing the variations in the trailing edge noise as a relative noise measure when different flow conditions and parameters are changed. Parametric studies were performed to investigate the effect of different design variables on the noise metric. Two-dimensional parametric studies were done using two symmetric NACA four-digit airfoils (NACA 0012 and NACA 0009) and two supercritical (SC(2)-0710 and SC(2)-0714) airfoils. The three-dimensional studies were performed with two versions of a conventional transport wing at realistic approach conditions. The twist distribution of the baseline wing was changed to obtain a modified wing which was used to investigate the effect of the twist on the trailing edge noise. An example study with NACA 0012 and NACA 0009 airfoils demonstrated a reduction in the trailing edge noise by decreasing the thickness ratio and the lift coefficient, while increasing the chord length to keep the same lift at a constant speed. Both two- and three-dimensional studies demonstrated that the trailing edge noise remains almost constant at low lift coefficients and gets larger at higher lift values. The increase in the noise metric can be dramatic when there is separation on the wing. Three-dimensional effects observed in the wing cases indicate the importance of calculating the noise metric with a characteristic velocity and length scale that vary along the span. The twist change does not have a significant effect on the noise at low lift coefficients, however it may give significant noise reduction at higher lift values. The results obtained in this study show the importance of the lift coefficient on the airframe noise of a clean wing and favors having a larger wing area to reduce the lift coefficient for minimizing the noise. The results also point to the fact that the noise reduction studies should be performed in a multidisciplinary design and optimization framework, since many of the parameters that change the trailing edge noise also affect the other aircraft design requirements. It's hoped that the noise metric developed here can aid in such multidisciplinary design and optimization studies. / Ph. D.

No / A novel technique for the laboratory characterization of the frequency-dependent acoustic surface admittance of partly saturated samples of sands is presented. The technique is based on a standard laboratory de-watering apparatus coupled with a standard acoustic impedance tube. The dependence of the surface admittance on the degree of water saturation is investigated for two samples of sand with widely different flow resistivities. It is shown that a relatively small change (e.g., from 0% to 11% by volume) in the degree of water saturation can result in a much larger change (e.g., twofold) in the acoustic surface admittance. An empirical relationship is found between the peaks observed in the real part of admittance spectra for the low flow resistivity sand and the degree of water saturation. The data are compared with predictions of two widely used ground impedance models: a semiempirical single parameter model and a two parameter model. A modified two-parameter version of a single-parameter model is found to give comparable fit to the two-parameter model. However, neither model provides an accurate fit.

Nonlinear propagation and shock formation are shown in noise radiated from full-scale military jet aircraft. Perception of sound is not only affected by the overall sound pressure level of the noise, but also characteristics of the sound itself. In the case of jet noise, acoustic shocks within the waveforms result in a characteristic commonly referred to as"crackle." The origin of shocks in the far-field of jet noise is shown to be through nonlinear propagation. Metrics characterizing the shock content of a waveform are explained and given physical significance, then applied to jet noise at various distances and engine conditions to show areas where shock formation is significant. Shocks are shown to develop at different distances from the aircraft, dependent on the amplitude and frequency, and nonlinear propagation is shown to be important in determining time and frequency characteristics of jet noise at distances of up to 1220 m from the aircraft. The shock content is also characterized during flyover experiments, and the shock content between the two scenarios is compared. While some reduction in overall level and shock content is seen in the maximum radiation region, level increases in the forward direction during flight result in increased shock content. Variation at distances of 305 m and beyond is considered and shown as a result of small atmospheric changes. Finally, a nonlinear numerical propagation scheme is used to model the propagation, showing accuracy in predicting frequency-domain and time-domain features that are evidence of nonlinear propagation.

The excitation mechanism of trapped diametral acoustic modes within a rectangular cavity-duct system is investigated both numerically and experimentally. The asymmetry inherent within the rectangular geometry introduces a preferred orientation, ensuring the excited diametral modes remain stationary. Three separate cavities are manufactured and tested. This included two asymmetric rectangular cross-sections and one symmetric square cavity. Experimental results indicate that the aeroacoustic responses of the three cavities are dominated by the strong excitation of trapped diametral modes. Numerical simulations indicate that the resolved radial acoustic particle velocity distributions are non-uniform at the upstream separation edge where the formation of vortical structures is initiated. As the cavity became smaller, and more asymmetric, the trapped nature of the acoustic modes decreased with an accompanied increase in the radiation losses and reduction in pulsation amplitude. Observations of the aeroacoustic measurements show evidence of three unique modal behaviours. The first case is the independent excitation of a single stationary mode where specific circumferential sections of the shear layer were excited and initiating the formation of vortical disturbances. These circumferential sections, and distribution of disturbances, were akin to the excited mode shape. The second case involved simultaneous excitation of two stationary modes. This suggested that the shear layer was exciting two modes simultaneously. Neighbouring circumferential sections, at the initial region of the shear layer, were being excited independently and at different resonant frequencies. Finally, a spinning trapped acoustic mode was observed in the symmetric square cavity. Due to the spinning nature, the excited circumferential portions and formation of vortices were non-uniform and rotated with the spinning acoustic mode. This resulted in the formation of a three-dimensional helical structure. / Thesis / Master of Applied Science (MASc)

In the field of supersonic jet noise, measurements of full-scale afterburning jet engines are infrequent and provide unique opportunities to better understand jet noise phenomena. This thesis represents a phenomenological jet noise analysis using far-field noise data collected from a T-7A-installed GE F404-103 turbofan engine. One issue with the far-field acoustic data from the T-7A was the effects of ground reflections present in the spectra generated from the measured waveforms. A previously developed ground reflection model was implemented into the data to account for this interference. This work represents the first time this model has been implemented in data collected from a full-scale aircraft. Spectra and spatiospectral maps are used to show that, while imperfect, this model represents a step in the right direction for accounting for ground reflections. From the ground reflection-corrected data, sound power values were calculated at varying engine powers. These values were compared against mechanical power values calculated using provided engine parameters at the corresponding engine conditions. It is shown that the observed increase in sound power with mechanical power at supersonic engine conditions follows classical jet noise theory, while the increase between transonic engine conditions is much greater than predicted by classical jet noise theory. This divergence is currently unexplained. Finally, far-field noise directivity measured from the T-7A is connected back to both physics-based and empirically derived definitions of the convective Mach number, a dimensionless parameter used to describe the velocities of coherent structures in the turbulent mixing layer of a jet. For supersonic jets, where Mach wave radiation is the dominant noise source, the convective Mach number should be useful for predicting peak directivity angles. The evaluated definitions show that the convective Mach number associated with Kelvin-Helmholtz instability waves best predicts the peak directivity of the T-7A.

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