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DNS solutions for a pipe/jet configuration are re-computed with the pipe alone to investigate suppression of previously identified internal noise source(s) with an acoustic liner, using a time domain acoustic liner model developed by Tam and Auriault (AIAA Journal, 34 (1996) 913–917). Liner design parameters are chosen to achieve up to 30 dB attenuation of the broadband pressure field over the pipe length without affecting the velocity field statistics. To understand the effect of the liner on the acoustic and turbulent components of the unsteady wall pressure, an azimuthal/axial Fourier transform is applied and the acoustic and turbulent wavenumber regimes clearly identified. It is found that the spectral component occupying the turbulent wavenumber range is unaffected by the liner whereas the acoustic wavenumber components are strongly attenuated, with individual radial modes being evident as each cuts on with increasing Strouhal number. [ABSTRACT FROM AUTHOR]

The surface pressure spectrum characteristics are studied on a Controlled Diffusion airfoil at varying Mach numbers and a fixed Reynolds number. Compressible large eddy simulation datasets have been used for this analysis. The study focusses on the noise generation mechanism and the source locations responsible for it. A mid-high frequency hump is observed for the higher Mach number cases in the pressure spectrum and the amplitude of this observed hump increases with Mach number. To investigate the cause of this hump, different hydrodynamic and acoustic pressure modes traveling at specific speeds are obtained by a wavenumber–frequency decomposition. The acoustic component originating from the separation bubble region near the leading edge is found to be responsible for the hump in the pressure spectrum. The interaction of hydrodynamic and acoustic waves also plays a significant role in the appearance of the hump in the spectrum. There is also an increase in the amplitude of the high frequency region in the acoustic spectrum as we go towards the trailing edge for higher Mach number which can be attributed to an increase in turbulent fluctuation amplitude in the leading edge region. Furthermore, stability analysis has been performed on one-dimensional mean boundary-layer profiles. The higher reverse velocity in the separation bubble with increasing Mach number seems to be causing higher growth rates of instabilities, ultimately resulting in the significant surface spectrum variations. The directivity in the farfield also confirms a significant contribution of the hump in the spectrum in the overall sound pressure level. [ABSTRACT FROM AUTHOR]

Abstract: In this study the role of acoustic feedback instabilities in the tonal airfoil self-noise phenomenon is investigated. First, direct numerical simulations are conducted of the flow around a NACA-0012 airfoil at and four angles of attack. At the two lowest angles of attack considered the airfoil self-noise exhibits a clear tonal contribution, whereas at the two higher angles of attack the tonal contribution becomes less significant in comparison to the broadband noise. Classical linear stability analysis of time-averaged boundary layer profiles shows that the tonal noise occurs at a frequency significantly lower than that of the most convectively amplified instability wave. Two-dimensional linear stability analysis of the time-averaged flowfield is then performed, illustrating the presence of an acoustic feedback loop involving the airfoil trailing edge. The feedback loop is found to be unstable only for the cases where tonal self-noise is prominent, and is found to self-select a frequency almost identical to that of the tonal self-noise. The constituent mechanisms of the acoustic feedback loop are considered, which appear to explain why the preferred frequency is lower than that of the most convectively amplified instability wave. [Copyright &y& Elsevier]

Three compressible-flow direct numerical simulations (DNS) of aerofoils were conducted at a chord based Reynolds number of R e c = 150 , 000 : a Controlled-Diffusion (CD) aerofoil at two incidences and an untripped NACA 6512-63 aerofoil at 0 ∘ angle of attack (AoA). The aim of the simulations was to help improve the understanding of the generation mechanisms of aerofoil self-noise. To this end, the time-averaged flows around the aerofoils were investigated using a multi-dimensional global stability analysis based on the response to very low-amplitude hydrodynamic initial pulses. For the CD aerofoil at 5 ∘ AoA and the untripped NACA 6512-63 aerfoil at 0 ∘ AoA, the base flows are found to be globally unstable and an acoustic feedback loop occurs regardless of the perturbation locations. For the CD aerofoil at a slightly higher incidence (8 ∘ AoA) where a laminar separation bubble appears at the leading edge and a turbulent attached boundary layer is present at the trailing edge, the initial perturbations decay with time for all initial pulse locations indicating that the base flow is globally stable. The results confirm that for cases where a laminar separation bubble is present on the aerofoil suction side at the trailing edge, independent of whether tones in the DNS exist or not, an unstable feedback loop exists. The laminar separation bubble at the aerofoil trailing edge on the suction side seems to be the only required prerequisite for triggering the acoustic feedback loop. [ABSTRACT FROM AUTHOR]

The effect of the structural compliance of an elastic trailing-edge (TE) on its aeroacoustic response is investigated by means of Direct Numerical Simulation. For that purpose, a two-dimensional semi-infinite rigid plate provided with an elastic TE is considered. Two different test cases of increasing complexity from an aerodynamic perspective are simulated. First, the boundary layer of the plate is forced at a single frequency to produce laminar instabilities that are scattered to the far field as acoustic pressure fluctuations from the TE. Four elastic TEs with different structural properties are simulated with a state-of-the-art high fidelity computational framework and their acoustic characteristics are analyzed. An acoustic decomposition is performed to separate the contribution of the motion-induced noise from the scattering due to the interaction of the incident fluctuations with the TE. We show that the noise reduction offered by the elastic TEs is achieved when the relative phase and amplitude of these two acoustic contribution ensure their mutual cancellation. For the second test case, a cylinder positioned above the upper surface of the plate is used as a vortex generator to produce an incident pressure field that is convected over the TE, resulting in scattered noise from the edge. The same four elastic TEs adopted in the previous case are tested in this new configuration. The results are discussed and compared with the response of an equivalent rigid TE. Insights are given into the mutual interaction between the structural motion of the elastic TE and the incident pressure field by a detailed analysis of the aerodynamic near field of the TE and the cylinder. Furthermore, the individual contributions to the noise radiation of the incident and of the motion-induced surface pressure are separated by performing an additional campaign of simulations in which the cylinder is removed and the structural motion is prescribed as an input. This analysis shows that the motion of the elastic TE produces two opposite effects on the acoustic spectra: noise attenuation at the characteristic frequencies of the incident pressure field generated by the wake of the cylinder and noise excess at the natural frequencies of the fluid-loaded beam. [ABSTRACT FROM AUTHOR]

The self-noise of an isolated controlled-diffusion aerofoil is investigated using direct noise computations. The high-order finite difference solver HiPSTAR is used to conduct four large eddy simulations and one direct numerical simulation. The motivation is to investigate the multiple sources of aerofoil self-noise on a realistic compressor blade geometry, the physical mechanisms leading to the generation of sound, and the influence of compressibility effects. The angle of attack is 8°, the chord-based Reynolds number is 105, and results obtained for four values of the free-stream Mach number [0.2, 0.3, 0.4, 0.5] are compared. For those flow parameters, the pressure side is fully laminar, whereas a separation bubble is present on the suction side close to the leading edge that promotes transition to turbulence. The size of the separation bubble is found to increase with the Mach number. Two noise sources are observed, one at the trailing edge and one in the leading edge transition/reattachment region. The first has a broadband, low frequency spectrum, while the second displays a tone whose frequency depends on the local Mach number. Because the leading edge separation bubble is very small, the associated tone frequency is high and requires a significantly finer grid to faithfully resolve the acoustic propagation than what is typically deemed sufficient. Finally, cross correlations between the surface pressure and the far-field pressure reveal that the pressure fluctuations reaching the trailing edge are initially generated in the transition/reattachment region, which indicates that the trailing edge noise is a consequence of the pressure fluctuations generated by the separation bubble. [ABSTRACT FROM AUTHOR]

A compressible direct numerical simulation of the flow and near-field acoustics over a Controlled-Diffusion aerofoil at 8∘ geometrical angle of attack, a Reynolds number of R e c = 1.5 × 10 5 based on the chord and a free-stream Mach number of M = 0.25 , with a spanwise extent of 12% chord was conducted to produce a space-time database to determine the aerofoil noise generation mechanisms. Three main noise sources respectively from the flow separation and reattachment at the leading-edge on the suction side, the interaction between the attached turbulent flow on the suction side and the trailing-edge and an additional near-wake source due to secondary instability are observed. It should be noticed, however, that such an additional source may strongly depend on the aerofoil shape and also on the flow conditions. The contribution of these three noise sources to the noise radiation at different frequencies is analysed through a wavenumber-frequency filtering procedure. In the low to mid frequency range, the contribution of the traditional trailing-edge noise is significant, whereas at high frequencies, the noise radiation is shown to be related to the flow transition/reattachment in the leading-edge area and the flow in the near wake further downstream of the trailing-edge noise source center. The near-wake source is observed to dominate the high-frequency noise radiation. [ABSTRACT FROM AUTHOR]

This paper presents a new data-driven approach for the establishment of empirical models describing turbulent boundary layer wall-pressure spectra. Unlike other models presented in literature, the new models are not derived by extending previously existing ones, but are directly built from a given dataset through symbolic regression using a machine learning algorithm known as Gene Expression Programming. Two modifications of the GEP algorithm presented in literature are proposed in this work to cope with some issues that are specific to the modelling of wall pressure spectra: a new power terminal and a local optimization loop. The validity of the new approach is first demonstrated using as input a dataset synthesized following the Chase-Howe and Goody models. The method is then applied to experimental data for a flat plate boundary layer. The results indicate that the wall pressure model obtained with the proposed approach remains consistent with previous formulations for zero pressure gradient, while showing a better match with the data and suggesting new ways to predict the influence of moderate pressure gradient [ABSTRACT FROM AUTHOR]

Hybrid methods for trailing edge noise prediction are applied to a NACA 6512-63 airfoil at zero angle of attack embedded in the actual narrow stream of the Siegen open-jet anechoic wind-tunnel. Incompressible large-eddy simulations (LES) of the airfoil trailing-edge flow yield the noise sources, and different analytical acoustic analogies then predict the acoustic far-field pressure. The chord-based Reynolds number of 1.9 × 105 triggers long laminar flow regions along the airfoil, leading to instability waves with an associated broadband noise radiation. To avoid this laminar-instability noise, the airfoil boundary layer is experimentally tripped on both sides. The actual serration tripping is included in the numerical grid as well as a simplified stair-step trip. Aerodynamic and acoustic results of the performed LES are compared with measurements and compressible direct numerical simulation (DNS) results. No noticeable influence of the trip geometry is found whereas the numerical schemes play a more important role. LES predictions with boundary layer trip compare well with experimental measurements and DNS cases. If the boundary-layer trip is not sufficiently thick in the numerical model, or not present at all, the far-field acoustic pressure is overpredicted and resembles more closely the sound radiation from the untripped airfoil with laminar instability noise. [ABSTRACT FROM AUTHOR]