Instability Of Laminar Boundary Layer Research Articles

Aerodynamic and Aeroacoustic Effects of Different Transition Mechanisms on an Airfoil

The paper delves into the detailed sound generation and propagation mechanisms associated with tripping-induced flow perturbations and trailing-edge scattering using wall-resolved large-eddy simulations. Two distinct boundary-layer tripping techniques, namely a geometrically resolved stair strip and an artificially modeled trip using suction and blowing, are investigated. To facilitate comparison, the natural boundary-layer transition is also simulated as a baseline scenario. The analysis takes into account a Reynolds number of 4×105, a Mach number of 0.058, and a nonzero angle of attack of 6.25° over a NACA 0012 airfoil configuration. The mechanisms for sound generation and propagation related to trailing-edge noise remain consistent across the three transition scenarios. However, boundary-layer tripping notably leads to intricate, scenario-specific noise generation: there is an interaction between the laminar separation bubble and tripping for the stair strip case, whereas laminar boundary-layer instability is evident for the suction and blowing scenario. Aerodynamic flowfields involving acoustic noise sources, their propagating natures near the wall, and far-field acoustics are cross-examined in detail. The comprehensive analysis of observed phenomena provides valuable insights for understanding nonlinear flow and acoustic interactions relevant to airfoil noise and designing new types of trips under adverse pressure gradient flows.

Leading edge erosion detection for a wind turbine blade using far-field aerodynamic noise

In this paper, the feasibility of using far-field acoustic measurements as a non-contact monitoring technique for wind turbine blade leading edge erosion is assessed. For this purpose, a DU96 W180 airfoil with several eroded leading edge configurations of different severities is experimentally investigated. The eroded leading edges are designed with pits, gouges and coating delamination scaled from a real eroded blade. To assess the feasibility of the technique in quasi-realistic configurations, experiments are carried out under clean and turbulent inflow conditions. Acoustic measurements are performed with a phased microphone array. In the absence of inflow turbulence, because of the low Reynolds number at which the experiments are carried out, the case with minor erosion severity shows similar far-field noise spectra as the clean leading-edge cases, i.e., the presence of tonal peaks caused by laminar boundary layer instability noise through a self-sustained feedback loop but with higher tonal amplitudes. Increasing the damage level (considered as moderate erosion), the spectra of the noise scattered from the suction side show that the tonal peaks shift to higher frequencies and have lower amplitudes, thus suggesting that the damage alters the flow features responsible for the acoustic feedback loop; whereas, the spectra from the pressure side show a broadband noise distribution. For heavy erosion, the far-field noise spectra show broadband features from both airfoil sides, thus suggesting that the damage has fully forced the transition to turbulent flow; in which case, an increase in the low-frequency content is observed. Conversely, in the presence of turbulent inflow, when comparing the noise scattered at the trailing edge, no difference is found. However, leading edge impingement noise decreases at medium–high frequency compared with the baseline case at a chord-length-based Strouhal number StC∼10. The experimental results also suggest that the delamination feature is the one which is the most easily detectable and the approach is valid for a wide range of angles of attack and inflow velocity.

Experimental study of airfoil instability noise with wavy leading edges

Airfoils generate intensive laminar boundary layer instability noise at low-to-moderate Reynolds numbers. Experimental studies are conducted to investigate the noise radiation characteristics of a NACA 0012 airfoil operated at various angles of attack and Reynolds numbers. It is found that the instability noise changes from a broadband hump to intensive tonal noise with increasing angle of attack, while it changes from tonal noise to a broadband hump with increasing Reynolds number. The overall sound power is observed to follow the 3rd power of the inflow velocity. In addition, wavy leading edges are employed to reduce the instability noise. The wavy leading edges are in the form of sinusoidal profiles with two main design parameters of serration amplitude and wavelength. Parametric studies of the amplitude and wavelength are performed to understand the effect wavy leading edges on noise reduction. It is observed that the sound power reduction level is sensitive to both the amplitude and wavelength. Overall, the wavy leading edge with a large amplitude and small wavelength can achieve the most considerable noise reduction effect.

Computational Aeroacoustic Investigation of Airfoil Cascades

At moderate Reynolds numbers and angles of attack, the Laminar Boundary Layer (LBL) becomes unstable on the surface of airfoils, and causes periodic vortex shedding, which means undesired tonal peaks in the spectrum of the emitted aeroacoustic noise along with increased vibration and decreased aerodynamic performance. In the past, numerous research campaigns focused on the LBL vortex shedding, including measurements and numerical simulations as well. The results of these investigation showed that, the formation of the LBL instability related to the presence of the laminar separation bubble. It was also shown that, the spectrum of the emitted noise has a multitonal behavior, and the scaling of the mean frequency with the free stream velocity has a ladder structure. Based on these results, the LBL instability is a complex phenomenon; however, in the preliminary design of axial flow turbomachines the prediction of the frequency of the vortex shedding is essential, therefore the use of semi-empirical formulas is usual to achieve this goal. The previous researches mostly focused on separated airfoils, however, in case of turbomachines, the blades form a cascade, which can significantly affect the aerodynamic of the airfoils, i.e. it can affect the behavior of the LBL instability as well. According to this, in the present paper the LBL instability of NACA 0012 cascades are investigated, using 2D computational fluid mechanics and aeroacoustics simulations. The investigation involves the variation of the angle of attack, the chord based Reynolds-number and the spacing. The results are compared to the semi-empirical Brooks-Pope-Marcolini model.

Modal analysis of the laminar boundary layer instability and tonal noise of an airfoil at Reynolds number 150,000

A direct numerical simulation of the flow field around a controlled-diffusion airfoil within an anechoic wind-tunnel at [Formula: see text] incidence and a high Reynolds number of [Formula: see text] is performed for the first time using a lattice Boltzmann method. The simulation compares favorably with experimental measurements of wall-pressure, wake statistics, and far-field sound. The simulation noticeably captures experimentally observed high-amplitude acoustic tones that rise above a broadband hump. Both noise components are related to a breathing of a recirculation bubble formed around 65–70% of the chord, and to Kelvin-Helmholtz instabilities in the separated shear layer that yield rollers that break down into turbulent vortices whose diffraction at the trailing edge produces a strong dipole acoustic field. A wavelet analysis of the wall-pressure signals combined with some flow visualization has shown that the flow statistics are dominated by intense events caused by intermittent, large and intense bursting rollers. Several modal analyses of these events are performed on both the wall-pressure fluctuations and the span averaged flow field in order to analyse the boundary layer instability which triggers the typical sharp tones over a broadband hump in airfoil noise. A suction side Kelvin-Helmholtz instability is observed to be coupled with a pressure side vortex shedding induced by the sudden transition to turbulence and the blunt trailing edge.

Effect of external acoustic excitation on NACA0015 discrete tonal noise

The purpose of this work is to investigate the effect of external acoustic excitation on the discrete tonal noise generated by laminar boundary layer instability. Experimental testing inside anechoic wind tunnel was done on NACA0015 in order to measure the sound spectrum at low Reynolds number of Re∼104 and angle of attack of 0°, 3° and 5°. The effects of excitation amplitude of 70 dB and 90 dB and excitation frequency of 3000 Hz and 2000 Hz on the airfoil discrete tones were investigated. Exciting at excitation frequency closed to the primary frequency was found able to suppress tonal noise, however, is highly dependent on the angle of attack and freestream velocity. The airfoil was found to have a strong preference in its tonal frequency regardless of excitation. No change was found on the dependency of the secondary frequency. Each individual frequency was found to have ∼U0.8 dependency with freestream velocity that is similar to that without excitation. In general, observation shows acoustic excitation only affect the behavior of the primary frequency. The primary frequency dependency at α = 0° was found changed to ∼U2.0 with acoustic excitation at 3000 Hz and 90 dB and at 2000 Hz and 70 dB.

Laminar boundary layer instability noise produced by an aerofoil

When a laminar boundary layer exists on the surface of an aerofoil up to the trailing edge, a tone or a number of tones are sometimes produced. These tones have been the subject of a number of investigations which have proposed a variety of different mechanisms regarding their production. This paper gives a brief overview of the previously proposed mechanisms and then describes the development of a theoretical model to estimate the tone frequencies. The model is validated against a number of well-known published experiments and also against the results of an experimental investigation undertaken by the authors. The model is compared with other models available for predicting laminar boundary layer instability noise and is shown to be accurate and robust. Unlike previous models, which are empirical, the model presented in this paper is purely theoretical and could be used to predict the frequency of laminar boundary layer instability noise produced by an arbitrary aerofoil.

Stability and Transition of a Laminar Boundary Layer in an Axially Rotating Pipe. (Linear Stability Theory and Experiment).

Instability of laminar boundary layer developing in the inlet region of an axially rotating pipe was studied theoretically using small-disturbance stability theory. Stability diagrams were obtained by numerically solving the perturbation equations as an eigenvalue problem. The disturbance wave number at the maximum amplification factor agrees closely with that of the velocity fluctuations due to the spiral vortices measured by a hot wire probe. It is found that the superimposition of small disturbances generated by a loud speaker on a laminar boundary layer causes an upstream shift of the transition from laminar to turbulent flow.

Elongational Viscosity and Secondary Instability of Laminar Boundary Layer

Drag reduction phenomena of liquid flows due to an addition of polymer molecules are caused by an interaction of the elongational viscosity of the polymer solution and the elongational-type secondary flow of the base flow. In transition form laminar to turbulent flow on a flat plate boundary layer the elongational viscosity acts on the secondary vortices appearing by amplification of Tollmien-Schlichting waves and extends the stable region of Tollmien-Schlichting waves. At the same time the wave number corresponding to the critical state of neutral stability for the waves becomes larger with an increase in elongational viscosity. By this analysis it is suggested that the addition of polymer can delay the turbulent transition of boundary layer flow.

The laminar boundary layer instability excitation of an acoustic resonance

Tollmien-Schlichting waves on a 10C4 airfoil spanning the working section of a low speed wind tunnel have been discovered as the source of excitation of the Parker β-mode resonance. The feedback loop of the self-excited resonance is completed by acoustic coupling to the origin of the instability at the critical Reynolds number and occurs within the inner region of this laminar boundary layer instability. The coupling frequency range and amplification rate are predictable from instability theory. In the self-exciting system the sound enables the boundary layer to remain attached and laminar to the trailing edge which is far past the mid-chord separation point calculated by using Thwaites' method. In this case the unsteady Kutta condition does not hold to the order of the instability pressure fluctuations in the boundary layer.

Reynolds Number Effects in Cascades and Axial Flow Compressors

A series of tests on guide vane and compressor cascades is reported. The Reynolds number was varied in the guide vane cascade tests, and the Reynolds number and the cascade aspect ratio were varied in the compressor cascade tests. The substantial laminar separation observed in the compressor cascades at high aspect ratio (and low Reynolds number) was suppressed in the cascade tests at low aspect ratio, 2:1. Effects of Reynolds number variation on the performance of a single stage axial flow compressor are also given, and compared with predictions of performance using the cascade tests. Calculations of laminar separation points agree quite well with the experimental observations. It appears that transition due to laminar boundary layer instability is unlikely to occur on compressor blades, in the normal operating range of Reynolds number.