Wings operated at low and moderate Reynolds number such as the ones of UAVs or the blades of small turbine and of compressors, can be the source of an aero-acoustic phenomenon called laminar boundary layer instability noise. The narrow band noise can be attenuated using trailing edge (TE) serrations (Chong et al. 2013, Moreau et al. 2012) but the mechanism behind tonal attenuation is yet to be investigated in detail. An experimental study (1.32X105 < Re < 5.30X105) is conducted in an open-jet low turbulence wind tunnel using NACA 0018 airfoil (c = 0.20m) with modified TE. The acoustic emission in the far field is recorded using far-field microphones whereas the developing field near the TE is studied using Time resolved planar particle image velocimetry and oil flow visualization techniques. Study showed the serration-3 (2h = 20mm, λ = 10mm) has maximum tonal noise attenuation and effect the point of separation of laminar separation bubble. The span-wise change in flow characteristics in case of serrations can be attributed to the effective chord-length of the wing at each span-wise position. Further span-wise correlation of chord-wise flow fluctuations show the flow being turbulent upstream of serration-3 TE modification.

The aerodynamic noise generated by trailing edge of aircraft wings at take-o_ and approach con_guration continues to be a critical factor in the future development of aviation. Trailing edge noise is also a serious concern for wind turbines, marine propellers, helicopter rotors, axial and centrifugal turbines and compressors. A fundamental understanding of how the energy contained in turbulent motions around trailing edges is converted to sound will be bene_cial for designing silent airfoils and control surfaces. Current research aims at applying new tools of ow _eld interrogation i.e. high-speed planar and tomographic PIV to the canonical problem of trailing edge noise. In analytical models, the far _eld noise can be described using turbulent quantities like the spanwise coherence. The use of high-speed tomographic PIV creates a link between the phenomenological description of the ow and acoustic _eld in terms of coherence of structures and noise. The geometric con_guration used involves an airfoil with an elliptic leading edge and an asymmetric beveled trailing edge with 25 deg tip angle (_) and three di_erent radii of curvature (R). The measurements are taken at four di_erent Reynolds numbers.

One of the main limitations of microphone measurements in an open-jet wind tunnel is spectral broadening. The phenomenon of spectral broadening is observed when a single frequency sound wave passes through a free shear layer. Inside the turbulent shear layer the wave interacts with the turbulence, which leads to broadening of the tone in the frequency domain. The spectrum of a single frequency tone measured with a microphone will not only include the peak but also two shoulders containing the broadened frequencies, one on each side of the peak. Spectral broadening does not only a_ect single microphone measurements but also has a detrimental e_ect on acoustic array measurements. The distortion caused by the shear layer will reduce the coherence between the microphones in the array. This limits the size and thus resolution of the array. This spectral broadening is particularly problematic when examining objects which produce a tonal noise, e.g. open rotors, sharp edges.

Understanding, predicting and controlling laminar to turbulent transition is of great interest due to the wide range of practical applications where transition is significant. Generally, turbulent flow has an order of magnitude higher skin friction. Fuel savings of up to 25% would be possible for a large commercial aircraft if laminar flow could be maintained over the wings. At low to moderate Reynolds number several types of airfoils show the development of Tollmien-Schlichting (T-S) instabilities. These T-S waves become the dominant cause of transition which is further amplified by an adverse pressure gradient.