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Experimental investigations of a trailing edge noise feedback mechanism on a NACA 0012 airfoil

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Abstract

Discrete frequency tones in the trailing edge noise spectra of NACA 0012 airfoils are investigated with the Coherent Particle Velocity method. The Reynolds number and angle of attack range, in which these discrete frequency tones are present, are consistent with published results. The discrete tones are composed of a main tone and a set of regularly spaced side peaks resulting in a ladder-type structure for the dependency on the free stream velocity. The occurrence of this discrete frequency noise could be attributed to the presence of a laminar boundary layer on the pressure side opening up into a separation bubble near the trailing edge, which was visualized using oil flow. Wall pressure measurements close to the trailing edge revealed a strong spanwise and streamwise coherence of the flow structures inside this laminar separation bubble. The laminar vortex shedding frequencies inferred from the streamwise velocity fluctuations, which were evaluated from hot-wire measurements at the trailing edge, were seen to coincide with the discrete tone frequencies observed in the trailing edge noise spectra. Previous findings on discrete frequency tones for airfoils with laminar boundary layers up to the trailing edge hint at the existence of a global feedback loop. Hence, sound waves generated at the trailing edge feed back into the laminar boundary layer upstream by receptivity and are, then, convectively amplified downstream. The most dominant amplification of these disturbance modes is observed inside the laminar separation bubble. Therefore, the frequencies of the most pronounced tones in the trailing edge noise spectra are in the frequency range of the convectively most amplified disturbance modes. Modifying the receptivity behavior of the laminar boundary layer on the pressure side by means of very thin, two-dimensional roughness elements considerably changes the discrete tone frequencies. For roughness elements placed closer to the trailing edge, the main tone frequency was seen to decrease, while the frequency spacing in-between two successive tones increased. Based on the stability characteristics of the laminar boundary layer and the characteristics of the upstream traveling sound wave, a method for predicting the discrete tone frequencies was developed showing good agreement with the measured results. Hence, with a controlled modification of the laminar boundary layer receptivity behavior, the existence of the proposed feedback loop could be confirmed. At the same time, no significant influence of a second feedback loop previously proposed for the suction side of the NACA 0012 airfoil was observed neither by influencing the boundary layer with a receptivity–roughness element nor by tripping the boundary layer at the leading edge.

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  1. Institute of Aerodynamics and Gas Dynamics, Pfaffenwaldring 21, 70569, Stuttgart, Germany

    B. Plogmann, A. Herrig & W. Würz

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  1. B. Plogmann
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  2. A. Herrig
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  3. W. Würz
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Correspondence to B. Plogmann.

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Plogmann, B., Herrig, A. & Würz, W. Experimental investigations of a trailing edge noise feedback mechanism on a NACA 0012 airfoil. Exp Fluids 54, 1480 (2013). https://doi.org/10.1007/s00348-013-1480-z

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  • DOI: https://doi.org/10.1007/s00348-013-1480-z

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