A prediction method is developed for the self-generated noise of an airfoil blade encountering smooth flow. The prediction methods for the individual self-noise mechanisms are semiempirical and are based on previous theoretical studies and data obtained from tests of two- and three-dimensional airfoil blade sections. The self-noise mechanisms are due to specific boundary-layer phenomena, that is, the boundary-layer turbulence passing the trailing edge, separated-boundary-layer and stalled flow over an airfoil, vortex shedding due to laminar boundary layer instabilities, vortex shedding from blunt trailing edges, and the turbulent vortex flow existing near the tip of lifting blades. The predictions are compared successfully with published data from three self-noise studies of different airfoil shapes. An application of the prediction method is reported for a large scale-model helicopter rotor, and the predictions compared well with experimental broadband noise measurements. A computer code of the method is given.

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Airfoil self-noise and prediction

Trailing edge noise prediction from measured surface pressures

Refraction of sound by a shear layer

Hohenemser, K. H. and Yin, Shen-Kuang, "Some Applications of the Method of Multi-blade Coordinates," Journals of the American Helicopter Society, Vol. 17, No. 3, July 1972, pp. 3-12. Coleman, R. P., "Theory of Self-Excited Mechanical Oscillations of Hinged Rotor Blades," NACA Advanced Restricted Rept. 3G29, Republished as Rept. 1351,1943, NACA. Miller, R. H., "Helicopter Control and Stability in Hovering Flight," Journal of the Aeronautical Sciences, Vol. 15, No. 8, Aug. 1948, pp. 453-472. Azuma, A., "Dynamic Analysis of the Rigid Rotor System," Journal of Aircraft, Vol. 4, No. 3, May-June 1967, pp. 203-209. Curtiss, H. C., Jr. and Shupe, N. K., "A Stability and Control Theory for Hingeless Rotors," American Helicopter 27th Annual National V/STOL Forum, Washington, D. C., May 1971. Ward, J. F., "Exploratory Flight Investigation and Analysis of Structural Loads Encountered By A Helicopter Hingeless Rotor System," D-3676, Nov. 1966, NASA. Harris, F. D., "Articulated Rotor Blade Flapping Motion at Low Advance Ration," Journal of the American Helicopter Society, Vol. 17, No. 1, Jan. 1972, pp. 41-48. Ormiston, R. A. and Peters, D. A., "Hingeless Helicopter Rotor Response with Nonuniform Inflow and Elastic Blade Bending," Journal of Aircraft, Vol. 9, No. 10, Oct. 1972, pp. 730736.

Vortex Noise of Isolated Airfoils

The present paper is devoted to the experimental study of the noise generated by an airfoil profile placed in a uniform laminar flow. The far-field acoustic spectrum is shown to be composed of a broadband contribution around frequency fs and a discrete contribution at equidistant frequencies fn, which follow power laws of the forms fs ∼ U1.5 and fn ∼ U0.85. Both contributions can be accounted for by a simple model derived from the original suggestions of Tam (1974) and Fink (1975). It is essentially assumed that the diffraction of the Tollmien-Schlichting instabilities by the trailing edge generates acoustic waves which propagate in the far field and also trigger an aeroacoustic feedback loop, whose length is equal to the distance between the trailing edge and the maximum velocity point of the airfoil.

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Noise generated by airfoil profiles placed in a uniform laminar flow

A theoretical and experimental investigation of the nosie and unsteady surface pressure characteristics of an isolated airfoil in a uniform mean velocity, homogeneous, nearly isotropic turbulence field was conducted. Experiments were performed with a 23-cm chord, two-dimensional, NACA 0012 airfoil in the UTRC Acoustic Research Tunnel over a Mach number range of 0.1 to 0.5. Far-field noise spectra and directivity as well as surface pressure spectra and cross-spectra were obtained. Incident turbulence statistics were documented. Theory applied to predict far-field noise and surface pressure characteristics from measured inflow turbulence statistics showed good agreement with measurement over the dominant frequency range for all Mach numbers investigated. The theoretical formulation represents a first-principles solution providing absolute level prediction without recourse to empirical or adjustable constants. It takes into account compressibility as well as source noncompactness effects. Correlation measurements demonstrated that all chordwise portions of the airfoil radiated directly to the far-field, but that the leading edge was the dominant noise producing region. b c c0 d / g kx,ky,kz

Noise and Surface Pressure Response of an Airfoil to Incident Turbulence

This new wind tunnel, designed specifically for aerodynamic noise research, has a variable area, open jet test section enclosed in an anechoic chamber. Speeds up to a Mach number of 0.65 are possible with a test section area of 5 ft2. A separate, low noise level, high pressure air source is provided for jet noise studies. Design and construction considerations for the major tunnel sections (i.e., inlet, anechoic chamber, diffuser, muffler, driver) are described. Aerodynamic and acoustic calibration of the tunnel is discussed in relation to design criteria. Problems applicable to other acoustic tunnels such as background noise, edge tone suppression, deflected jet noise and distortion of noise directivity patterns by shear layer refraction are discussed. Initial testing demonstrated that a sufficiently low background noise level has been obtained to permit measurement of the noise radiated by the relatively weak noise source of an isolated airfoil placed in a uniform, low turbulence stream.

Design and Development of the United Aircraft Research Laboratories Acoustic Research Tunnel

An experimental investigation was conducted in the UARL Acoustic Research Tunnel to define the noise characteristics associated with the interaction of a stationary tip vortex and a downstream stationary airfoil. This model test geometry simulated, in its simplest form, the tip vortexblade interaction which occurs on single rotor helicopters during hover. For moderate to high lift test conditions, the vortex-airfoil interaction was found to cause local blade stall with an attendant increase in the blade far-field noise. These results indicated that this interaction may be an important source of helicopter broadband noise during hover. Cross-correlation measurements conducted amongst surface-mounted and far-field microphones demonstrated that the operative noise mechanism was "trailing edge noise" arising from the interaction of stall generated eddies with the airfoil trailing edge. This mechanism would be expected to be responsible for increased noise at stall conditions in other, nonrotary wing, applications.

Isolated Airfoil-Tip Vortex Interaction Noise

: The purpose of the study was to determine the vortex shedding noise characteristics of isolated airfoils in a Reynolds number range applicable to full-scale helicopter rotors. Measurements of far-field noise, airfoil surface pressure fluctuations, and correlation coefficients were obtained for three airfoils. Two models of differing thickness were tested as two-dimensional models (NACA 0012 and 0018 airfoils) and the third has a span of half the tunnel width (NACA 0012 airfoil). Airfoil vortex shedding noise was found to be discrete rather than broadband. The frequency dependence of this noise on velocity and chord was found to be well predicted by a constant Strouhal number near 0.2 referenced to laminar wake thickness at the airfoil trailing edge. The presence of a laminar boundary layer on the pressure surface of the airfoils was found to be critical to the presence of vortex shedding noise. (Author)

Robert W. Paterson

Papers

Vortex Noise of Isolated Airfoils

Noise and Surface Pressure Response of an Airfoil to Incident Turbulence

Design and Development of the United Aircraft Research Laboratories Acoustic Research Tunnel

Isolated Airfoil-Tip Vortex Interaction Noise

Vortex Shedding Noise of an Isolated Airfoil.