Wind Tunnel Aeroacoustic 

Tests of Six Airfoils for Use 

on Small Wind Turbines

Preprint 

December 2003 • NREL/CP-500-35090 

P. Migliore 

National Renewable Energy Laboratory

S. Oerlemans 

National Aerospace Laboratory 

NLR Department of Aeroacoustics 

To be presented at the 2004 AIAA Wind Energy Symposium 

Reno, Nevada 

January 5–8, 2004 

National Renewable Energy Laboratory 

1617 Cole Boulevard 

Golden, Colorado 80401-3393 

NREL is a U.S. Department of Energy Laboratory 

Operated by Midwest Research Institute • Battelle 

Contract No. DE-AC36-99-GO10337 
NOTICE 

The submitted manuscript has been offered by an employee of the Midwest Research Institute (MRI), a 

contractor of the US Government under Contract No. DE-AC36-99GO10337. Accordingly, the US 

Government and MRI retain a nonexclusive royalty-free license to publish or reproduce the published 

form of this contribution, or allow others to do so, for US Government purposes. 

This report was prepared as an account of work sponsored by an agency of the United States 

government. Neither the United States government nor any agency thereof, nor any of their employees, 

makes any warranty, express or implied, or assumes any legal liability or responsibility for the accuracy, 

completeness, or usefulness of any information, apparatus, product, or process disclosed, or represents 

that its use would not infringe privately owned rights. Reference herein to any specific commercial 

product, process, or service by trade name, trademark, manufacturer, or otherwise does not necessarily 

constitute or imply its endorsement, recommendation, or favoring by the United States government or any 

agency thereof. The views and opinions of authors expressed herein do not necessarily state or reflect 

those of the United States government or any agency thereof. 

Available electronically at http://www.osti.gov/bridge

Available for a processing fee to U.S. Department of Energy 

and its contractors, in paper, from: 

U.S. Department of Energy 

Office of Scientific and Technical Information 

P.O. Box 62 

Oak Ridge, TN 37831-0062 

phone: 865.576.8401 

fax: 865.576.5728 

email: reports@adonis.osti.gov 

Available for sale to the public, in paper, from: 

U.S. Department of Commerce 

National Technical Information Service 

5285 Port Royal Road 

Springfield, VA 22161 

phone: 800.553.6847 

fax: 703.605.6900 

email: orders@ntis.fedworld.gov 

online ordering: http://www.ntis.gov/ordering.htm

Printed on paper containing at least 50% wastepaper, including 20% postconsumer waste 
1

WIND TUNNEL AEROACOUSTIC TESTS OF SIX AIRFOILS 

FOR USE ON SMALL WIND TURBINES*§

Paul Migliore 

National Renewable Energy Laboratory, National Wind Technology Center 

1617 Cole Boulevard, Golden, Colorado 80401, USA 

paul_migliore@nrel.gov

Stefan Oerlemans 

National Aerospace Laboratory NLR, Department of Aeroacoustics 

P.O. Box 153, 8300 AD Emmeloord, The Netherlands 

stefan@nlr.nl

 ABSTRACT *§

Aeroacoustic tests of seven airfoils were performed in 

an open jet anechoic wind tunnel. Six of the airfoils are 

candidates for use on small wind turbines operating at 

low Reynolds number. One airfoil was tested for comparison to benchmark data. Tests were conducted with 

and without boundary layer tripping. In some cases a 

turbulence grid was placed upstream in the test section 

to investigate inflow turbulence noise. An array of 48 

microphones was used to locate noise sources and separate airfoil noise from extraneous tunnel noise. Trailing 

edge noise was dominant for all airfoils in clean tunnel 

flow. With the boundary layer untripped, several airfoils exhibited pure tones that disappeared after proper 

tripping was applied. In the presence of inflow turbulence, leading edge noise was dominant for all airfoils. 

 

*

 This work was performed at the National Renewable Energy 

Laboratory in support of the U.S. Department of Energy under Contract No. DE-AC36-99GO10337. 

§

 This material is declared a work of the U.S. Government and 

is not subject to copyright protection in the United States. 

INTRODUCTION

The U.S. Department of Energy, working through its 

National Renewable Energy Laboratory (NREL), is 

engaged in a comprehensive research effort to improve 

the understanding of wind turbine aeroacoustics. The 

research is motivated by the desire to make use of the 

large expanse of low wind speed sites that tend to be 

closer to U.S. load centers. Quiet wind turbines are an 

inducement to widespread deployment, so the goal of 

NREL’s aeroacoustic research is to develop tools for 

use by U.S. industry in developing highly efficient, 

quiet wind turbines for deployment at these low wind 

speed sites. NREL’s National Wind Technology Center 

(NWTC) is implementing a multi-faceted research approach that includes wind tunnel tests, field tests and 

theoretical analyses in direct support of low wind speed 

turbine development by its industry partners. NWTC 

researchers are working hand-in-hand with industry 

engineers to ensure that research findings are available 

to support ongoing design decisions. 

Figure 1. Wind tunnel airfoil models.
2

The work described herein focuses on the experimental 

aeroacoustic analysis of six airfoils that are candidates 

for use on small wind turbines. However, without 

knowledge of both the aerodynamic and aeroacoustic 

performance of airfoils, engineers are frustrated in making decisions on new blade designs. This is particularly 

true for small wind turbines, which operate at low Reynolds numbers where airfoil aerodynamic characteristics are both sensitive and difficult to predict. Thus, the 

present work must be considered in the context of the 

broader research effort that includes wind tunnel aerodynamic and aeroacoustic tests. The aerodynamic tests 

were conducted at the University of Illinois at UrbanaChampaign, and the aeroacoustic tests were conducted 

at the Netherlands National Aerospace Laboratory 

(NLR) in Emmeloord, Netherlands. The results, documented in NREL reports [1,2] and several papers in 

these conference proceedings, provide a valuable airfoil 

database for designers who wish to consider the tested 

airfoils. 

For the acoustic tests described in this paper, the twodimensional airfoil models were mounted between two 

acoustically treated endplates and tested with and without boundary layer trips and a turbulence grid, at three 

angles of attack and five wind tunnel speeds corresponding to Reynolds numbers from 0.2 to 1.0 million. 

Approximately 500 configurations were tested. Results 

were obtained in the form of 1/3-octave band acoustic 

“source plots” using conventional sum-and-delay beam 

forming, which allows separation of source and background noise. Further processing provided noise spectra and overall sound pressure1

 levels. The principal 

objective was to obtain a relative comparison of the 

prominent noise sources for the tested airfoils and conditions. Significant results are presented as a compari-

 

1

 Sound is characterized by small pressure fluctuations overlaying atmospheric pressure, but the human ear does not respond linearly to the amplitude of sound pressure [4]. Doubling the amplitude produces the sensation of louder noise, 

but it seems far less than twice as loud. For this reason, the 

scale used to characterize sound pressure amplitudes is logarithmic, which is an approximation of the actual response of 

the human ear. The definition of sound pressure level Lp is 

Lp = 10 · log [ p2 ÷ p2

ref ] expressed in decibels, dB, 

where, p is the root mean square sound pressure and pref has a 

value of 20 μ Pa corresponding to the weakest audible sound 

– the threshold of human hearing – at a frequency of 1000 Hz. 

A sound spectrum shows the distribution of acoustic energy 

as a function of frequency in either 1/3-octave or narrow 

bands. The overall sound pressure level represents the total 

acoustic energy, obtained by summing the spectral values 

over the whole frequency range. 

son of noise emissions for the various airfoils, the influence of inflow turbulence on leading-edge noise and the 

effect of boundary layer tripping. Because this paper is 

a significant condensation of the larger NREL report 

[2], many details have been omitted. 

SOURCES OF AIRFOIL NOISE

There are six different sources that independently generate airfoil acoustic emissions [3,4]: inflow turbulence, 

turbulent boundary layer trailing edge interaction, separating flow, laminar boundary layer vortex shedding, 

trailing edge bluntness (von Karman) vortex shedding 

and tip vortex formation. These sources are superimposed to form the total noise spectrum from a wind 

turbine blade. The spectra are often summed to calculate an overall sound power2

 level. 

Inflow turbulence noise caused by the interaction of the 

leading edge of an airfoil with a turbulent inflow is often called leading edge noise. Researchers currently 

think that sharp leading edge geometries are more susceptible to inflow turbulence noise. 

The other sources of noise are collectively called airfoil 

self noise, because they are caused by the airfoil interacting with its own boundary layer and near wake. If 

the trailing edge thickness of the airfoil is very thin 

relative to the boundary layer thickness, as was the case 

for the models tested, there will be no trailing edge 

bluntness noise. And two-dimensional airfoil models 

tested between endplates do not have a tip vortex or any 

associated noise, although interaction between the endplate boundary layer and model-endplate juncture may 

cause extraneous noise. 

Turbulent boundary layer trailing edge noise is generally considered to be the most important source of airfoil self noise for modern wind turbine blades. In this 

phenomenon, the unsteady pressure waves in the turbulent boundary layer are amplified and radiated by the 

 

2

 Whereas sound pressure level is a property of the observer 

location [4], the total strength of a source of sound is characterized by the sound power emitted by the source. In general, 

the sound power P transmitted through a surface S is the integral of the sound intensity I (energy transmitted per unit time 

and unit area) over the surface. If the surface S encloses the 

source of the sound, then P is the total sound power emitted 

by the source. The definition of sound power level is 

Lw = 10 · log [ P ÷ Pref ] expressed in decibels, dB, 

where Pref =10-12 watts is the standard reference sound power. 

The eardrum can detect incoming sound power as weak as 

one picowatt, and exposure to incoming sound power of more 

than one watt will result in some hearing loss. 
3

sharp trailing edge. As the angle of attack increases, 

the thickness of the turbulent boundary layer increases 

and large-scale unsteady structures can dominate noise 

production from the trailing edge. For fully separated 

flow, noise can radiate from the entire chord. 

Laminar boundary layer vortex shedding noise is created by a feedback loop between vortices being shed at 

the trailing edge and instability waves in the laminar 

boundary layer upstream. This source of noise can occur on either the suction or pressure side of the airfoil, 

and it can be particularly annoying because it is often 

manifested in pure tones that result from feedback amplification. It is not likely to be important for large 

turbines operating at high Reynolds number, but it may 

be significant for small wind turbines. 

 

AIRFOIL MODELS

Tests were conducted on seven airfoil models, each 

having a 22.86 cm chord and a 0.51 m span. Six of the 

airfoils are either being used or considered for use on 

small wind turbine blades. These are the FX63-137, 

S822, S834, (Selig-Donovan) SD2030, (Selig-Giguere) 

SG6043 and (Selig-Hanley) SH3055. The models are 

shown in Figure 1 along with a NACA 0012 profile 

(center of photograph) used as a benchmark for comparison to previous results obtained by the National 

Aeronautics and Space Administration (NASA) [5]. 

The airfoil models were built to exacting tolerances, 

where the difference between the specified and as-built 

surface was required to be within 0.05% of the model 

chord, or 0.1143 mm. Trailing edge thickness was to 

be no greater than 0.375 mm with 0.1875 mm as a preferred upper limit. A coordinate measuring machine 

was used to verify the final dimensions, and in virtually 

every case, the model geometry was within the specified tolerances. Figure 2 shows a typical result. The 

noticeable deviation at the trailing edge is because the 

airfoil coordinates result in a trailing edge thickness of 

zero, which is impossible to fabricate. 

Figure 2. S822 airfoil model accuracy (difference 

between specified and measured coordinates) 

The NACA 0012 model was constructed of carbonfiber fabric pre-impregnated with epoxy resin and cured 

at room temperature. Although the accuracy of this 

model was acceptable, we decided to construct the remaining models of solid aluminum to accelerate the 

fabrication process. 

 

The S822 model was also used to test micro electro 

mechanical systems (MEMS) tabs, which are being 

investigated for aerodynamic controls. These devices 

were mounted at 95% chord on the pressure side of the 

airfoil, using double-sided adhesive tape. Two sets of 

four MEMS were tested, comprising various tab heights 

and spacing. Although space limitations preclude reporting the test results in this paper, they are detailed in 

the NREL report [2]. 

For tests with a fully turbulent boundary layer, tripping 

was initiated using zigzag tape of width = 5% chord 

over the entire model span at 2% and 5% chord on the 

suction and pressure sides of the airfoil, respectively. 

The standard trip thickness was 0.25 mm, but for some 

conditions, trips of up to 0.5 mm were used. In some 

cases a stethoscope was employed to verify whether or 

not the trips induced the desired boundary layer transition. The stethoscope was attached to an L-shaped totalpressure tube, which was traversed manually over the 

surface of the model. Transition from a laminar to a 

turbulent boundary layer was observed by listening. 

The geometric angle of attack (α) set in the wind tunnel 

was corrected [5] to account for flow curvature, resulting in an effective angle of attack (αeff). 

MEASUREMENTS AND DATA ANALYSES

The aeroacoustic tests were conducted in NLR's Small 

Anechoic Wind Tunnel KAT (Figure 3). This open circuit wind tunnel, the test section of which is surrounded 

by an anechoic room that is completely covered with 

foam wedges, yields more than 99% sound absorption 

above 500 Hz. Horizontal endplates are mounted to the 

upper and lower sides of a rectangular 0.38 m x 0.51 m 

nozzle, providing a semi-open test section for airfoil 

self noise measurements. To suppress reflections, the 

endplates are acoustically lined with layers of sound 

absorbing foam covered by perforated plates. For inflow turbulence measurements, a turbulence grid consisting of diagonally oriented, cylindrical 12 mm bars 

with a mesh width of 60 mm is installed in the nozzle. 

Although the tunnel is equipped with a force balance, 

that information was not used in these tests. 

Test Section Calibration

To verify acceptable flow quality, a calibration study 

was performed using hot-wire anemometers in the 

empty test section. Cross hot-wires were used to measure the turbulence intensity and flow angularity with 
4

and without the turbulence grid. Hot wire traverses 

were made in both cross flow directions for two Mach 

numbers, M=0.12 and M=0.18, at two axial locations 

roughly corresponding to the position of the leading and 

trailing edge of the tested models. 

The test section calibration resulted in profiles of axial 

velocity, axial turbulence intensity, lateral turbulence 

intensity and flow angularity. Without the turbulence 

grid, the central part of the test section had turbulence 

levels of 1% or less. With the turbulence grid, the levels 

increased to 9% and 5% at locations corresponding to 

the model leading and trailing edges, respectively. Although the flow angularity increased slightly with increasing speed and turbulence level, it always remained 

within 1o

 in the central area of the test section. 

Microphone Array

Sound pressure level data were acquired using an 

acoustic array consisting of 48 1⁄2-inch LinearX M51 

microphones mounted in an open grid. Frequency dependent sensitivities of the individual microphones 

were taken from calibration sheets. No corrections were 

applied for microphone directivity, because this effect 

is the same for all airfoils and less than 2 dB for angles 

up to 45o

 and frequencies up to 15 kHz. Phase matching 

of the microphones was checked prior to the test using a 

calibration source at a known position. 

 

To obtain high resolution at low frequencies, the array 

dimensions needed to be rather large (0.8 m x 0.6 m2

). 

The microphone pattern was designed for maximum 

side lobe suppression at frequencies between 1 kHz and 

20 kHz. The array was placed outside the tunnel flow 

0.6 m from the tunnel axis, either on the suction or 

pressure side of the model as dictated by the test matrix. 

The relatively small distance between the array and the 

model was chosen to obtain a maximum signal-to-noise 

ratio. The center of the array was placed at the same 

height as the tunnel axis. Thus, the levels measured by 

the array represent airfoil noise levels radiated in the 

average direction of the array microphones. Because the 

source directivity for trailing edge and inflow turbulence noise is expected to be the same for all airfoils, 

the comparisons of noise from different airfoils is valid. 

Acoustic data from the array microphones were synchronously measured at a sample frequency of 51.2 

kHz and a measurement time of 30 s. A 500 Hz highpass filter was used to enhance the dynamic range. 

Test Program

Array measurements were made on the suction side of 

the six candidate airfoils, with and without tripping and 

turbulence grid for three angles of attack. The NACA 

0012 airfoil was tested for the same conditions as in 

Reference 4: with and without trip, without turbulence 

grid, for four Reynolds numbers and five angles of attack. An overview of the suction-side test matrix is 

shown in Table 1. For a number of conditions, pressure 

side array measurements were also made to determine 

directivity effects. Some measurements were repeated 

with thicker zigzag tape to assess its effectiveness in 

tripping the boundary layer. 

Data Processing

Processing methods and computer programs are referenced in the test report [2], but details are not discussed 

because of their proprietary nature. 

Conventional sum-and-delay beam forming was applied 

to obtain acoustic source plots as illustrated in Figure 4. 

The effect of sound refraction by the tunnel shear layer 

was corrected using the Amiet method [6], and the array scan plane was placed in the plane of the model and 

rotated in accordance with the angle of attack. Using 

these source plots, noise originating from the model 

was separated from background noise. 

A number of special measures were taken in the beamforming process. First, the main diagonal in the cross 

power matrix (auto powers) was discarded to suppress 

the influence of tunnel background noise. Second, a 

spatial window was applied to the microphone signals 

to reduce the effective array aperture with increasing 

frequency and reduce coherence loss effects. 

For quantitative comparison of different airfoils and 

conditions, the array results were processed to obtain 

narrowband or 1/3-octave-band spectra for specific 

source regions (again, the main diagonal in the cross 

power matrix was discarded). The acoustic data were 

processed using a block size of 2048, yielding a narrowband frequency resolution of 25 Hz. By defining an 

integration contour around the mid-span of the model, 

extraneous noise sources at junctions of the model and 

end plates were suppressed. A preliminary review of 

the test results showed that for measurements without 

the turbulence grid, noise was radiated from the trailing 

edge of the model. Therefore, in those cases, the midspan integration area was centered on the trailing edge. 

For measurements with the turbulence grid, the dominant noise source was observed to be located at the 

leading edge of the model. In those cases, the integration contour was centered on the leading edge. This 

process is illustrated in Figure 5. 

Because the integration area cuts through the line 

source region at the leading or trailing edge, “leakage” 
5

from sources outside the integration area into the integration contour, and vice versa, will occur. The magnitude of this effect depends on array resolution, and 

therefore on frequency. To account for this effect, a 

“line source correction” was applied using simulations 

similar to those described by Oerlemans and Sijtsma 

[7]. The resulting spectral levels are Sound Power Levels produced by 10 cm of span. 

Extraneous Noise Sources

In some cases, the airfoil noise levels were so low that, 

despite the procedures described above, the spectra 

were dominated by extraneous wind tunnel noise 

sources. To facilitate rapid judgment of the validity of 

the measured levels, procedures were developed to indicate the importance of tunnel noise in the spectra. 

For a significant number of conditions, the trailing edge 

noise levels were influenced by extraneous sources at 

the model-endplate junctions. These “corner sources” 

are illustrated in Figure 6. A routine was developed to 

determine the importance of these corner sources and, 

in cases where their influence on the trailing edge noise 

level is more than 1 dB, to calculate an upper limit for 

the trailing edge noise level. In graphs of the trailing 

edge noise spectra, these upper limits are indicated by 

the absence of a marker (symbol) at that specific frequency. If significant corner sources are absent, calculated noise levels are assumed to be absolute and a 

marker is placed at that specific frequency on the spectral plots. These data processing methods were successful in isolating the corner sources and mitigating 

their influence on calculated noise levels. Furthermore, 

a technique was identified for eliminating the corner 

sources entirely in future tests [2]. 

For measurements with the turbulence grid in place, 

background noise from the grid itself, rather than corner 

sources, often obscured the leading edge noise levels. In 

those cases, leading edge noise levels were compared to 

those obtained in the empty test section (with turbulence grid) for the same speed. If the leading edge noise 

level was at least 6 dB higher than the background 

noise level, a marker is placed at that specific frequency 

on the spectral plots. Absence of a marker indicates that 

the spectral level was influenced by grid noise. 

RESULTS

A large amount of data was acquired and analyzed in 

the experiments, but only a small amount can be presented in this paper. The reader may wish to obtain the 

full NREL report [2], which includes graphical presentations and narrative discussion of the following topics: 

 Test section calibration data 

 Sound power level trailing edge noise spectra for 

the six small wind turbine airfoils tested, including 

the effect of boundary layer tripping 

 Comparison of trailing edge noise data to benchmark NACA 0012 data [5] 

 Sound power level leading edge noise spectra for 

the six small wind turbine airfoils tested, including 

the effect of boundary layer tripping 

 Normalized leading and trailing edge noise spectra 

illustrating collapse of data for different speeds 

 Emission of pure tones for several airfoils, including suppression by aggressive boundary layer trips, 

 Directivity tests with the microphone array on both 

pressure and suction sides of the model 

 Leading edge noise and trailing edge noise sound 

power level comparison for all airfoils at 32 m/s 

and several angles of attack 

 Trailing edge noise spectra for various MEMS configurations tested on the S822 airfoil 

 Discovery of techniques to significantly reduce 

extraneous corner source noise. 

Trailing Edge Noise Spectra

Figure 7 illustrates typical trailing edge noise spectra 

obtained from acoustic source plots such as those 

shown in Figure 4. Results for the tripped boundary 

layer exhibit smooth broadband spectra typically associated with trailing edge noise. Results for the untripped boundary layer show some significant peaks 

that can be associated with laminar boundary layer vortex shedding [5]. The frequency of the peaks increases 

with speed at the angle of attack shown. 

We investigated the directivity of trailing edge noise by 

conducting certain tests with the array first on the suction side and then on the pressure side. The directivity 

was found to be symmetrical about the chord, as demonstrated by Figure 8. Speed dependence, occurrence 

of tones, and comparison between different airfoils will 

be discussed in subsequent sections. 

Leading Edge Noise Spectra

When the airfoils were tested with the turbulence grid 

installed in the tunnel, leading edge inflow turbulence 

noise became the dominant source for all airfoils. This 

was clearly indicated in the source plots (Figure 9), 

which reveal no prominent trailing edge emissions. 

The dynamic range of the source plots was 12 dB, 

which implies that the leading edge source was much 

greater than the trailing edge source. For the S822 airfoil at the same wind speed and angle of attack, the 

peak leading edge noise level (Figure 10) with the turbulence grid was around 90 dB compared to a peak 
6

trailing edge noise level (Figure 7) of approximately 70 

dB without the grid. With the turbulence grid, tripped 

and untripped results were identical, and leading edge 

noise directivity was found to be symmetrical around 

the chord. 

 

It is important to note that the level of turbulence in the 

tunnel was much greater than is typically experienced 

in the atmosphere at typical rotor speeds. The literature 

does not provide any evidence that leading edge noise 

is a prominent source for large wind turbines, and conventional wisdom holds that trailing edge noise is 

dominant. NREL researchers are currently formulating 

an experimental approach to investigate this hypothesis. 

Normalized Noise Spectra

To investigate the speed dependence of the airfoil noise 

levels, it is useful to normalize the noise spectra (Figure 

11). Here, normalized sound power level, given by 

PWL norm = PWL - 10 · log (Um); PWL is the airfoil sound 

power level determined from array measurements; the 

exponent m denotes the speed dependence of airfoil 

noise levels. Sound intensity p2

 is proportional to tunnel 

speed Um, with p the acoustic pressure. PWL norm is 

plotted against Strouhal number St=f · c/U, where f is 

the acoustic frequency and c is the model chord. Typically, Strouhal scaling is based on boundary layer 

thickness, but because that information was not available, the model chord was used. 

The best data collapse was obtained for a value of the 

exponent m = 4.5 determined iteratively. This is slightly 

lower than the value of 5 found in Reference 4. Some 

explanations are discussed in the full NREL report [2]. 

The S822 airfoil data in Figure 11 is provided for illustration, but review of the spectra for all airfoils showed 

very good data collapse for trailing edge noise in 

tripped conditions. For a given angle of attack, the 

trailing edge noise levels at different speeds coincided 

within 1-2 dB. Because normalization in terms of St and 

PWLnorm works well for the turbulent boundary layer 

trailing edge noise, experimental results for one speed 

can be extrapolated to other speeds. 

For untripped conditions, peak Strouhal numbers for 

different speeds were found to coincide within about 

30%. A slight increase in St with increasing speed suggests that a better collapse of peak frequencies could be 

obtained by using the boundary layer thickness as the 

length scale in St rather than chord, because boundary 

layer thickness at the trailing edge will decrease with 

increasing Reynolds number [5]. 

The spectral levels for the untripped results did not collapse as well as for the tripped data. Other values of the 

exponent m did not significantly improve the collapse. 

This illustrates that the normalization in terms of St and 

PWLnorm was not very successful for the complex feedback mechanism associated with laminar boundary 

layer vortex shedding noise, which is an important 

source for untripped airfoils at low Reynolds numbers. 

We also normalized the leading edge (inflow turbulence) noise spectra in a manner similar to that described for trailing edge noise. The best collapse of 

data was obtained for the exponent m = 6, which is in 

good agreement with theoretical predictions for low 

frequency inflow turbulence noise [4]. Figure 12 provides an example for the S822 airfoil. 

Comparison to Benchmark

Acoustic measurements on the NACA 0012 airfoil enabled direct comparison to benchmark data from NASA 

[5]. The tested airfoil shapes and chords were identical, 

and in both studies, the model was mounted between 

endplates attached to opposite sides of a rectangular 

tunnel nozzle. There were some differences in the manner of boundary layer tripping, but the NLR measurements were done at the same tunnel speeds and effective angles of attack as the NASA study. 

Detailed comparisons of the test results are beyond the 

scope of this paper, but in general, the spectral characteristics (Figure 13) agreed quite well. Broadband spectra for the tripped cases and spectral humps (tones) for a 

number of untripped cases were reproduced. The 

tripped results revealed an interesting difference between the NASA and NLR data. Although the sound 

power levels compared quite well for intermediate frequencies, the NASA data consistently exhibited a hump 

around 1 kHz that did not appear in the NLR data. Possible explanations for this behavior are the topic of continuing dialogue among NLR, NASA, and NREL researchers. Issues include differences in the manner of 

boundary layer tripping (NASA trips were more severe) 

and differences in measurement technique (NASA used 

a 2-microphone correlation method and NLR used a 48-

microphone array). 

Pure Tones

One of the most interesting observations of the test 

campaign was the presence of intense, narrowband 

peaks in the trailing edge noise spectra for several airfoils at different operating conditions. These are called 

“pure tones” and are perceived as such by a listener. 

They are illustrated in Figure 14 for the trailing edge 

noise spectra of the untripped S834 (α=10°), SG6043 

(α=0°) and SD2030 (α=0°) airfoils. The nature of these 

tones was investigated in more detail at 22.4 m/s, be-
7

cause they were most pronounced at this tunnel speed. 

The spectra for these cases show peaks at around 1 kHz 

and 2 kHz for all three airfoils. The angle of attack 

range for which these tones occurred was estimated by 

listening in the test section during a sweep of angle of 

attack. This gave the following ranges: 7.5°<α<13° for 

S834, -8°<α<2° for SG 6043, and -10°<α<4° for 

SD 2030. 

Although such tones sometimes result from blunt trailing edge vortex shedding, the extremely thin trailing 

edges of the models made this unlikely. Calculations of 

the Strouhal number (St), which is approximately 0.2 

for von Karman vortex shedding, confirmed this was 

not the cause. We hypothesized, therefore, that laminar 

boundary layer vortex shedding caused the tones. To 

investigate this hypothesis, we applied a different 

thickness of zigzag tape and observed the effects. By 

tripping one side of the airfoil at a time, we could determine where the tones originated. For the S834 and 

SD2030 airfoils, it was the pressure side, whereas for 

the SG6043 airfoil, the tones originated from the suction side. Stethoscope measurements on the pressure 

side of the untripped S834 model indicated a laminar 

boundary layer up to about 80% chord. With tripping, 

transition to a turbulent boundary layer occurred directly behind the trip at 5% chord. This observation 

supported the hypothesis that the tones were due to 

laminar boundary layer vortex shedding. 

The sensitivity of the tones to trip thickness was investigated in more detail for the S834 airfoil, with results 

shown in Figure 15. The untripped case (Figure 14) 

shows the narrowband peak at 925 Hz and the harmonic 

at 1850 Hz. It can be seen that the standard trip thickness of 0.25 mm on both sides of the airfoil was not 

effective. In fact, the level of the tones increased 

slightly with respect to the untripped case, and a harmonic appears at 2775 Hz. Application of a slightly 

thicker trip (0.30 mm) on the pressure side caused the 

spectral level to decrease dramatically and the 925 Hz 

and 1850 Hz tones to vanish completely. Interestingly, 

the broadband level decreases even further after the 

addition of a 0.30 mm trip on the suction side. While 

analysis of one case is not conclusive, this suggests that 

the suction side boundary layer dominates broadband 

noise production, whereas the pressure side may generate tones if it is not properly tripped. 

Another interesting observation was that the tones disappeared in the presence of upstream turbulence. Observation of the source plots with and without the turbulence grid showed that the inflow turbulence removed 

the trailing edge tones, and the grid noise became 

dominant. Evidently, the inflow turbulence interrupts 

the feedback mechanism responsible for the tones. 

Comparison of Airfoils

An important objective of the wind tunnel tests was to 

observe the noise levels of the different airfoils. All the 

models were not tested under identical conditions, but a 

common point for comparison does exist at 32 m/s and 

angles of attack of 0°, 10° and 18°. (The NACA 0012 

data were taken at a slightly different speed of 31.7 m/s 

and angles of attack of 0°, 9.5° and 16.5°.) 

 

The NREL report [2] compares the trailing edge noise 

spectra of the different airfoils. For the untripped condition, the noisiest airfoils were the SD 2030 and SG 

6043 at 0° and the NACA 0012 and S834 at 10°. For 

the tripped condition, in which pure tones were largely 

eliminated, noise levels were reduced in virtually every 

case. Figure 16, in which tripped and untripped spectra 

are displayed side-by-side, shows this trend. Although 

the comparison among airfoils is obscured by the fact 

that for many frequencies only upper limits are available (see ‘Extraneous Noise Sources’ above), it appears 

that the FX63-137 and SH3055 airfoils are somewhat 

noisier than the others. These observations are reflected 

in Figure 17, which shows the A-weighted3

 overall 

sound power level obtained by summing the 1/3-octave 

band sound pressure levels between 0.8 kHz and 12.5 

kHz. These overall sound power levels are an upper 

limit for the actual two-dimensional trailing-edge noise. 

We also compared the inflow turbulence noise spectra 

for the airfoils. Because the results were virtually identical for tripped and untripped conditions, only tripped 

data were examined. Differences in leading edge noise 

of 8 dB were observed between the quietest and noisiest 

airfoils. By examining these results in relationship to 

the airfoil shapes, we observed a general trend—the 

sharper the airfoil leading edge, the higher the inflow 

turbulence noise. This trend is demonstrated in Figure 

18, which plots the A-weighted overall noise levels for 

three angles of attack. These were obtained by summing the 1/3-octave band levels between 1 kHz and 2.5 

kHz where airfoil noise is attributed to inflow turbulence. Data in Figure 18 are presented from left to right 

in the order of decreasing airfoil leading edge radius. 

It is important to note that comparing airfoil noise data 

at the same angle of attack can be misleading, because 

 

3 The ear is not equally sensitive to tones of different frequencies. Maximum response occurs between 2 kHz and 3 kHz, 

where the hearing threshold is somewhat less than 0 dB. A 

100 Hz tone, however, must have an intensity of 40 dB to be 

heard [4]. Therefore, weighted sound levels have been introduced where lower frequencies are de-emphasized in a manner similar to human hearing. A-weighting is most commonly 

used and is well suited for sound levels that are not too high.
8

the airfoils may operate at different angles of attack. 

For example, peak aerodynamic efficiency for the 

FX63-137 occurs around α = 4o

 while the S822 is best 

around α = 8o [1]. Therefore, to estimate which airfoil 

will be quieter, we must compare at their respective 

design points (angle of attack, Reynolds number, surface condition, and tip speed). Following this logic, 

Figure 17 suggests that the FX63-137 (αeff=4.4o

, α=10o) 

is likely to be quieter than the S822 (αeff=7.9o

, α=18o). 

However, inspection of the S822 trailing edge noise 

spectra in Figure 16 shows a sharp peak between 1 kHz 

and 2 kHz, indicating laminar boundary layer vortex 

shedding and a boundary layer that was not fully 

tripped. This shows that simply comparing data at the 

same Reynolds number is not sufficient. The character

of the boundary layer, in particular the extent of laminar 

flow, must also be simulated. In fact, the data showed 

[2] that the FX63-137 and S822 with their boundary 

layers fully tripped had no tones and suggested that the 

S822 was really quieter than the FX63-137. 

SUMMARY

Large amounts of high quality data were obtained for 

the airfoils tested in this project. This is attributed to 

precise models and rigorous, mature test methods. One 

consequence of this precision was that trailing edge 

bluntness vortex shedding noise did not materialize 

owing to the extremely thin trailing edge of the models. 

We believe model precision also contributed to the 

good agreement obtained for the spectral characteristics 

of the NACA0012 airfoil tested at NASA and compared 

to the NLR results, although a discrepancy in sound 

power level at 1 kHz is still being investigated. 

In quiescent inflow, trailing edge noise is dominant. 

Test results suggest that untripped airfoils operating at 

low Reynolds numbers (< 1 million) can be expected to 

exhibit pure tones at some angles of attack. For example, it was common to observe tones of 10 dB – 15 dB 

above the broadband level. This was attributed to laminar boundary layer vortex shedding. It was also observed that proper tripping eliminated pure tones and 

reduced broadband noise. Even those airfoils that did 

not exhibit pure tones experienced a reduction in sound 

power level of up to 3 dB(A) when tripped. In highly 

turbulent inflow, pure tones disappeared, probably due 

to the suppression of the laminar boundary layer and/or 

disruption of the feedback mechanism responsible for 

laminar boundary layer vortex shedding noise. 

 

The tested airfoils exhibited notably different turbulent 

boundary layer trailing edge noise levels. This source, 

which is likely to be dominant for typical wind turbines, 

differed by as much as 8 dB(A) among the airfoils 

tested. Leading edge inflow turbulence noise becomes 

the dominant source, masking trailing edge noise, in the 

presence of severe upstream turbulence. There appears 

to be a trend of increasing inflow turbulence noise with 

increasing airfoil leading edge sharpness.4

In considering the noise of different airfoils, it may not 

be appropriate to compare at the same angle of attack. 

It is more important to compare at the angle of attack 

expected at the design condition, which may vary significantly from one airfoil to another. 

 

ACKNOWLEDGEMENTS

A lengthy and ambitious project such as this has many 

contributors. In our case, credit must first be given to 

the extraordinary craftsmanship of Yvan Tinel of 

Northbrook, Illinois, who fabricated the airfoil models 

to exacting specifications. Michael Selig of the University of Illinois at Urbana-Champaign performed the 

coordinate measurements and plotting. The test engineers and technicians at NLR deserve special mention. 

They operate a superb facility and exhibit great professionalism and a commitment to quality. 

REFERENCES

[1] Selig, M. and McGranahan, B., 2003, Wind Tunnel 

Aerodynamic Tests of Six Airfoils for Use on Small 

Wind Turbines, NREL SR-500-34515. 

[2] Oerlemans, S., 2003, Wind Tunnel Aeroacoustic 

Tests of Six Airfoils for Use on Small Wind Turbines, NREL SR-500-34470. 

[3] Moriarty, P. and Migliore, P, 2003, Semi-empirical 

Aeroacoustic Noise Prediction Code for Wind Turbines, NREL TP-500-34478. 

[4] Wagner, S., Bareiß, R. and Guidati, G, 1996, Wind 

Turbine Noise, Springer-Verlag, Berlin, pp. 14-21. 

[5] Brooks, T., Pope, D. and Marcolini, M., 1989, Airfoil Self-Noise and Prediction, NASA Reference 

Publication 1218. 

[6] Amiet, R., 1978, “Refraction of Sound by a Shear 

Layer,” Journal of Sound and Vibration, Vol. 58, 

No. 2, pp 467-482. 

[7] Oerlemans and Sijtsma, 2002, “Determination of 

Absolute Levels from Phased Array Measurements 

Using Spatial Source Coherence,” AIAA-2002-

2464.

 

4

 The turbulence intensity in the NLR tests was greater than 

anticipated in the atmosphere because of differences in mean 

velocity between tunnel speeds and typical blade tip speeds. 

Although evidence is not seen in the literature, researchers 

generally agree that leading edge noise is not significant for 

typical large wind turbines. This may not be true for small 

turbines having thin airfoils with sharp leading edges, especially downwind rotors operating in a tower wake. 
9

Figure 3. NLR anechoic wind tunnel set up with acoustically lined endplates and microphone array 

 Turbulence grid off Turbulence grid on 

Airfoil Trip No trip Trip No trip 

S822 0.20/ 0.35/ 0.50/ 0.75/ 1.0 0.20/ 0.35/ 0.50/ 0.75/ 1.0 0.20/ 0.35/ 0.50/ 0.75/ 1.0 0.20/ 0.35/ 0.50/ 0.75/ 1.0 

S834 0.20/ 0.35/ 0.50/ 0.75 0.20/ 0.35/ 0.50/ 0.75 0.20/ 0.35/ 0.50/ 0.75 0.20/ 0.35/ 0.50/ 0.75 

FX 63-137 0.20/ 0.35/ 0.50/ 0.75 0.20/ 0.35/ 0.50/ 0.75 0.20/ 0.35/ 0.50/ 0.75 0.20/ 0.35/ 0.50/ 0.75 

SG 6043 0.11/ 0.20/ 0.35/ 0.50 0.11/ 0.20/ 0.35/ 0.50 0.11/ 0.20/ 0.35/ 0.50 0.5 

SH 3055 0.50/ 0.75/ 1.0 0.50/ 0.75/ 1.0 0.50/ 1.0 0.50/ 1.0 

SD 2030 0.20/ 0.35/ 0.50 0.20/ 0.35/ 0.50 0.20/ 0.50 0.20/ 0.50 

NACA0012 0.50/ 0.62/ 0.87/ 1.12 0.50/ 0.62/ 0.87/ 1.12 0.50/ 0.62/ 0.87/ 1.12 0.50/ 0.62/ 0.87/ 1.12 

Table 1. Test matrix showing measured Reynolds numbers (in millions) with the array on the suction side 

of the model. For the six small-wind-turbine airfoils, all Reynolds numbers were tested at geometrical angles of attack of 0°, 10° and 18°, except the shaded boxes (0°, 5° and 10°). The NACA 0012 airfoil was 

tested at geometrical angles of attack of 0°, 4.5°, 9.0°, 12.0° and 16.5°, to obtain the same effective angles of attack (0°, 2°, 4°, 5.3° and 7.3°) as in Reference 2 (0°, 2°, 4°, 5.3° and 7.3°). 

Figure 4. Acoustic source plots for the untripped NACA 0012 airfoil at 39.6 m/s and α=0° (array on pressure side) illustrating prominent trailing-edge emissions. Flow direction is from left to right. 
10

 

Figure 5: Acoustic source plot (left) indicating noise source locations in the plane of the model. The model 

contour is indicated by the (black) vertical rectangle: flow goes from left to right. The (pink) horizontal rectangle indicates the trailing edge integration contour used to translate acoustic source plots to airfoil noise 

spectra. For measurements with the turbulence grid, a leading edge integration contour was used. 

Figure 6: Acoustic source plots for tripped S822 airfoil at 47.9 m/s and α=0° (array on suction side). Note 

the extraneous “corner sources” in contrast to the uniform sources shown in Figure 4. 

Figure 7. Trailing edge noise spectra for the S822 airfoil (array on suction side) plotted versus frequency 

in Hz. __ 22.4 m/s; _._ 32.0 m/s; ... 47.9 m/s; _ _ 63.9 m/s. * indicates a trip thickness of 0.5 mm. 

PWL

(dB) 

Frequency
11

Figure 8. Trailing edge noise spectra for S822 airfoil at 32 m/s. __ array on pressure side; _._ array on 

suction side. Note the symmetry about the chord, suggesting uniform directivity. 

Figure 9. Leading edge noise from SD 2030 airfoil with trip, with turbulence grid, at a tunnel speed of 

32.0 m/s and α=18° (array on suction side). Note the dominance of the leading edge sources. 

Figure 10. Leading edge noise spectra for S822 airfoil (array on suction side). __ 22.4 m/s; _._ 32.0 m/s; 

... 47.9 m/s; _ _ 63.9 m/s. Note that the levels are much higher than the trailing edge noise spectra 

shown in Figure 7. 
12

Figure 11. Normalized trailing edge noise spectra for the S822 airfoil (array on suction side) plotted versus Strouhal number, St = f · c ÷ U. __ 22.4 m/s; _._ 32.0 m/s; ... 47.9 m/s; _ _ 63.9 m/s. 

Figure 12. Normalized inflow-turbulence noise spectra for tripped S822 (left) and S834 (right) airfoils 

(array on suction side). __ 22.4 m/s; _._ 32.0 m/s; ... 47.9 m/s; _ _ 63.9 m/s. 

Figure 13. Trailing edge noise spectra for the NACA 0012 airfoil at 55.5 m/s. __ NLR data (array on suction side); _._ NASA data [4]. 
13

Figure 14. Narrowband trailing edge noise spectra for three untripped airfoils that showed intense tones 

(U=22.4 m/s; α=10° for S834, α=0° for SG 6043 and SD 2030). 

Figure 15. Narrowband trailing edge noise spectra for S834 airfoil at 22.4 m/s and α=10° as a function of 

trip thickness on pressure side (PS) and suction side (SS). A thin line indicates that these spectral values 

are an upper limit for the trailing-edge noise level. 

20

30

40

50

60

70

0 2000 4000 6000 8000

Frequency (Hz)

PWL (dB)

S834

SG 6043

SD 2030

10

20

30

40

50

60

70

0 2000 4000 6000 8000

Frequency (Hz)

PWL (dB)

0.00 / 0.00

0.25 / 0.25

0.30 / 0.00

0.30 / 0.30

PS / SS
14

Figure 16. Trailing edge noise spectra for S822 airfoil (array on suction side). __ 22.4 m/s; _._ 32.0 m/s; 

... 47.9 m/s; _ _ 63.9 m/s. * indicates a trip thickness of 0.5 mm. 
15

 Figure 17. A-weighted overall trailing edge noise levels at 32 m/s.

Figure 18. A-weighted overall leading edge noise levels at 32 m/s. Airfoils are presented from left to right 

in the order of decreasing leading edge radius and increasing leading edge inflow turbulence noise. 

50

55

60

65

70

75

80

85

90

S822

S834

FX63-137

SG6043

SD2030

SH3055

NACA0012

PWL (dBA)

alpha=0, trip

alpha=0, no trip

alpha=10, trip

alpha=10, no trip

alpha=18, trip

alpha=18, no trip

50

55

60

65

70

75

80

85

90

SH3055

S822

S834

FX63-137

SG6043

SD2030

PWL (dBA)

alpha=0

alpha=10

alpha=18
REPORT DOCUMENTATION PAGE Form Approved 

 OMB NO. 0704-0188 

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson 

Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503. 

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE

December 2003 

3. REPORT TYPE AND DATES COVERED 

Conference paper 

4. TITLE AND SUBTITLE

Wind Tunnel Aeroacoustic Tests of Six Airfoils for Use on Small Wind Turbines: 

Preprint 

6. AUTHOR(S) 

P. Migliore and S. Oerlemans 

5. FUNDING NUMBERS

WER3.1830 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 

National Renewable Energy Laboratory 

1617 Cole Blvd. 

Golden, CO 80401-3393

8. PERFORMING ORGANIZATION 

REPORT NUMBER 

NREL/CP-500-35090

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING 

AGENCY REPORT NUMBER

11. SUPPLEMENTARY NOTES

12a. DISTRIBUTION/AVAILABILITY STATEMENT

National Technical Information Service 

U.S. Department of Commerce 

5285 Port Royal Road 

 Springfield, VA 22161

12b. DISTRIBUTION CODE

 

13. ABSTRACT (Maximum 200 words)

Aeroacoustic tests of seven airfoils were performed in an open jet anechoic wind tunnel. Six of the airfoils are 

candidates for use on small wind turbines operating at low Reynolds number. One airfoil was tested for comparison 

to benchmark data. Tests were conducted with and without boundary layer tripping. In some cases a turbulence grid 

was placed upstream in the test section to investigate inflow turbulence noise. An array of 48 microphones was 

used to locate noise sources and separate airfoil noise from extraneous tunnel noise. Trailing edge noise was 

dominant for all airfoils in clean tunnel flow. With the boundary layer untripped, several airfoils exhibited pure 

tones that disappeared after proper tripping was applied. In the presence of inflow turbulence, leading edge noise 

was dominant for all airfoils.

15. NUMBER OF PAGES

14. SUBJECT TERMS 

airfoils; small wind turbines; aeroacoustic tests 16. PRICE CODE

17. SECURITY CLASSIFICATION 

OF REPORT

Unclassified 

18. SECURITY CLASSIFICATION 

OF THIS PAGE

Unclassified

19. SECURITY CLASSIFICATION 

OF ABSTRACT

Unclassified

20. LIMITATION OF ABSTRACT

UL 

 NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89) 

Prescribed by ANSI Std. Z39-18 

298-102