Trailing-Edge Noise Measurements using a Hot-Wire

Based Coherent Particle Velocity Method

A. Herrig∗, W. Wurz†, Th. Lutz‡ and E. Kramer§

University of Stuttgart, D-70550 Stuttgart, Germany

This paper describes a hot-wire based method for two-dimensional trailing edge noisemeasurements mainly for use in wind tunnels with high background noise levels. Themethod is based on the cross correlation of two hot-wire signals, allowing to measure theCoherent Particle Velocity (CPV) of the emitted sound waves. The strong directionalsensitivity of the hot-wires leads to a suppression of parasitic noise, which significantlyimproves the signal-to-noise ratio. Due to the small dimensions of the sensors together withthe assured laminar flow condition they are well suited for inflow measurements. To obtainquantitative results in terms of sound pressure level, the sensitivity of the measurementsetup is derived by simulation of the response of the hot-wires to a line source. Forvalidation, trailing edge noise measurements were performed at 60 m/s and a Reynoldsnumber of 1.6×10

6 in the closed test section of the Laminar Wind Tunnel Stuttgart (LWT),using the CPV-method. Finally, the same airfoils were investigated in the open jet of theAeroacoustic Wind Tunnel Braunschweig (AWB) using a phased microphone array.

The quantitative comparison of the experimental results obtained in the two wind tun-nels required the application of appropriate wind tunnel corrections. The obtained soundpressure frequency spectra are basically found to be parallel in the frequency range ofsufficient measurement accuracy. The total sound pressure levels vs. lift coefficient showa more or less constant offset of about 2 dB between AWB and LWT. Given the totallydifferent measurement principles this can be regarded as a good agreement. Finally resultsof a NACA 0012 airfoil are presented and compared to published data.

Nomenclature

α angle-of-attackαg geometric angle-of-attackρ0 air density, kg/m3

A overheat ratio Tw/Ta, -c chord length, mc0 sound velocity, m/sE hot-wire bridge output voltage, VL wetted length of trailing edge, mLp sound pressure level, dB re 20µPaMa Mach number, -

Tux longitudinal turbulence level√

u′2/U∞, -U, U∞ velocity at the hot-wire, in the free stream, m/sv′ particle velocity in the sound waves, m/s

∗Research engineer, Institute of Aerodynamics and Gas Dynamics (IAG), Pfaffenwaldring 21, AIAA Member.†Senior Researcher, Institute of Aerodynamics and Gas Dynamics (IAG), Pfaffenwaldring 21.‡Senior Researcher, Institute of Aerodynamics and Gas Dynamics (IAG), Pfaffenwaldring 21, AIAA Member.§Professor, Head of the Institute of Aerodynamics and Gas Dynamics (IAG), Pfaffenwaldring 21.

1 of 12

American Institute of Aeronautics and Astronautics

24th Applied Aerodynamics Conference5 - 8 June 2006, San Francisco, California

AIAA 2006-3876

Copyright © 2006 by Andreas Herrig. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

I. Introduction

Turbulent Boundary-Layer Trailing-Edge (TBL-TE) interaction noise is the dominant far-field noiseradiated from wind turbines for typical inflow conditions. Wind park acceptance and growingly strict

regulations concerning noise emissions require a reduction of this noise. The aim is therefore to develop newoptimized airfoil contours, which provide a reduction of total sound pressure levels, but must not suffer fromdecreased aerodynamic performance.

Before application of the airfoils on wind turbines the aerodynamic and aeroacoustic characteristics haveto be verified in two-dimensional wind tunnel tests. Typically, acoustic measurements are performed in opentest section wind tunnels with a surrounding anechoic room. Acoustic damping measures can be appliedto nearly achieve free-field conditions without disturbing reflections of sound waves. Phased microphonearray systems are the established method,1, 2 because the processing technique provides the possibility tolocate noise sources and suppress background noise. Furthermore, a gain in signal-to-noise ratio (SNR) canbe achieved by the large number of sensors. Elliptic mirrors are also applied successfully in open-jet windtunnels.3

Open-jet wind tunnels are not suited for aerodynamic measurements, on the other hand. Measurementsfor the verification of aerodynamic performance must be performed in low turbulence closed test section windtunnels. So a drawback of the combined aero-acoustic verification is that two wind tunnel campaigns arenecessary. The comparison of the obtained data is also complicated by the different aerodynamic boundaryconditions (open-jet effect, turbulence level). Therefore, it is desirable to develop methods for acousticmeasurements in closed test section wind tunnels with typically high background noise levels. Wall-mountedarrays4 and in-flow mounted microphones arrays5, 6 are commonly used for this task. At the Institute ofAerodynamics and Gas Dynamics (IAG) of the University of Stuttgart a new hot-wire based method forthe in-flow measurement of trailing edge noise has been developed.7 Hot-wires are used to measure theparticle velocity of the sound waves. Cross-correlation is used to extract coherent signals – Coherent ParticleVelocity (CPV) method. The setup is similar to the Coherent Output Power method chosen by Hutchesonand Brooks,8 but has some important advantages, which will be outlined in this paper.

First CPV experiments were performed on a symmetric 4.2% thick ’flat plate like’ airfoil (c = 0.5m) atzero angle of attack and Ma = 0.175.7 This represents some benchmark case due to the low noise emission incomparison to real airfoil sections. Subsequently the CPV-method was successfully applied for the validationof cambered airfoils. Very valuable information for the improvement of IAG’s aeroacoustic prediction codecould be obtained, as it allows the resolution of very small differences in sound pressure level in the orderof 0.5 dB. In addition comparisons of the CPV results to an established aeroacoustic measurement methodwere performed.9

In this paper some experimental trailing edge noise results from the SIROCCO10 project obtained in openjet (microphone array) and closed test section wind tunnel tests (CPV) will be discussed in more detail. Insection II the experimental set-up in both wind tunnels is described, concentrating on the CPV-method. Itis followed by aerodynamic aspects of the measurements. In section III the acoustic results are discussedand compared to each other. Finally recently obtained data of a NACA 0012 model is shown and comparedto published benchmark data.

II. Experimental setup

A. Coherent particle velocity method

1. LWT wind tunnel environment

The acoustic CPV-measurements were carried out in the Laminar Wind Tunnel (LWT)11 of the IAG. It isof Eiffel type (Fig. 1) and has a maximum velocity of 90m/s. The closed rectangular test section measures2.73 × 0.73m2 and is 3.15m long. The 2D airfoil models vertically span the short distance of the testsection and the gaps to the walls are sealed. The high contraction ratio of 100:1 (referred to the total inletarea) as well as five screens and filters result in a very low turbulence level. This makes it well-suited forlaminar boundary layer investigations and research on Natural Laminar Flow (NLF) airfoils. A longitudinalturbulence level of about Tux = 0.02% could be determined with hot-wire measurements in the frequencyrange of 20-5000Hz at U∞ = 30m/s.12 In the frequency range below 250Hz the dominating part of thedisturbances is of acoustic nature. This could be shown by correlation measurements13 with two hot-wires

2 of 12

American Institute of Aeronautics and Astronautics

Figure 1. Side view of the Laminar Wind Tunnel, total length 46m.

traversed transversally. For frequencies of interest (> 50Hz) the turbulent longitudinal velocity fluctuationsare correlated in the cross-flow direction over less than 15mm.a

Despite the LWT is very well suited for aerodynamic measurements, it is not optimal for acoustic mea-surements. The fan is located 12m downstream in the diffuser and in straight line to the test section. Upto now, no acoustic damping measures are applied. Due to the area relation between the location of the fanand the test section of 5.7:2 the total sound pressure in the empty test section of the LWT is rather high.An A-weighted sound pressure level of Lp = 94dBA at 60m/s flow velocity was measured with 1/2” Bruel& Kjær free-field microphones type 4190 with a nose cone.

This is comparable to other closed test section wind tunnels (Fig. 2). The velocity dependence shows alower slope though, which might be caused by the large ratio of fan area to test section area. The backgroundnoise has broadband character and most of the energy is concentrated in the low frequency range (Fig. 3).Special methods are required to measure airfoil trailing edge noise under such conditions, because it mustbe separated from the high background noise.

U [m/s]

L p[d

BA

]

0 10 20 30 40 50 60 7050

60

70

80

90

100

110

120

LWTDNW-LLF 8x6m², closed jetDNW-LLF 8x6m², acoustic open jetIVK tunnel Stuttgart, open jet

Figure 2. Measured A-weighted total background noiselevel of LWT compared to other European wind tunnels(data from14).

f [Hz]

Lp

[dB

re2e

-5P

a]

102 103 10450

60

70

80

90

100 20 m/s30 m/s40 m/s50 m/s60 m/s70 m/s

Figure 3. Unweighted 3rd-octave sound pressure spec-tra of LWT background noise for different velocities.

2. CPV hot-wire sensors

The setup for the CPV-method is derived from Hutcheson & Brooks8 by replacing the microphones withhot-wires. Two 45 degree slanted hot-wires (Dantec P15 probes, ø2.5µm×1.4mm platinum-plated tungstenwires) are placed in the x-y-plane (perpendicular to the trailing edge, sketched in Fig. 4), with their axis inthe same plane. Typically the distance is a = ±75mm and b = −75mm, which is many times larger thanthe boundary layer thickness for a typical chord length of 0.4-0.6m. Due to the large distance of the wiresand their separation by the airfoil section, the influence of coherent parts in the signals resulting from windtunnel turbulence and turbulent boundary layer (TBL) vortices is significantly reduced. Remaining coherentsignals are therefore mainly due to acoustic particle velocity fluctuations.

aThis could not be quantified exactly, as the measurement of these very low fluctuations is at the limit of the signal to noiseratio and due to the presence of the large acoustic contributions.

3 of 12

American Institute of Aeronautics and Astronautics

The strongly non-isotropic directional sensitivity of the wires is exploited to improve the signal-to-noiseratio (SNR) further, which is the main advantage of using hot-wire sensors. For a single yaw probe, thepolar distribution of the sensitivity in the plane of the wire can be described by a yaw function f(α), anextended cosine law is used

f(α) =√

cos2 α + k2 sin2 α , (1)

with α being the angle between the incident wave vector and the wire normal and k expressing the sensitivityin the direction of the wire. For standard hot-wires k is about 0.2, for the thin wires with l/d ≈ 600 it can besmaller.15 So noise approaching from the direction of the wire axis is damped by at least 14 dB. The -3 dBview angle is 45◦ which is not particularily good, but still a large amount of parasitic noise is rejected muchmore efficiently than when using omnidirectional microphones.

In the out-of-plane direction it turns out that due to the presence of a large mean velocity vector and onlyvery small fluctuating velocities (v′/U → 0) a cosine law is also found for the directivity in the y-z-plane.

So the wires are positioned such that the reception of the airfoil noise is maximized, but backgroundnoise approaching from downstream is damped by 3 dB. TE corner noise sources, often posing problems toacoustic measurements,16 are damped by at least 11.2 dB for the test conditions at the LWT.

Figure 4. Schematic arrangement of the CPV-system.

Figure 5. The CPV-system mounted in the LWT test section downstream of a wind tunnel model of 0.6mchord.

The probe supports are mounted approximately 0.6m downstream of the wind tunnel model in such away that its wake can pass between them, avoiding extra turbulent inflow noise from the supports (Fig. 5).

The hot-wires are driven by very low-noise Dantec 55M10 CTA-bridges, which are adjusted to deliver a flatfrequency response up to about 80-120kHz (compare square wave test15 result in Fig. 6). So a measurement

4 of 12

American Institute of Aeronautics and Astronautics

of the frequency response of the sensors can be avoided. The wires are operated with an overheat ratio ofA = 1.8. In principle the fluctuating voltage signal is composed of contributions from density, temperatureand pressure fluctuations

E′ = Euu′[

1 +Eρρ

′

Euu′+

ET T ′

Euu′+

Epp′

Euu′

]

. (2)

According to Davis17 it is possible to choose a wire temperature so that the three terms on the rightcancel each other and the velocity fluctuations indicated by the wire correspond to the true particle velocityfluctuations in a plane acoustic wave. For velocities of 60m/s this results in an overheat ratio of about 1.7.Unfortunately, Davis presented only results for a ø5µm wire and there is no direct relation for transferringthe result to a 2.5µm×1.4mm wire. The main unknown is the change of the sensitivity to pressure, whichis influenced by the Knudsen number Kn being the ratio of molecular mean free path to wire diameter.With the larger Kn of the thinner wires the pressure term is expected to rise, which would mean that themeasured velocity fluctuations are somewhat too high. So in the experiments A might be chosen too large,but it reduces the sensitivity to temperature fluctuations from the environmental air on the other hand. Theinfluence of a phase shift between u′ and p′ in the acoustic near field still is to be investigated.

The calibration of the hot-wires is done in-situ by variation of the tunnel speed U∞ and approximation of∂U/∂E from the hot-wire mean voltage changes ∆E recorded simultaneously. The derivatives (linearizationof King’s law) are evaluated from fitted polynomials of second order. Fig. 7 shows an example. Thisprocedure improved the accuracy in ∂U/∂E to about 0.1-0.2 dB.

The influence of the airfoil velocity field is taken into account. The ratio of potential velocity at thehot-wire position to freestream U/U∞ is obtained with XFOIL to also capture the effect of the boundarylayer displacement thickness and possible turbulent separation.

Figure 6. Step response of hot-wires indicating a cor-ner frequency of about 130 kHz.

−0.03 −0.02 −0.01 0 0.01 0.02 0.0356

57

58

59

60

61

62

63

64

files 305 − 373, tK = 284.8, p

K = 727.1, φ = 66%

∆ E [V]

U [m

/s]

dU/dE1 = 121.46 m/sV−1

dU/dE2 = 122.67 m/sV−1

HW 1 HW 2 approx. 1approx. 2

Figure 7. Calibration of the wires by variation of hot-wire voltage with mean velocity.

Further advantages of the wires are that the flow over the sensors is always laminar, independent of thelocal flow direction and the disturbance of the flow field due to the sensors is very small.

3. Data acquisition and processing

The signals of the hot-wire bridges are AC-amplified by AMI-321A ultra low-noise amplifiers (1 nV/√

Hzeqv. input noise) with a gain of 1,000. The 200Hz high-pass filter characteristic was used to improve thedynamic range. Final AD-conversion is done by a 24bit audio-system (RME Hammerfall DSP) at a samplingrate of 44.1 kHz per channel. The Σ∆-converters with 64 times over-sampling provide excellent anti-aliasingfiltering at half the data rate. Additionally one-pole RC low-pass filters at 15 kHz remove excessive high-frequency noise. Typically, time traces of 10min are recorded and then processed by Fast Fourier Transformsof blocks of 4096 points yielding a frequency resolution of 10.78Hz. The cross spectrum G12 is calculatedfrom the Fourier coefficients of the two simultaneous streams 1, 2 and averaged over the whole record length

5 of 12

American Institute of Aeronautics and Astronautics

of n ≈ 6500 blocks:18

G12(f) =1

n

n∑

k=1

[

X∗

1,k(f)X2,k(f)]

(3)

The phase differenceϕ(f) = arctan(Im(G12(f))/Re(G12(f))) (4)

is obtained from the cross spectrum. This is a crucial information for selecting the frequency range where theairfoil noise is sufficiently higher than the background noise and can therefore be measured. Sound pressure(and particle velocity) radiated from the TE is of dipole type, leading to 180◦ phase shift with a symmetricsetup, while the tunnel background noise approaching from downstream shows up with approx. 0◦ phase.

Figure 8. Coordinate system insimulation of the line source andretarded position of the wires.

For obtaining quantitative far-field values of the sound pressure, asimulation of the response of the whole CPV system to the TE line sourceis performed. Incoherent point sources are uniformly distributed alongthe trailing edge and their particle velocity contributions at the positionof the wires are added up. This way the non-isotropic sensor response, thedistance scaling and the source directivity is corrected. Sound pressure p′

is related to particle velocity v′ by the radial impedance p′ = ZRv′, whichis given by

ZR = ρ0c0

1

1 − ic0

ωr

for monopole and ZR = ρ0c0

ic0

ωr− 1

2 c0

ωr+ 2i c0

ωr− 1

(5)

for a dipole source. For the present investigations monopoles were chosenfor the evaluation, as it could not be verified experimentally that thedistance scaling corresponds to a pure dipole law. The convection of thesound waves causes a retardation of the effective hot-wire position, which is also taken into account (compareFig. 8). The results are finally given as the sound pressure level Lp produced by a trailing edge of L = 1 mat a distance of r = 1m and an observer placed at a reference angle of 90◦ to the airfoil chord.

B. Aeroacoustic Wind Tunnel Braunschweig

Acoustic verification measurements were performed by comparing CPV measurements in the LWT to mi-crophone array measurements in the Aeroacoustic Wind Tunnel Braunschweig (AWB). The maximum flowvelocity is about 60m/s, the turbulence level is significantly higher than in the LWT, but not exactly spec-ified. In contrast to the LWT the AWB is a closed return open-jet tunnel with a rectangular nozzle of1.20× 0.80m2 and a surrounding anechoic test chamber. Besides the low background noise level this makesit possible to achieve nearly free-field conditions without disturbing reflections of sound waves.

The same wind tunnel models as in the LWT were used. They were mounted horizontally between twoendplates coated with open cell foam for reducing acoustic reflections. Fig. 9 shows the nozzle of the AWBwith the acoustic array of the National Aerospace Laboratory (NLR) above the model (standard position).The 1 m diameter array consists of 96 microphones in an open metal grid. The phased array processingis similar to19 and results in 1/3-octave band spectra of the trailing edge noise radiated from the central0.2m of the model span. Special measures were taken to physically reduce extraneous noise sources atthe model-endplate junctions. In some cases, noise from these corners influenced the measured TE noiselevels. Therefore, a routine was used which automatically determines the importance of these ’corner sources’and which, in case the influence of the corner sources on the trailing edge noise level is more than 0.5 dB,calculates an upper limit for the actual 2D trailing edge noise level.

Prior to the array measurements a lift balance was used for the determination of lift curves and sub-sequently of reference angles of attack. Transition location measurements using a stethoscope and oil flowvisualizations were performed similar to the investigations at the LWT.

C. Wind tunnel corrections

It is commonly accepted that comparisons of airfoil TE noise spectra should be performed on the basisof equal lift coefficients. Wind tunnel corrections are therfore important. In the LWT the aerodynamiccoefficients are by default corrected for streamline curvature (SC) and solid blockage to represent values that

6 of 12

American Institute of Aeronautics and Astronautics

Figure 9. Test section of the Aeroacoustic Wind Tunnel Braunschweig (AWB) with acoustic array installedabove the model.

would be obtained in an infinite flow field (IFF). The corrections are in the order of 1-2% for typical modelchord lengths. For open-jet wind tunnels much larger SC corrections are necessary, because the air stream isdeflected by the airfoil. Due to the latter the effective airfoil shape changes quite significantly (compare20)and is different for different lift coefficients. Depending on the ratio of chord length and jet dimensions, thesame lift coefficient cl,t as in the IFF is reached at a much higher geometric angle-of-attack (AOA) αg.

To illustrate the problem Fig. 10 shows lift curves of one of the airfoils measured in LWT and AWB(symbols). It is apparent, that the slope is reduced to half the IFF value in the AWB. In addition, for equallift coefficient the transition location in the AWB and turbulent separation were occurring significantly moreupstream on the upper surface.b

Normally correction terms are applied only to αg to obtain the AOA of equivalent lift α in IFF (comparee.g.2). Fuglsang et al.21, 22 chose an AOA correction based on the cl and cm and also taking the downwashinto account23

α = αg −√

3σ

πcl −

2σ

πcl −

σ

π(4cm) , (6)

where σ = π2

48

(

cH

)2with H the tunnel height. For cambered airfoils this usually provides much better

agreement than using α = αg/[(1 + 2σ)2 +√

12σ] like in2 as the zero-lift angle is taken into account.Using (6) with the AWB dimensions, the slope of the corrected AWB lift curve is somewhat too low.

It has to be noted, that the aeroacoustic behaviour is strongly influenced by the boundary layer devel-opment and therefore by even small details in the cp distribution. The methods mentioned above howeverneglect the fact that the cp distributions are not similar for equal lift coefficient. In order to obtain a goodcomparison between the two wind tunnel results, investigations were performed to match the cp pressuredistributions as close as possible.

Therefore the influence of the open jet was investigated in more detail using MSES.24 MSES solves thecoupled Euler and boundary layer equations by a Newton scheme. Constant pressure along the upper andlower grid boundaries can be forced to represent a free jet. The transverse grid dimensions were chosen as inthe experiment. The resulting lift curve (Fig. 10, orange) corresponds to the AWB already quite good, butthere is still a too steep lift slope. The reason is the induction of a downwash at the airfoil position. Due

bThe influence of the higher turbulence level on transition was assumed to be relatively small, because close to cL = 0 thetransition locations were found to be quite similar.

7 of 12

American Institute of Aeronautics and Astronautics

αg

CL

0

0.5

1

LWT exp.AWB exp.MSES, infinite flow fieldMSES, open-jetMSES, open-jet+downw.AWB exp., corr. α Eqn. 6

Re = 1.6e6, clean

5°

Figure 10. Comparison of lift curves measured in LWT(corrected to infinite flow field already) and AWB aswell as MSES calculations.

Figure 11. Vortex system for calculation of lift slope.The wake vortices begin downstream the endplates.

to the deflected jet, vortices develop downstream of the finite end plates of the AWB setup, not consideredin the 2D MSES simulations. This changes the effective angle of attack and reduces the lift. An equationfor the lift slope of a ’clipped’ horseshoe vortex (Fig. 11) was derived based on the principle of Pistolesi.25

It was assumed, that the shear between the jet and the surrounding air corresponds to vortices havingthe same circulation as the bound vortex at quarter chord and that they start at the trailing edge of theendplates. Using this approximation, a downwash correction factor for cl can be approximated and the liftcurve matches the AWB experiments sufficiently well (Fig. 10, red line).

When comparing the MSES calculated pressure distributions of open-jet and IFF it is found, that com-plete equivalence cannot be achieved with the same airfoil contour, even for only one angle of attack. Dueto the reduced effective camber and thickness a suction peak forms ealier in open-jet. But the slopes ∂cp/∂xagree fairly well for a reduction of the IFF-cl of about ∆cl/cl = 0.17. The calculated open-jet transitioncurves almost correspond to the IFF curves scaled down the same amount. The observed downscaling of 25%cl between the experimental xtr − cl-curves of AWB compared to LWT expresses the effect of the additionaldownwash and corresponds to the result when applying all derived corrections.

Thus to summarize, when the trailing edge noise results from the AWB are compared to those from theLWT, lower lift coefficients have to be used for the AWB, to account for the open-jet effect on the pressuredistribution and boundary layer. Using a correct reduction, the pressure gradients are nearly the same andthe transition locations are similar in both tunnels. This results in boundary layer states at the trailing edge,which correspond to each other reasonably. It should be noted that in the nonlinear lift range near cl,max

the accuracy is only limited and more detailed investigations are necessary then.

III. Results and discussion

First results presented here were obtained for airfoils for wind turbine application measured in the frame-work of the SIROCCO project. The wind tunnel models have a chord length of c = 0.4m, the TE thicknesswas 0.3mm. Tests were performed at a velocity of U∞ = 60m/s. All airfoils showed smooth spectra withoutblunt trailing edge (BTE) noise and laminar boundary-layer vortex-shedding noise. Fig. 12 shows a typicalplot of the raw data obtained at a lift coefficient of cl = 1.0 (upper plot) serving to discuss the main aspects

8 of 12

American Institute of Aeronautics and Astronautics

of the CPV method here.c With reduction of frequency the levels rise strongly and the airfoil noise getsmasked by the background noise from the tunnel. At high frequencies above 3 kHz there is an increase ofelectronic noise because of hot-wire thermal noise amplified by the bridge amplifier. The signals are gettingmore incoherent here, which can be seen from the level difference between the single spectra (G1, G2) andthe cross spectrum G12, as well as the random phase.

At least four types of ’noise’ contribute to the measured signal: acoustic noise from the airfoil, acousticbackground noise, electronic noise and turbulent velocity fluctuations. Completely incoherent noise is reducedby the averaging procedure proportional to 1/

√n, with n the number of averages (typically 6500). Assuming

that all parasitic noise is incoherent, a signal-to-noise ratio can be approximated according to SNR =10 · lg(γ

√n) with γ2 = G2

12/G1G2 being the coherence squared (lower plot in Fig. 12). In the plot of thephase difference between the two sensor signals four ranges can be identified:a) Up to about 300Hz the phase tends to zero. Coherent background noise from downstream dominates.b) From 300 to 600Hz tunnel background noise seems to be higher than the airfoil noise. The phase showssome fluctuations, but the SNR is high as indicated in Fig. 12. Reflected sound waves and/or boundarylayer vortices seem to cause this behaviour and the mechanism is not fully understood, yet.c) Above 600Hz, the TE noise dominates the signals and can be measured reliably. The phase shows theexpected 180◦ behaviour, the small change to higher frequencies is probably caused by a time delay of noisedue to a slight asymmetry of the setup and a higher convectional speed on the suction side.d) The upper bound of the useful measurement range is given by the loss of SNR at about 3 kHz. In principle,it is possible to increase this upper bound by significantly longer averaging. In contrast the lower bounddoes not change then, it depends on the setup and the coherent background noise of the tunnel.

From the narrow-band cross spectra third-octave spectra of sound pressure level are calculated by ener-getic summation. Based on the phase distribution a criterion was established to ensure that only valid datapoints, which fit to c), are accepted.

A. Comparison of LWT and AWB measurements

For comparison of the results of both tunnels, the AWB microphone array data are converted to LWT resultsfor convenience. As the AWB results are given as sound power levels PWL (equals the monopole soundpressure level in a distance of 1/

√4π m) for a trailing edge segment of 0.2m length, the re-scaling is as

follows:

Lp = PWL − 20 lgr2

r1

+ 10 lgL2

L1

+ 50 lgU2

U1

= PWL − 3.26 dB . (7)

The TBL-TE noise is incoherent, therefore Lp ∼√

L. The velocity term is to correct the small differencein tunnel velocity (AWB U1 = 58m/s, LWT U2 = 60m/s). From theory a Ma5 dependence is expected forthe non-compact frequency range,26 but as the boundary layer properties depend on the Reynolds number,lower values are often found in practice. Nevertheless, for the AWB tests an exponent of 5 turned out to bea very good value that matches measured spectra of 50 and 58m/s.

When comparing the third-octave spectra obtained in both wind tunnels (Fig. 13), lower test lift co-efficients for the AWB have been used, to account for the open-jet effect and the wake influence on thepressure distribution and boundary layer (see section II.C). The general similarity of the spectra is good. Inthe high-frequency region where the measurement certainty is good, only small differences exist. In the lowfrequency region the values differ a bit more, which is probably due to effects of the measurement techniques.The CPV values tend to lower values when an additional coherent signal with opposed phase exists (mainlythe fan noise).

The total sound pressure levels of the two airfoils obtained in AWB and LWT are compared in Fig. 14.The values are obtained by summation in the measurable frequency range. The minimum frequency forsummation is indicated at the data points, the upper end was at least 2500Hz. As the peak frequencydepends strongly on the angle of attack, the integration must be started at the appropriate lower frequenciesfor every lift coefficient. When calculating total levels, the inclusion of the third-octave band where the peaklevel occurs is very important to get representative values (this is illustrated in27). The TL-132 is a newairfoil section designed in the framework of the SIROCCO project.28 In comparison to the reference sectiona reduction of 1-1.5 dB is obtained for tripped boundary layer in both wind tunnels.

cThe few distinct peaks in the CPV spectra around f = 600 Hz mainly result from vibrations of the hot-wires struts, whichcould be shown by adding a mass to one of them which reduced the resonance frequencies.

9 of 12

American Institute of Aeronautics and Astronautics

f [Hz]

SN

R[d

B]

1000 3000 5000 7000 90000

5

10

15

20

400

f [Hz]

|G1|

,|G

2|,|

G12

|[(

m/s

)2 /Hz]

arg(

G12

)[r

ad]

1000 3000 5000 7000 900010-13

10-12

10-11

10-10

10-9

10-8

10-7

10-6

-3

-2

-1

0

1

2

3

|G1||G2||G12|arg(G 12)

GAMESA airfoil, cL=1.0, trip 5%

400

Figure 12. Single and cross-spectra of particle velocity (above). The signal-to-noise-ratio (below) togetherwith the phase information indicates the region of reliable values (600-3000Hz).

Frequency [Hz]

L P(r

=1m

)[d

B/m

]

1000 2000 3000 4000 5000

reference, CPV, c l=1.0reference, AWB, c l=0.83TL132, CPV, c l=1.0TL132, AWB, c l=0.83

60 m/s, trip 5%

5 dB

250

Figure 13. Measured third-octave sound pressure spec-tra obtained for the reference and TL132 airfoils inLWT and AWB (scaled to U∞ = 60 m/s according to(7).

cL

LP

(r=

1m

)[d

B/m

]

0.6 0.7 0.8 0.9 1 1.1 1.2 1.3

reference, AWBreference, LWTTL132, AWBTL132, LWT

1000

900

800

630

500

2 dB

Figure 14. Total SPL polars for the tripped referenceand TL132 airfoils. The lower frequency bound forstarting the integration is denoted at the labels. AWBcl’s scaled down to represent IFF values.

B. Measurements on a NACA 0012 airfoil

Recently CPV measurements on a NACA 0012 airfoil of c = 0.4m (original TE thickness) were performed.Results obtained at α = 0◦, U∞ = 60m/s, Re = 1.6× 106 are compared (Fig. 15) to benchmark microphonearray data from NLR’s KAT wind tunnel29 and data measured at NASA.30 The data were converted tothe representation used at the LWT similar to (7). This results in 0.7 dB added to the NLR values and6.14+0.7dB to the NASA values. Then standard scaling laws for U∞ (p′2 ∼ Ma5,f ∼ U∞) were applied tocorrect for the remaining difference in free stream speed. The obtained spectra are very similar, differences arenot much larger than the differences between the two other methods.d The comparison with the theoreticalprediction code developed in Sirocco28 also shows very good agreement of the spectra. The model for blunttrailing edge noise was not activated, so the hump around 4.5 kHz occuring in the experiment for the clean

dUnfortunately only a model with c = 0.4m could be measured up to now, so despite applying standard scaling laws for U∞

there is still the different Reynolds number to be considered. Measurements on models of 0.2m and 0.6m are planned for thefuture.

10 of 12

American Institute of Aeronautics and Astronautics

case is not reproduced. The general quantitative agreement of the CPV results at the present state withwell established acoustic measurement techniques shows the potential of this new method.

f*60/U [Hz]

Lp

+50

lg(U

/60)

500 2500 4500 6500 850030

35

40

45

50

55

60

65

70

75

80

LWT 2006, c=0.4 m, 60 m/sNLR 2004, c=0.2286 m, 55.5 m/sBPM 1989, c=0.2286 m, 55 m/s

clean

f*60/U [Hz]500 2500 4500 6500 8500

30

35

40

45

50

55

60

65

70

75

80

LWT 2006, c=0.4 m, 60 m/sNLR 2004, c=0.2286 m, 55.5 m/sBPM 1989, c=0.2286 m, 55 m/s

tripped

Figure 15. Comparison of measured NACA 0012 3rd-octave spectra at α = 0◦ scaled to U∞ = 60 m/s.

IV. Conclusions

A hot-wire based CPV-method is presented for the measurement of trailing edge noise in ’noisy’ aero-dynamic wind tunnels. Several advantages arise from the present method in comparison to the microphonebased COP method: (i) due to the high directional sensitivity of the hot-wires parasitic noise is rejectedmuch more efficiently in comparison to microphones with an omnidirectional response, (ii) the hot-wires canbe calibrated in situ by a simple change of mean velocity, leading to a high accuracy in the determinationof the sensitivity of the sensors, (iii) the frequency response is flat up to the ultrasonic range, (iv) the flowover the sensors is always laminar, independent of the direction of the local flow vector with respect to thesensor orientation, (v) the disturbance of the flow field due to the sensors is very small.

The experimental procedure, including the necessary theoretical approach for obtaining quantitative noisespectra, is validated by comparison of measured TE noise spectra from wind turbine airfoil sections, developedin the framework of the SIROCCO project. The CPV-tests were performed in the Laminar Wind Tunnel(LWT) in a closed test section and the corresponding microphone array measurements in the open test sectionof the Acoustic Wind Tunnel Braunschweig (AWB). The same wind tunnel models were used at similar flowspeeds of 60m/s. It turned out that delicate wind tunnel corrections are necessary to ensure similar pressuredistributions, and therefore similar boundary layer development. The commonly used assumption, that TE-noise radiated from airfoils can be compared for different facilities if equal lift coefficients are adjusted, ismisleading. Mainly for open test sections with limited size of the end plates, the influence of the curved freestream on the local pressure distribution is significant. In the present case the conditions of the open testsection were modeled using MSES calculations and additional potential methods for the downwash due tovortex induction. The final results obtained show a very good quantitative agreement between the CPV andthe microphone array measurements, for the frequency spectra in the range of 800-2500Hz as well as for thetotal sound pressure level.

Additional CPV measurements were performed on a NACA 0012 section in order to compare the results toestablished data bases which are publically available. Again a good overall agreement is found. Consequently,the CPV-method is being used as a standard technique to perform combined aerodynamic and aeroacoustictests at high Reynolds numbers in the LWT. Further developments are pointed at the investigation of sourcedirectivities and distance scaling laws and the extension of the measurement range to lower frequencies.

V. Acknowledgements

This research project is supported by the European Commission’s Fifth Framework Programme, projectreference: ENK5-CT-2002-00702 SIROCCO, Silent Rotors by Acoustic Optimisation. The authors would

11 of 12

American Institute of Aeronautics and Astronautics

also like to thank the project partners for the permission to publish selected results.

References

1Mueller, T. J., editor, Aeroacoustic Measurements, Experimental Fluid Mechanics, Springer, Berlin Heidelberg New York,2002.

2Hutcheson, F. V. and Brooks, T. F., “Effects of Angle of Attack and Velocity on Trailing Edge Noise,” 42nd AIAAAerospace Sciences Meeting, January 5-8 2004.

3Dobrzynski, W., Nagakura, K., Gehlhar, B., and Buschbaum, A., “Airframe Noise Studies on Wings with DeployedHigh-Lift Devices,” 4th AIAA/CEAS Aeroacoustics Conference, Toulouse, France, June 2-4 1998, AIAA-1998-2337.

4Soderman, P. T. and Allen, C. S., Aeroacoustic Measurements, chap. Microphone Measurements In and Out of Airstream,Experimental Fluid Mechanics, Springer, Berlin Heidelberg New York, 2002, pp. 1–61.

5Holthusen, H., “The Phased Microphone Array Measurement Systems of the German-Dutch Wind Tunnels,” SWING+Workshop, Stuttgart, IAG, 2002.

6Guidati, S., Entwicklung eines akustischen Messsystems fur aerodynamisch optimierte Windkanale, Ph.D. thesis, Uni-versitat Stuttgart, 2003.

7Wurz, W., Guidati, S., Herrig, A., Lutz, T., and Wagner, S., “Measurement of Trailing Edge Noise by a Coherent ParticleVelocimetry Method,” Int. Conference on Methods of Aerophysical Research ICMAR, Novosibirsk, 28.06.-02.07. 2004.

8Hutcheson, F. V. and Brooks, T. F., “Measurement of Trailing Edge Noise Using Directional Array and Coherent PowerMethods,” Proceedings of 8th AIAA/CEAS Aeroacoustics Conference, No. 2002-2472 in AIAA paper, 2002.

9Herrig, A., Wurz, W., Lutz, T., Braun, K., Kramer, E., and Oerlemans, S., “Trailing Edge Noise Measurements ofWind Turbine Airfoils in Open and Closed Test Section Wind Tunnels,” First International Meeting on Wind Turbine Noise:Perspectives for Control, Berlin, Germany, 17.-18.10. 2005.

10Schepers, J. G., Curvers, A., Oerlemans, S., Braun, K., Lutz, T., Herrig, A., Wurz, W., and Lopez, B. M., “SIROCCO:SIlent ROtors by aCoustiC Optimisation,” First International Meeting on Wind Turbine Noise: Perspectives for Control, Berlin,Germany, October 17-18 2005.

11Wortmann, F. X. and Althaus, D., “Der Laminarwindkanal des Instituts fur Aero- und Gasdynamik an der TechnischenHochschule Stuttgart,” Zeitschrift fur Flugwissenschaften, Vol. 12, No. 4, April 1964, pp. 129–134.

12Herr, S., “Messung des Turbulenzgrades im Laminarwindkanal,” Tech. rep., Institut fur Aerodynamik und Gasdynamik,2003.

13Herrig, A. and Wurz, W., “New Measurements of the Turbulence Level in the Laminar Wind Tunnel,” Tech. rep.,Institute of Aerodynamics and Gas Dynamics, Jan. 2006.

14Kunstner, R., Potthoff, J., and Essers, U., “The Aero-Acoustic Wind Tunnel of Stuttgart University,” InternationalCongress and Exhibition, Feb. 27 - Mar. 2 1995, SAE Technical Paper Series 950625.

15Bruun, H. H., Hot-Wire Anemometry: Principles and Signal Analysis, Oxford University Press, Oxford OX2 6DP, 1995.16Oerlemans, S., “Acoustic Wind Tunnel Measurements on GAMESA Baseline and Optimised Airfoils,” Tech. Rep. NLR-

CR-2005-055, NLR, 2005, SIROCCO Deliverable D13a.17Davis, M. R., “Hot Wire Anemometer Response in a Flow with Acoustic Disturbances,” Journal of Sound and Vibration,

Vol. 56, No. 4, 1978, pp. 565–570.18Bendat, J. S. and Piersol, A. G., Random Data, Analysis and Measurement Procedures, John Wiley & Sons Inc., 3rd

ed., 2000.19Oerlemans, S. and Migliore, P., “Aeroacoustic Wind Tunnel Tests of Wind Turbine Airfoils,” AIAA paper 2004-3042,

2004.20Brooks, T. F., Marcolini, M. A., and Pope, D. S., “Airfoil Trailing-Edge Flow Measurements,” AIAA Journal , Vol. 24,

No. 8, 1986, pp. 1245–1251.21Fuglsang, P., Dahl, K. S., and Antoniou, I., “Wind Tunnel Tests of the Risø-A1-18, Risø-A1-21 and Risø-A1-24 Airfoils,”

Tech. Rep. Risø-R-1112, Risø, June 1999.22Fuglsang, P., Bak, C., Gaunaa, M., and Antoniou, I., “Design and Verification of the Risø-B1 Airfoil Family for Wind

Turbines,” Journal of Solar Energy Engineering , Vol. 126, November 2004, pp. 1002–1010.23Gaunaa, M., Fuglsang, P., Bak, C., and Antoniou, I., “Open-Jet Wind Tunnel Validation Using a NACA 0012 Airfoil,”

The Science of Making Torque from Wind , Risoe National Laboratory, Delft University of Technology, the Netherlands, April2004, pp. 37–48.

24Drela, M., “Newton Solution of Coupled Viscous/Inviscid Multielement Airfoil Flows,” 1990, AIAA-90-1470.25Pistolesi, E., “Betrachtungen uber die gegenseitige Beeinflussung von Tragflugelsystemen,” Hauptversammlung der

Lilienthal-Gesellschaft, Berlin, 1937.26Williams, J. F. and Hall, L., “Aerodynamic Sound Generation by Turbulent Flow in the Vicinity of a Scattering Half

Plane,” Journal of Fluid Mechanics, Vol. 40, 1970, pp. 657–670.27Lutz, T., Herrig, A., Wurz, W., Braun, K., and Kramer, E., “Constrained Aerodynamic & Aeroacoustic Design of Wind-

Rotor Airfoils,” First International Meeting on Wind Turbine Noise: Perspectives for Control, Berlin, Germany, 17.-18.10.2005.

28Lutz, T., Herrig, A., Wurz, W., Kamruzzaman, M., and Kramer, E., “Design and Wind-Tunnel Verification of Low-NoiseAirfoils for Wind Turbines,” 24th Applied Aerodynamics Conference, San Francisco, California, 5-8 June 2006, AIAA 2006-3322.

29Oerlemans, S., “Wind Tunnel Aeroacoustic Tests of Six Airfoils for Use on Small Wind Turbines,” Tech. Rep. NREL/SR-500-35339, National Aerospace Laboratory NLR, 2004, NREL report, www.nrel.gov.

30Brooks, T. F., Pope, D. S., and Marcolini, M. A., “Airfoil Self-Noise and Prediction,” Tech. Rep. 1218, NASA ReferencePublication, 1989.

12 of 12

American Institute of Aeronautics and Astronautics