Experimental Detection of Transition on WindTurbine Airfoils

Peter Bæk and Peter Fuglsang

LM Glasfiber A/S, Denmark

Abstract

A non-intrusive method for detection of laminar to turbulent boundary layer tran-sition, based on infrared thermography, was developed and implemented in thelarge low speed wind tunnel located at LM Glasfiber. The setup was tested onthree different wind turbine blade airfoils at Reynolds numbers up to 6 millions. Acomparison was made between the transition model of XFOIL and the measure-ments, showing that XFOIL predicts the transition point well, but in some isolatedcases has a significant deviation from the measurements. An exact determinationof the transition point is a key to correctly predict airfoil characteristics a wind tur-bine blades, and for development of high lift-to-drag airfoils.

Key words: transition, airfoil, thermography

1 Introduction

The need for a detection method is motivated primarily by the desire to validateand study various transition prediction algorithms that are currently used to de-sign airfoils. Secondarily it allows for verification of the airfoil model surface beingfree of roughness elements prior to a wind tunnel test, which could otherwise leadto erroneous conclusions on the performance of an airfoil. Finally future innova-tions in airfoil add-ons to control the position of the transition point or non-trivialgeometries can be investigated experimentally using a detection technique.

The knowledge of the position of transition is of vital importance in the designof new airfoils for several reasons. First of all the transition location to a greatextent determines the frictional drag of the airfoil, because the turbulent boundarylayer is associated with an order of magnitude higher skin friction than its laminarcounterpart. As a rule of thumb the drag coefficient is decreased by 10% when

Email address: pbk@lmglasfiber.com (Peter Bæk and Peter Fuglsang).URL: http://www.lmglasfiber.com (Peter Bæk and Peter Fuglsang).

EWEC 2009 March 19, 2009

the transition point is moved 10% of the chord length toward the trailing edge.Secondly a correctly positioned transition point will enable the turbulent boundarylayer to overcome a steeper pressure recovery without separation. Limiting theturbulent fraction of the boundary layer is an objective in the development of highlift to drag ratio airfoils. However, there are many dilemmas and compromises tobe made in this field, since the delay of transition leads to a more unreliable airfoil,which may be susceptible to deposition of roughness (dirt, insects, ice, salt, etc.)or prone to development of laminar separation which leads to a high drag. Thetransition point can to a large extent be controlled by modifying the airfoil pressuredistribution.

Since wind tunnel experiments are both expensive and time consuming, airfoildesign is usually based on computer optimization coupled to an airfoil designsoftware, for example the XFOIL code developed by Mark Drela, MIT [2]. Navier-Stokes equation solvers (CFD codes) are still considered too computationally de-manding to do airfoil optimization. XFOIL on the other hand is extremely fast,allowing many iterations in the parameter space, but at the expense of precision.The results must therefore be validated by doing wind tunnel experiments. Espe-cially the drag is usually under predicted by XFOIL compared to measurements.One of the primary reasons for this is the accuracy of the current transition pre-diction models that are incorporated in both panel codes such as XFOIL as wellas CFD codes.

2 Theory

Most of the current transition prediction schemes are based on (semi-)empiricalcorrelations for the linear stability of the laminar boundary layer. One of the mostextensively used is the en envelope model, which was first introduced in theXFOIL code [2]. The en envelope method estimates the amplification of a trav-eling waves in the boundary layer, along the surface of the airfoil. For this it usestabulated values that have been derived using linear stability theory for variousboundary layer shape factors. Once the amplitude ratio, n = log(A/A0), betweenthe initial wave and the amplified one reaches a certain user defined critical value,ncr , transition occurs immediately. Based on empirical knowledge acquired at LMGlasfiber the value, ncr , is usually set to 7, but this is a very vague transition cri-terion because the initial wave amplitude is unknown, and very dependent on thespecific flow conditions. The initial amplitudes that enter the boundary layer fromthe external flow are determined by the receptivity of the boundary layer - a pro-cess that remains to be fully understood by researchers. This also means that thetransition model completely fails, when initial disturbances are so large that tran-sition occurs with little or maybe no amplification - known as by-pass transition.For by-pass transition, other empirical prediction schemes have been developed.

A second short-coming of the en envelope transition model is that it only takesinto account the linear amplification phase, and not the following non-linear breakdown of the waves into turbulent vortices. But these following stages normally2

Fan

Test SectionContraction

Flow

Mesh ScreensHeat Exchanger

Corner Vanes

Figure 1. LM Glasfiber’s wind tunnel seen from above, with top shell removed.

occur very close to the transition point, whereas the linear amplification phasecan extend over the majority of the chord length.

3 Experimental Setup

3.1 Wind Tunnel Facility

The LM Glasfiber Low Speed Wind Tunnel was inaugurated in 2006, and wasspecifically designed for validation of airfoil sections under flow conditions veryclose those that are present on large wind turbine blades. The wind tunnel is aclosed circuit, variable fan speed tunnel, with temperature control [4]. The flowquality is very very high due to specially designed corner vanes, a honeycombstructure, three fine mesh screens and finally a contraction of 10 to 1, that damp-ens the turbulence to a level of 0.1% and ensures a uniform velocity profile of goodquality. This makes the wind tunnel ideal to study natural transition phenomena.See figure 1.

The turbulence intensity in the empty wind tunnel has been reported in [5] to beonly slightly dependent on the air speed as seen in table 1. The flow variationin both time and space of the free stream velocity is less than 0.2% and theangularity is found to be less than 0.2◦.

The maximum wind speed in the test section is 105 m/s which together with a 0.9m chord length yields a chord Reynolds number of 6×106 and a Mach number ofM = 0.3. The test section is shown in figure 2.3

U∞ [m/s] Re × 106 (c =0.9m) Tu [%]

50 3.0 0.098

80 4.8 0.103

100 6.0 0.108Table 1Turbulence intensities in the empty wind tunnel test section. Measured with hotwire. High-pass filtered at 10Hz. [5]

Turn Table

Wall Pressure Tabs

Flow

Load CellsAirfoil Model

Wake Rake

Figure 2. LM Glasfiber’s wind tunnel test section. Pressure distributions are measured onthe airfoil surface, on the ceiling and floor as well as on the wake rake. The load cellsmeasure the integrated force. This ensures a redudant force measurement setup.

3.2 Transition Detection Setup

The transition detection method relies on the principle that a higher skin friction onthe airfoil surface is associated with a higher convectional heat transfer from thesurface. If the surface of the airfoil is heated uniformly the laminar boundary layerwill approach a higher equilibrium temperature than the laminar boundary layer.Internal heat conduction in the model will tend to diminish the temperature stepbetween the two boundary layer types, and make the transition line less clear, butby using thermally insulating materials for the model surface, such as plastics, the4

IR Camera

View PortQuartz

Markers

Heat Lamp

Figure 3. Sketch of the setup. The field of view of the camera must be illuminated by theheat lamp (illustrated with colors on the airfoil).

temperature step can be maintained very sharp.

The airfoil suction side surface was illuminated with a heat lamp placed on theceiling of the test section with optical access through a quartz glass, which hada high transparency to the near infrared light of the lamp. The heat lamp sup-plied approximately 450 W/m2 in the focused region on the airfoil surface. Theinfrared camera was placed adjacent to the lamp with a separate optical accessthrough a barium fluoride glass, which was sufficiently transparent in the far in-frared spectrum. The camera used was an AGEMA 570 with a 150mK sensitivityon a 320x240 pixel focal plane and a 24x18 degree lens. Gold leaf markers wereplaced on the airfoil surface in order to identify the location of the transition pointon the thermographic images. See figure 3.

4 Results

4.1 Transition Point Determination

The transition point is determined from the temperature distribution. In order to getthe best ability to compare with XFOIL, an unsteady finite difference solver wasdeveloped to solve the temperature distribution on the surface of the airfoil, basedon XFOIL values exclusively. Thereby the correlation between the temperatureprofile and the transition point could be found [1].

A validation case was defined to demonstrate the method. On a small spanwisesection of the NACA 0015 airfoil model the transition point was fixed at x = 0.2cusing a piece of zigzag tape (0.4 mm thick). The zigzag tape caused prema-ture transition at zero degrees angle of attack with Reynolds numbers between1.5×106 and 6×106. Figure 4 shows the temperature distribution for a Reynoldsnumber of 6 millions. The transition line is clearly visible with a temperature dif-5

Re [106] xtr/c Picture Temperature Profile

6 0.29 x/c

y/c

0 0.1 0.2 0.3 0.4 0.5 0.6

0

0.5

x/c

T[K

]

0 0.1 0.2 0.3 0.4 0.5 0.6

295296297298299300

Figure 4. Transition detection using heat lamp. Flow from left to right on the picture. Theairfoil is a NACA 0015 with an angle of attack of zero degrees. Markers have been re-drawn. The reflection from the lamp is seen as a hot spot on the upper middle of thepicture. Furthermore the spanwise (vertical) distribution of light not completely homoge-neous. Nevertheless the transition line is clearly detectable. The free transition line isseen as a vertical dashed line.

ference of approximately 2◦C.

4.2 Validation

Validation of the thermographic transition detection method was made using mi-crophones. Thin 0.8 mm polyurethane tubes were placed with the opening point-ing upstream at each 5% chordwise position on the airfoil (from 10% to 40 %chord). The other ends of the tubes were connected to microphones placed out-side the wind tunnel. The time series measured for each angle of attack and foreach tube, were analyzed in a similar manner as the Risø transition measure-ments [3]. Briefly the Risø analysis method was to make a power spectrum of thetime series, Ps, and treat this power spectrum as a probability density function insuch a way that the average frequency was defined as

µ1 =

´ f2f1

fPsdf´ f2

f1Psdf

,

where f is the frequency and f1 and f2 represent the lower and upper frequencybounds limited by the measurement equipment. For this experiment the lowerfrequency was 100Hz and upper frequency was 10kHz. The turbulent boundary6

α[◦]

x tr

[c]

Microphones, Re = 3.0 × 106

Microphones, Re = 6.0 × 106

Infrared, Re = 3.0 × 106

Infrared, Re = 6.0 × 106

-5 0 5 100

0.1

0.2

0.3

0.4

0.5

Figure 5. Left: The transition point measured with infrared camera and with microphonesare compared for two Reynolds numbers. Right: Thermographic image of the tubes atRe=3 × 106 and α = 0◦. The transition cones caused by the tubes are seen not to enterinto the neighboring tubes.

layer will have a broad band of high energy fluctuations and the mean will there-fore normally be high. A laminar boundary layer will typically only contain low fre-quency background noise. µ1 then becomes a function of the angle of attack andthe tube locations, i.e. µ1(α, x). The transition point for a given α is then definedas being near the maximum slope of the µ1(x) curve

xtr (α) = max

(

∂µ1

∂x

∣

∣

∣

∣

∣

α

)

Figure 5 compares the transition point measurements using the thermographicmethod to the microphone measurements. It is clearly seen that the results arevery close despite the two methods were relying on measuring completely differ-ent physical quantities.

4.3 Wind Turbine Airfoils

Three different airfoils developed for wind turbine applications were character-ized with respect to transition location, namely a blade tip airfoil (NACA-64-618,t/c=18%), a blade midsection airfoil (DU-91-W-250, t/c=25%) and a blade rootairfoil (DU-40, t/c=40% modified by LM Glasfiber). See contours on figure 6. Es-pecially the two former airfoils were very interesting to test because they havebeen designed to have a very low lift-to-drag ratio, which requires a large portionof the flow over the airfoil to remain laminar. Such flows are especially challengingfor XFOIL to determine a correct transition points on. The airfoils were tested atReynolds numbers 1.5, 3 and 6 millions and at angles of attack between -20 and20 degrees, measuring the transition point on the suction side of the airfoil.7

Figure 6. Above is the tip section airfoil (NACA-64-618), in the middle a midsection airfoil(DU-91-W-250) and below a root section airfoil (DU-40, modified).

In figure 7 the measured results of the first two airfoils are compared with XFOILcalculations made with a critical amplification ratio of ncr = 7. On the wind turbinethese airfoils operate at angles of attack between 3 and 7 for normal operation,and this range is therefore of highest interest. For the NACA-64-618 case, XFOILis seen to compare very well with measurements above 3 degrees angle of at-tack, which is highly satisfactory. Below 3 degrees angle of attack the differenceincreases with Reynolds number, from roughly no disagreement at Re = 1.5×106

approximately 4% absolute difference at Re = 6×106. For the DU-91-W-250 case,XFOIL’s Reynolds dependency was not completely following the trend of the mea-surements for angles of attack below 3 degrees. Above 4 degrees angle of attackthe model fails to predict the correct transition points, especially at Re = 3 × 106

and α = 7◦, where the difference is more than 10% absolute position, which byrule of thumb corresponds to almost 10% difference in drag coefficient.

For the thick root section airfoil (DU-40, modified), the correspondence on tran-sition point between measurements and XFOIL showed up to be very good, asshown in figure 8. The transition points of XFOIL deviate less than 2% absolutechord length from the measurements. When comparing the measured lift anddrag coefficients to XFOIL values it is clear that XFOIL completely overpredictedthe maximum lift coefficient. This could be due to a difference between the exper-imental and XFOIL’s position of the pressure side transition point, but it is morelikely to be due to XFOIL’s turbulent boundary layer models. If these models arenot capable of determining the correct separation point the thick airfoil, the inte-grated forces will be wrong. But it is valuable information to know that XFOIL’stransition model is not necessary wrong.

5 Conclusion

A fast, reliable non-intrusive setup, based on infrared the rmography wasdeveloped and verified using an acoustic transition detection setup. The transi-tion detection setup is now an integrated part of the measurement setup at thelarge industrial wind tunnel facility at LM Glasfiber.A comparison with the XFOIL transition model showed good agr eement forthe overall trend of the transition point . Generally XFOIL was overpredictingthe laminar portion of the boundary layer, when this was very extended, such as8

α[◦]

x tr/

c

-5 0 5 10 150

0.10.20.30.40.50.60.7

α[◦]

x tr/

c

-5 0 5 10 150

0.10.20.30.40.50.60.7

IR Camera, Re = 1.5 × 106

IR Camera, Re = 3.0 × 106

IR Camera, Re = 6.0 × 106

XFoil, ncr = 7, Re = 1.5 × 106

XFoil, ncr = 7, Re = 3.0 × 106

XFoil, ncr = 7, Re = 6.0 × 106

Stuttgart LWT, c =0.6m, Re = 3.0 × 106

Figure 7. Left: Transition points for NACA-64-618 for different Reynolds numbers com-pared with XFOIL calculations. Transition measurements made with stethoscope fromStuttgart LWT included [6]. Right: Transition points for DU-91-W-250 compared withXFOIL, for the same conditions as on the left figure.

CD [1]

CL

[1]

0 0.05 0.1-0.5

0

0.5

1

1.5

2

α[◦]

LSWT, Re = 3.0 × 106

XFoil, ncr = 7, Re = 3.0 × 106

x tr[

c]

-5 0 5 10 150.2

0.22

0.24

0.26

0.28

0.3

0.32

0.34

0.36

0.38

0.4

-0.5

0

0.5

1

1.5

2

Figure 8. Polar for DU-40 (modified), measured in LM Glasfiber’s wind tunnel (LSWT) andcompared with XFOIL. The lift and drag are seen to deviate a lot above α = 7◦ and belowα = −2◦, while the transition points agree within a 2% range (absolute of chord) for allmeasured angles. 9

for high lift-to-drag airfoils at low angles of attack. In such cases of extensive lam-inar boundary layer, XFOIL overpredicted the transition point by approximately4 % absolute chordwise position, when the critical amplification ratio was set toncr = 7 and a Reynolds number of 6 millions.The XFOIL transition model was found not to be reliable near t he opera-tional design point of the airfoils. XFOIL mostly captured the transition pointwell, but for e.g. the DU-91-W-250 profile, XFOIL calculated the transition pointmore than 10% absolute chordwise position earlier than the measurements, whichcorresponds to a 10% drag difference. It is therefore not clear when XFOIL man-ages to find the correct transtion point and when the model under or over esti-mates the transtion point.Although XFOIL manages to find the correct transtion point, X FOIL doesnot necessarily determine the correct force coefficients. For a very thick pro-file such as the DU-40 modified, the force prediction by XFOIL was very poorcompared to measurements at moderate and especially high angles of attack,albeit that the transition point was correctly determined. This indicated that boththe transition model and the boundary layer models of XFOIL are unreliable forwind turbine airfoils, and that experimental verification of XFOIL results is crucial.

Acknowledgements

A special thanks to Dr. Knud-Erik Meyer from the DTU and Peter Fuglsang fromLM Glasfiber, who supervised Peter Bæk during his MSc thesis work in 2008 [1].

References

[1] Peter Baek. Experimental detection of laminar to turbulent boundary layertransition in an industrial wind tunnel facility. MSc Thesis, Department of Me-chanical Engineering, Technical University of Denmark, 2008.

[2] Mark Drela. Xfoil: An analysis and design system for low reynolds number air-foils. In Low Reynolds Number Aerodynamics. Springer-Verlag Lecture Notesin Engineering 54., 1989.

[3] Mads Døssing. High frequency microphone measurements for transition de-tection on airfoils (risoe r 1645 en). Technical report, Risø DTU, 2008.

[4] Peter Fuglsang and Stefano Bove. Wind tunnel testing of airfoils involves morethan just wall corrections. In EWEC 2008, Brussel, 2008.

[5] Heinz-Dieter Papenfuss. Aerodynamic commissioning of the new wind tun-nel at lm glasfiber (lunderskov). Technical report, Ruhr-Universität Bochum,Institut für Thermo- und Fluiddynamik, 2006.

[6] W. Würz, C. Vetter, and M. Langohr-Kolb. Wind tunnel measurements of thelm naca64-618 airfoil. Technical report, Universität Stuttgart, Institut Für Aero-dynamic und Gasdynamik, 2006.

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