A comparison of NACA 0012 andNACA 0021 self-noise at low Reynoldsnumber

A. Laratro, M. Arjomandi, B. Cazzolato, R. Kelso

Abstract The self-noise of NACA 0012 and NACA 0021 airfoils are recordedat a Reynolds numbers of 96,000 in an anechoic wind tunnel at an angle-of-attack range of −5◦ to 40◦. Results suggest that the low angle-of-attack tonalnoise of the airfoils behaves differently, with the NACA 0021 producing tonesat much higher angles-of-attack but not near 0◦. Noise generated at the onsetof stall is subtlely different, with signature of the NACA 0012 forming over alarger angular range compared to the NACA 0021 where the stall signatureforms suddenly.

Key words: aeroacoustics, stall noise, airfoil noise

1 Introduction

At low angles-of-attack airfoils produce a tonal noise at the trailing-edgewhen instabilities in the transitional boundary layer convect past the trailingedge, generating noise which further excites the boundary layer (McAlpine,1997; Arcondoulis et al, 2009). On NACA 0012 airfoils this noise typicallyoccurs over a small range of angles-of-attack, when the separation point isnear the trailing edge and as the angle increases the tonal frequencies tend toincrease or stay constant (McAlpine, 1997; Arcondoulis et al, 2009). A studyby Hansen et al (2010) on NACA 0021 airfoils indicates that tonal noise is notpresent at 0◦, persists to a higher angle of attack and decreases in frequencyas the angle-of-attack is increased. These differences were not discussed byHansen et al (2010) but suggest that the tonal noise properties of the two

A. Laratro, M. Arjomandi, B. Cazzolato, R. KelsoSchool of Mechanical Engineering, The University of Adelaide, South Australia, Australia,

e-mail: alex.laratro@adelaide.edu.au

1

2 A. Laratro, M. Arjomandi, B. Cazzolato, R. Kelso

airfoils differ. Howeverthis tonal noise is sensitive to environmental factorshindering direct comparison (Hansen et al, 2010).

At high angles-of-attack airfoils act similar to bluff bodies and shed largevortex streets. This generates sound at a frontal-height based Strouhal Num-ber of between 0.15 and 0.2 at moderate Reynolds numbers similar to a flatplate (Fage and Johansen, 1927; Colonius and Williams, 2011). During thetransition to fully separated flow airfoil noise is less well understood.

Brooks et al (1989) conducted extensive testing of the self-noise of NACA0012 airfoils over a range of angles-of-attack. They found that as angle-of-attack increases and the airfoil stalls the shed vortices become larger andself-noise shifts to lower frequencies. While the data of Brooks et al usesrange of angles and Reynolds numbers its detail is limited by the use of1/3-octave spectra which makes it difficult to discern peaks due to vortexshedding. More recent work by Moreau et al (2009) presented much higherresolution spectra and showed that there are peaks in the airfoil noise nearstall that are not resolved when the data is presented in third-octave bands.At angles of attack from approximately 14-20◦ they reported small peaksin the spectra attributed to separation noise that decreased in amplitudeas the angle-of-attack was further increased. Beyond this range the largerand sharper peaks attributed due to bluff body vortex shedding formed andmoved to lower frequencies with higher angles as expected.

The objective of the current work is two-fold; Firstly to expand upon theexperimental findings of Moreau et al (2009) by recording data in the light-stall regime for both NACA 0012 and NACA 0021 airfoils and identifying ifthere are differences in their spectra. Secondly to use the data collected toinvestigate the differences between tonal noise of the NACA 0012 and NACA0021 airfoils as seen in the data of Hansen et al (2010).

2 Method

In order to achieve high angles-of-attack with the available facilities the air-foils had to be mounted vertically. This resulted in a design span of 73mm, toaccount for small deflections in the end plates. A servo motor mounted to therig enabled remote control of the angle-of-attack, and did not significantlyaffect the background noise at high flow-speeds. The rig placed the airfoilseveral chord lengths from the exit of the nozzle (as shown in Figure 1) inorder to give a larger arc line-of-sight to the airfoil for future testing, andseveral chord lengths before the end of the plates to allow wake development.

The microphone used was located 0.61m from the rotation axis at polarand azimuthal angles of 91.2◦ and −0.4◦ respectively. Each spectrum wascreated with Welch’s method with a sampling frequency of 215Hz and a win-dow length of 213Hz with 150 averages. The experiment was conducted a

Title Suppressed Due to Excessive Length 3

Table 1: Experimental parameters

Nozzle size 75mm x 275mm

Airfoil profile NACA 0012 NACA 0021

Chord 50mm

Span 73mm

Thickness 6mm 10.5mm

Aspect ratio 1.46

Relative flow width 5.5

MotorMicrophone

Airfoil

Rig support

Nozzle

Fig. 1: Experimental setup

speed of 30m/s, corresponding to a Reynolds numbers of 96,000. Spectrawere produced for angles-of-attack from -5 to 40◦ at a resolution of 1◦.

3 Results

The experimental results, shown in Figure 2, show tonal noise at low angles-of-attack due to trailing-edge instability and at high angles-of attack as theairfoils begin to act as bluff bodies, as expected. The NACA 0021 airfoil doesnot generate tones until 2−3◦ which then decrease in frequency as the angle-of-attack is increased before fading at 12◦. This is in some agreement withthe results of Hansen et al. (2012) where tonal noise was detected as low as1◦ and corroborates both the trend of tonal noise not being generated at 0◦

and the frequency decreasing as the angle is increased. This suggests that thetonal noise behaviour observed in the NACA 0021 is intrinsic to the airfoil atthis Reynolds number and not a result of environmental conditions. Note thatthe geometric angles-of-attack is presented here and the true angle-of-attackis given by αt = 0.76α (Brooks et al, 1989).

4 A. Laratro, M. Arjomandi, B. Cazzolato, R. Kelso

(a) NACA 0012 (b) NACA 0021

Frequency (Hz)

α(◦

)

100 10000

10

20

30

40

50

60

4000

SP

Lre

20µPa

-5

0

5

10

15

20

25

30

35

40

Frequency (Hz)

α(◦

)

100 10000

10

20

30

40

50

60

4000

SP

Lre

20µPa

-5

0

5

10

15

20

25

30

35

40

Fig. 2: Spectra of airfoils at U = 30ms−1, Re = 96,000, ∆α = 1◦, ∆f = 4Hz

At post-stall angles-of-attack the airfoils display similar vortex sheddingbehaviour. In this regime the airfoils are acting as bluff bodies, and the changein thickness does not significantly affect the vortex shedding frequency. Alow frequency peak similar to that reported by Moreau et al (2009) is seenbetween the tonal and bluff body noise, referred to as light stall in thatstudy. Secondary peaks are also seen at slightly higher frequencies and willbe discussed in further detail below. It is important to note that as the airfoilbegins to stall there is a noticeable increase in broadband noise level belowapproximately 900Hz and a decrease in noise level above. The speed at whichthis change occurs is noticeably different for each airfoil, with the NACA 0021experiencing a more rapid change in noise signature. This is believed to beindicative of the NACA 0021 stalling more sharply in these conditions. Thebehaviour of the NACA 0012 is consistent with both lift data in literatureMarchman et al (1998) and a direct numerical simulation by Rodrıguez et al(2013) which indicated that the onset of stall occurred over a range of anglesfor the NACA 0012 in this Reynolds number range. The simulation indicatedthat the peaks in the noise at stall occur due to instabilities in the separatedshear layer near the leading edge as well as vorticity produced near the trailingedge. Similarly lift data for the NACA 0021 near the experimental Reynoldsnumber how a sharp decrease in lift at the onset of stall corresponding to thesharp increase in low-frequency noise observed in this study Marchman et al(1998). Comparable simulations could not be found in the literature, howeverit is reasonable to assume that similar flow phenomenon are responsible forthe observed noise for both airfoils.

As mentioned previously, some anomalous peaks appear in the data, witha wide peak near 950Hz that increases in strength with angle-of-attack andthen decreases as the airfoil stalls. What was initially believed to be secondarypeaks at around 550Hz may be due to a related phenomenon. These peaks atare located at frequencies that are prominent in the background noise whenthe end-plates are installed as shown in Figure 3, however they rise and fallwith the changes in noise at stall. Because of this it is currently believed that

Title Suppressed Due to Excessive Length 5

Frequency (Hz)

SPLre2

0µPa

100 100010

20

30

40

50

60

70No airfoilNACA 0012, 14◦NACA 0021, 16◦

4000

Stall noise

Fan noise

Potentialaeroacousticfeedback

Fig. 3: Spectra of airfoils at stall for U = 30ms−1, Re = 96,000

these peaks are a result of an aeroacoustic coupling between the airfoil andthe end-plates, and steps are being taken to attempt to reduce the effect.Regardless, the large difference in the behaviour of this coupling betweenthe NACA 0012 and NACA 0021 airfoils suggests that after the coupling issuppressed that some difference in spectrum will remain.

4 Conclusion

NACA 0012 and NACA 0021 airfoils with 50mm chord were tested in theanechoic wind tunnel at the University of Adelaide at a Reynolds number of96,000 at various angles-of-attack. Noticeable differences were found in thecharacteristics of how their self-noise spectra change as the angle-of-attack isincreased, including confirmation of a lack of NACA 0021 tonal noise near 0◦

as seen in the data of Hansen et al (2010). The onset of bluff body behaviouroccurred later for the NACA 0012 airfoil and the onset of stall took placemore gradually compared to the NACA 0021. There is evidence that thereare differences in the broadband behaviour of these airfoils as they approachstall, however due to a possible aeroacoustic coupling between the airfoil andthe experimental rig this cannot be determined conclusively.

References

Arcondoulis E, Doolan C, Zander A (2009) Airfoil noise measurements atvarious angles of attack and low Reynolds number. In: Proc Acoustics2009

6 A. Laratro, M. Arjomandi, B. Cazzolato, R. Kelso

Brooks T, Pope D, Marcolini M (1989) Airfoil self-noise and prediction. Tech.rep., NASA Reference Publication 1218

Colonius T, Williams D (2011) Control of vortex shedding on two- and three-dimensional aerofoils. Phil Trans R Soc A 369(1940):1525–1539

Fage A, Johansen F (1927) On the flow of air behind an inclined flat plateof infinite span. In: Proceedings of Royal Society of London. Series A, vol116, pp 170–197

Hansen K, Kelso R, Doolan C (2010) Reduction of flow induced tonal noisethrough leading edge tubercle modifications. AIAA Paper 2010 3700

Marchman F, Gunther C, Gundlach J (1998) Semi-span testing at lowreynolds number. AIAA Paper 608

McAlpine A (1997) Generation of discrete frequency tones by the flow aroundan aerofoil. PhD thesis, School of Mathematics, University of Bristol

Moreau S, Roger M, Christophe J (2009) Flow features and self-noise of air-foils near stall or in stall. In: 15th AIAA/CEAS Aeroacoustics Conference

Rodrıguez I, Lehmkuhl O, Borrell R, Oliva A (2013) Direct numerical sim-ulation of a naca0012 in full stall. International journal of heat and fluidflow 43:194–203