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Experiments in Fluids (2021) 62:106

https://doi.org/10.1007/s00348-021-03204-9

RESEARCH ARTICLE

Aeroacoustic noise reduction byapplication ofend plates

onwall‑mounted finite airfoils

ErikSchneehagen1· ThomasF.Geyer2· EnnesSarradj1· DanielleJ.Moreau3

Received: 16 November 2020 / Revised: 1 April 2021 / Accepted: 6 April 2021 / Published online: 24 April 2021

Abstract

One known method to reduce vortex shedding from the tip of a blade is the use of end plates or winglets. Although the

aerodynamic impact of such end plates has been investigated in the past, no studies exist on the effect of such end plates on

the far-field noise. The aeroacoustic noise reduction of three different end-plate geometries is experimentally investigated.

The end plates are applied to the free end of a wall-mounted symmetric NACA0012 airfoil and a cambered NACA4412

airfoil with an aspect ratio of 2 and natural boundary layer transition. Microphone array measurements are taken in the

aeroacoustic open-jet wind tunnel at BTU Cottbus-Senftenberg for chord-based Reynolds numbers between 75,000 and

225,000 and angles of attack from 0

◦

to 30

◦

. The obtained acoustic spectra show a broad frequency hump for the airfoil base

configurations at higher angles of attack that is attributed to tip noise. Hot-wire measurements taken for one configuration

show that the application of an end plate diffuses the vorticity at the tip. The aeroacoustic noise contribution of the tip can

be reduced when the endplates are applied. This reduction is most effective for higher angles of attack, when the tip vortex

is the dominant sound source.

Graphic abstract

1 Introduction

For all lifting airfoils in a flow, regions of high and low

pressure are created by the flow over the pressure side sur-

face and the suction side surface, respectively. If the airfoil

is finite and hence has a free end exposed to the flow, the

resulting pressure gradient between both sides at this end

* Erik Schneehagen

erik.w.schneehagen@tu-berlin.de

Extended author information available on the last page of the article

Experiments in Fluids (2021) 62:106

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can force the flow around the tip side edge. The tip vor-

tex generated by this spanwise flow is not only responsi-

ble for induced drag but can also be a significant source

of broadband flow-induced noise. Guo (2011) developed a

semi-empirical model for flap side edge noise in which two

noise generation mechanisms are identified: the flow separa-

tion near the sharp corners of the flap and the interaction of

large-scale vortex structures with the sharp corners of the

flap. For low Mach number flows, the resulting pressure fluc-

tuations on the surface are efficient noise radiators. Thereby,

the magnitude of the sound radiation strongly depends on

the strength of the vorticity, which is created by the sweep

of the flow around the edge as well as on the distance of the

vortex from the edge (Hardin 1980). Studies on airfoil tip

noise by Moreau etal. (2016), Moreau and Doolan (2016)

demonstrated that the dominant tip noise contribution mani-

fests itself as a broad peak or hump in the noise spectrum.

The broad tip noise peak increases in amplitude as the angle

of attack is increased and, to a lesser extent, when flow speed

is increased.

For aircraft airframe components including the main wing

and flap side edge, it is crucial to not only reduce noise to

meet regulations but also to increase the aerodynamic effi-

ciency. Especially for airfoils used in rotating applications,

like wind turbine blades, the noise contribution from the

tip is very important. This is due to the fact that the flow

velocity along the span takes maximum values in this region

(George etal. 1980) and that the resulting tonal noise occurs

at frequency ranges where it can be perceived as particularly

annoying.

While the physics of this phenomenon is quite well under-

stood and corresponding scaling models exist (George etal.

1980; Brooks and Marcolini 1986; Moreau etal. 2016),

measures for preventing the tip vortex formation are not

extensively studied. In general, three different, passive

measures exist: one is the introduction of porous material

at the tip in order to reduce the pressure jump between suc-

tion and pressure side and therefore reduce the maximum

strength of the vorticity (Revell etal. 1997; Angland etal.

2009). Another measure is to modify the geometry of the

tip region with the aim to minimize the interaction between

the vortices and the side edge (Kinzie etal. 2013) and the

third is the application of winglet-type fences to diffuse the

vorticity at the edge by creating a physical obstacle from top

to bottom (Slooff etal. 2002; Zaman etal. 2017). The latter

is also commonly used to reduce airfoil drag in flight. While

the flow field and vortex generation is investigated in detail

in these studies, the aeroacoustic effect of these measures is

considered very little or not at all.

This paper aims to study the effect of end plates applied

to the free end of a wall-mounted finite airfoil on the result-

ing noise generation. For this purpose, two airfoil shapes

are considered, which are widely used in studies on airfoil

aerodynamics and aeroacoustics: the symmetric NACA0012

airfoil and the cambered NACA4412 airfoil. They are often

used in low-speed applications such as drone/UAV propel-

ler blades at low-to-mid Reynolds numbers. Due to the dif-

ference in camber and the resulting differences in surface

curvature, these airfoils are known to show quite different

behaviour regarding flow separation and transition (Zhang

etal. 2015). The symmetrical NACA0012 airfoil develops

a stronger adverse pressure gradient, which leads to the fact

that flow detaches at a position further upstream than for

the cambered NACA4412 airfoil. To illustrate the differ-

ences, Fig.1 shows the pressure coefficient calculated for a

NACA0012 airfoil and a NACA4412 airfoil at an angle of

attack of 15

◦

using XFOIL (Drela 1989).

The investigated end-plate geometries are taken from an

early study on the aerodynamic characteristics of an unswept

NACA

641A412

airfoil at a Reynolds number (based on

Fig. 1 Pressure coefficient of a

2D NACA0012 airfoil versus

that of a NACA4412 airfoil

[calculated for inviscid flow at

Re=125,000 and M=0.0787

using XFOIL (Drela 1989)]

Experiments in Fluids (2021) 62:106

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airfoil chord length) of

106

with multiple different end-plate

designs (Riley 1951). Three of these end-plate geometries

which are easy to manufacture and show the most benefi-

cial aerodynamic characteristics are considered in the cur-

rent paper. To show the favourable aerodynamics of these

end plates, results from the original study by Riley (1951)

are displayed in Fig.2. In Fig.2(a), it can be seen that the

lift coefficient

is larger for all angles of attack when end

plates are applied to both sides of the airfoil. It is possi-

ble that the end plates effectively change the shape of the

wing or slightly increase the effective span. Figure2(b) then

shows that the lift-drag ratio is higher with end plates for

large lift coefficients. The width of the range where the lift-

drag ratio takes maximum values also increases, although

the maximum value is higher for the base configuration

without end plate.

Apart from these characteristics, the effect of an applied

end plate on the flow around an airfoil was examined quali-

tatively by performing basic flow visualization experiments

in the UNSW wind tunnel (Doolan etal. 2018). The results

are shown exemplary in Fig.3. For these photographs, taken

from the suction side, smoke was introduced to the flow

upstream of a NACA4412 airfoil at 15

◦

angle of attack and

Re = 50,000. The photographs show clear paths of the flow

as white smoke streamlines. The top view of the baseline

configuration (Fig.3a) shows an absence of smoke in the

wake of the airfoil close to the free end. It can be assumed

that this is caused by the tip vortex. For the case when the

circular end plate is applied to the free end of the airfoil

(Fig.3b), no such absence of smoke can be seen, as the

smoke seems to be distributed more homogeneously in the

wake. Thus, it appears that the end plate suppresses the vor-

tex formation at the tip.

(b)(a)

Fig. 2 Aerodynamic characteristics of single NACA

641A412

(baseline) and in combination with various end plates, data replotted from Riley

(1951)

Fig. 3 Flow visualization performed for a NACA4412 airfoil with/without end plate, Re = 50,000, 15

◦

angle of attack, top view suction side

with airfoil edges as white dotted lines, flow from left to right

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For this modification of the flow field, the current study

aims to add an aeroacoustic point of view and present the

noise reduction potential of these end plates for flow con-

figurations where tip noise is dominant. Therefore, the

noise reduction effect is experimentally investigated in

an aeroacoustic wind tunnel for symmetric (NACA0012)

and cambered (NACA4412) airfoils with natural bound-

ary layer transition. The dataset presented in this paper

includes microphone array measurements for chord-based

Reynolds numbers from 75,000 to 225,000 in increments of

25,000. The angle of attack is varied from 0

◦

to 25

◦

for the

NACA0012 airfoil and from 0

◦

to 30

◦

for the NACA4412

airfoil in 2.5

◦

-steps. Additionally, hot-wire measurements

were taken for one configuration in order to obtain more

details on the flow field and to relate this to the noise reduc-

tion mechanism.

The paper is structured as follows: In Sect.2, the experi-

mental set-up and methods are described. Section3 presents

acoustic spectra and sound maps from microphone array

measurements and flow velocity fields from hot-wire meas-

urements in the wake of the airfoil tip. Finally, Section4

gives concluding remarks.

2 Experimental method

2.1 Airfoil models

The two airfoil profiles NACA0012 and NACA4412 have a

theoretical chord length C of 70mm and span L of 140mm

resulting in the aspect ratio L/C = 2. The actual chord

length is 67mm due to a truncated rounded trailing edge

with curvature diameter of 1.0mm. These airfoils follow the

NACA four-digit series and have 0% and 4% camber at 40%

chord, respectively. For some selected configurations, forced

boundary layer transition was achieved by using 60-degree

zigzag trip tape (manufactured by Glasfaser Flugzeugser-

vice) with 0.4mm thickness and 6mm point-to-point dis-

tance on both sides of the airfoil at 10% chord. The three

analysed end-plate geometries (Riley 1951) are shown

in Fig.4. All end plates have the theoretical chord length

of 70mm. The height of the square plate (a) is 0.5chord

lengths, the radius of the circular plate (b) is one chord

length, and the height of the trapezoidal end plate is 0.1

chord lengths at the leading edge and 1.2chord lengths at

the trailing edge. The end plates were attached to the airfoils

with double-faced adhesive tape. The end plates were manu-

factured out of 2-mm-thin plastic (polyurethane) plates. The

edges are straight and untreated.

2.2 Wind tunnel

The measurements were taken in the aeroacoustic open jet

wind tunnel at BTU Cottbus-Senftenberg (Sarradj etal.

2009). The experimental set-up can be seen in Fig.5a. The

two airfoil profiles, NACA0012 and NACA4412, were

attached to an acrylic glass plate, which was fixed on the rec-

tangular nozzle exit (green), which has the width of 280mm

and height of 230mm. At the nozzle exit, the flow is charac-

terized by an even profile and a very low turbulence intensity

in the order of 0.2% (Geyer etal. 2015). The thickness of the

boundary layer on the side plate at the position of the airfoil

leading edge is in the order of 5mm to 10mm at flow speeds

ranging from 10 to 50m/s (Moreau etal. 2017). Due to the

small size of the airfoils compared to the nozzle exit area, no

blockage correction was performed. The microphone array

was mounted on the ceiling of the test section, and all other

surrounding walls were covered with sound-absorbing mate-

rial, which provides an anechoic environment for frequen-

cies above 125Hz.

(a) (b) (c)

Fig. 4 End-plate geometries with NACA0012 (solid, black line) and NACA4412 (dashed, blue line) outline: a square (height 0.5chord lengths),

b circle (radius 1.0 chord length), c trapezoid (height 0.1 to 1.2chord lengths)

Experiments in Fluids (2021) 62:106

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The angle of attack was adjusted with an electronic angle

meter, which was placed on a ledge parallel to the airfoil

chord line. It has to be noted that for a finite length airfoil

with flow over the tip, the effective angle of attack changes

along the span, which will also affect the lift coefficient. For

the configurations without end plates, the angle of attack

distribution can be calculated using the procedure reported

by Awasthi etal. (2018). No method exists for the calcula-

tion of an effective angle of attack that considers the pres-

ence of end plates. Hence, it was decided to perform all

far-field noise comparisons at the same geometric angles

of attack and not at the same lifting conditions. However,

it should be kept in mind that the airfoil generates more lift

at the same geometric angle of attack when end plates are

applied (see Fig.2a) and also the lift-drag ratio is altered

(Fig.2b). Although the original study by Riley (1951) was

done with a different airfoil profile and two free ends, the

general trend of the aerodynamic characteristics still pro-

vides some insight.

2.3 Acoustic measurement set‑up anddata

processing

The acoustic measurements were taken using a planar micro-

phone array consisting of 47 1/4-inch Panasonic microphone

capsules (WM-61A), which were flush-mounted on the rigid

ceiling. The microphone distribution together with the coor-

dinate system is shown in Fig.5b. The axis origin matches

the array centre, while

- and

- were the streamwise and

spanwise axis, respectively. The distance to the airfoil trail-

ing edge at 0

◦

angle of attack was 710mm.

All pressure signals were recorded over a duration

of40seconds with a National Instruments 24-bit multichan-

nel system, (including PXI-4472 cards in a NI PXI-1044

chassis). The sampling frequency was 51.2kHz.

The Acoular 19.02 framework (Sarradj and Herold 2017)

together with the CLEAN-SC beamforming algorithm

(Sijtsma 2007)was used to locate and evaluate the sound

sources in a 40cm x 40cm focus grid with a grid resolution

of 1cm. This algorithm was chosen due to its fast and stable

performance, especially for aeroacoustic testing (Herold and

Sarradj 2017; Merino-Martínez etal. 2019; Bahr etal. 2017;

Sarradj etal. 2017). Hanning window with the block size of

4096samples and 50% overlap was applied for calculating

the cross-spectral matrices resulting in the frequency reso-

lution of 12.5Hz. Additionally, the downstream convection

of the sound waves by the flow was accounted for by using

a ray-tracing approach together with an analytical approxi-

mation of the flow field of a slot jet (Albertson etal. 1950)

over the nozzle width. The main diagonal of the averaged

cross-spectral matrix was removed to avoid contamination

of the signals by incoherent noise (Hald 2017). A value of

6dB was subtracted from the sound pressure level spectra to

account for reflections at the rigid microphone array plate.

For all the baseline cases (no end plate attached) at angles

of attack from 0

◦

to 12.5

◦

the dataset from Zhang etal.

(2020) was used. These measurements were taken with the

same measurement set-up as described here.

The beamforming results were used to quantitatively ana-

lyse the sound pressure at the centroid of the array micro-

phone positions (0, 0.07, 0), which is not exactly at the array

centre due to the non-symmetric microphone distribution. In

(a) (b)

Fig. 5 Acoustical measurement set-up in aeroacoustic wind tunnel, nozzle (green), microphone array in ceiling and microphone positions (red)

Experiments in Fluids (2021) 62:106

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106 Page 6 of 17

order to investigate the locations of dominant sound sources,

the source region was subdivided into three different areas

from which the contribution to the overall sound pressure

level was evaluated. In Fig.6, these integration sectors are

shown.

The whole integration area expands to half a chord length

upstream the leading edge and downstream the trailing edge.

In the spanwise direction, it expands a third chord length

further than junction and tip. The tip integration area is one-

fourth of the whole span, and the integration sectors of lead-

ing and trailing edge are three quarters of span, respectively,

and end at half chord. The integrated sound pressure levels

from these areas were calculated from 1kHz to 20kHz. The

frequency range is chosen so that the Helmholtz number

of the array aperture and the lowest frequency is not lower

than 4 to ensure an accurate performance of the CLEAN-SC

algorithm (Herold and Sarradj 2017).

2.4 Hot‑wire measurement set‑up

In order to investigate the flow structures in the wake of

the tip, constanttemperature anemometry (CTA) measure-

ments were taken on the NACA0012 airfoil for an angle

of attack of 15

◦

and a Reynolds number of 125,000 for the

base configuration and the square and circular end plate. A

Dantec miniature X-wire probe 55P64 was used to measure

the velocity fluctuations in both the streamwise direction

x and in the direction z perpendicular to the flow and the

span/trailing edge. Additionally, a Dantec 90P10 tempera-

ture probe was used for temperature correction during the

measurements. The velocity calibration of the hot-wire

probe was performed at twelve logarithmically spaced cali-

bration points against a vane anemometer with an accuracy

±0.2

m/s/

±2

% and a measurement range of 0.4m/s

…

40m/s using a fourth-order polynomial curve fit. The hot-

wire measurement system contains an internal low-pass filter

with a cut-off frequency of 10kHz, and both probes were

connected to a 24Bit National Instruments digital signal

acquisition module with a sampling frequency of 25.6kHz.

The probe was positioned using a three-dimensional traverse

system with a minimum step size of 0.1mm. The traverse

speed was set low enough to avoid any vibrations of the

probe.

The flow velocity time signal was measured for 8seconds

at discrete points in the yz-plane, 0.1chord lengths down-

stream of the trailing edge. The plane dimensions stretched

spanwise from 5mm over the tip to 22mm spanwise of the

airfoil and in the other dimension from 55mm above the

trailing edge to 33mm beneath it. The resolution was 1mm

in the direct vicinity of the trailing edge and 2mm else-

where. This resulted in a total of 1196measurement points

per configuration. At each point, the mean velocity compo-

nents U (in streamwise direction x) and W (in vertical direc-

tion z) as well as the total turbulence intensity

√

(u2

rms +w2

rms)∕U

were calculated, with

urms

and

wrms

being

the root-mean-square values of the velocity fluctuations in

streamwise and vertical (upwash) directions, respectively,

and

being the outer flow speed.

Fig. 6 Different integration sec-

tors for sound pressure calcula-

tion out of beamforming results,

the airfoil (black) is seen from

top view, LE (leading edge), TE

(trailing edge)

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3 Experimental results

3.1 Beamforming results—NACA0012

First the procedure of evaluating the noise reduction effect

is explained by means of an example. Figure7 shows the

sound maps for one frequency band of the far-field noise

measured for the NACA0012 airfoil at 15

◦

angle of attack

at a Reynolds number of 125,000 for the baseline configura-

tion and the circular end plate attached. For every frequency,

the sound pressure level (SPL) of the sources is integrated

within the shown area around the whole airfoil (Tip + LE +

TE, Fig.6). This results in the spectra shown for both con-

figurations and is also done for the empty wind tunnel at the

same flow speed. The overall sound pressure level (OASPL)

is calculated for all frequencies between 1kHz and 20kHz

where the signal-to-noise ratio to the empty wind tunnel is

above 6dB for both configurations. The total noise reduc-

tion achieved by applying the end plate then corresponds to

the difference of these two overall sound pressure levels and

amounts to 1.3dB in this example.

This procedure is done for all three end-plate configura-

tions of the NACA0012 airfoil. For the square and the circle

end plates, the angle of attack is varied from 0

◦

to 25

◦

in 2.5

◦

increments and for the trapezoid from 10

◦

to 20

◦

. The Reyn-

olds number is varied from 75,000 to 225,00 in 25,000 incre-

ments. Every colour map in Fig.8 represents one end-plate

configuration, and every tile is the described difference of

overall sound pressure level between this end plate and the

base configuration for the entire airfoil integration region.

The range of the colour scheme is saturated at ±10dB.

Overall, the three different end plates show a quite similar

behaviour. For zero angle of attack, the end plates increase

the self-noise sound pressure level for almost all flow speeds.

For low angles of attack up to 12.5

◦

, the sound pressure level

difference fluctuates around 0dB, which results in a seem-

ingly random pattern of decrease and increase in the overall

sound pressure level. A possible explanation could be that

for lower angles of attack the noise contribution from the

tip becomes less important (Moreau etal. 2016). For lower

angles of attack the pressure difference between suction and

pressure side is reduced. This pressure difference drives

the cross-flow around the tip which is the source of the tip

vortex. Other studies on NACA0012 airfoils have shown

that the tip vortex strength is related to lift produced by the

airfoil (Zaman etal. 2017) and that the tip vortex size and

radius of curvature increase with angle of attack (Moreau

and Doolan 2016). If the tip vortex strength is reduced, the

sound radiation from interaction with the airfoil solid surface

also decreases. Therefore, other source regions like the trail-

ing edge could be dominant and not affected positively by

the end plates. In addition, it is possible that aerodynamic

noise is generated by flow over the trailing edge of the end

plates themselves.

At higher angles of attack it can be seen that the end

plates become more effective in reducing the noise at all

flow speeds. The highest overall noise reduction achieved by

the investigated end plates is 7dB at 20

◦

and

Re =75, 000

The noise reduction does not seem to depend strongly on the

shape of the end plates because all three configurations yield

similar trends and values.

To further investigate the effect of the end plates, the

power spectral densities at the reference point for the three

Fig. 7 Example for calculating overall sound pressure level differ-

ences of baseline and end plate: sources in sound maps on the left

(examples for 6300 Hz, airfoil black dashed line) are integrated

over sector (black, red) for every frequency band to obtain spectrum

(right) and calculate OASPL difference for frequencies where the sig-

nal-to-noise ratio to the empty wind tunnel is more than 6dB

Experiments in Fluids (2021) 62:106

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angles of attack with the highest noise reduction for all con-

figurations and every second flow speed are plotted in Fig.9.

The Strouhal number and Reynolds number are based on the

airfoil chord length. Results obtained for the lowest flow

speed are shown as the spectra at the bottom, so that one

row from left to right from Fig.8 is shown here from bottom

to top. The Strouhal number range shown is for the highest

flow speed and frequency range from 1 to 20kHz where the

signal-to-noise ratio is large enough.

The spectra for all angles of attack show a decrease in

sound pressure level with increasing frequency except for

the tripped baseline case at 10

◦

. For 10

◦

and 15

◦

angle of

attack, the baseline configuration shows a broadband hump

that is located at Strouhal numbers between 18 and 20. This

shows that the peak frequency is proportional to the flow

velocity. With forced boundary layer transition, this peak is

also apparent although the level decreases. This means that

no laminar boundary layer instabilities are responsible for

this hump.

Overall it can be seen that for all end-plate configura-

tions this broadband peak vanishes and the decrease in

sound pressure level continues with the same slope towards

higher frequencies. The trapezoidal end plate develops a

tonal peak that has a lower frequency than the broadband

peak. Its Strouhal number is between 7 and 8, and it most

likely can be attributed to the vortex shedding from the sharp

edges of the end plate in this configuration. The same applies

to the square end plate at a slightly higher Strouhal num-

ber. For lower flow speeds, the sound pressure level of the

high-frequency hump is higher compared to the broadband

noise at low frequencies. For 15

◦

and

Re =75, 000

, the level

difference is around 5dB while at

Re =225, 000

it is roughly

15dB. For 10

◦

, the trend is similar except for the lowest flow

speed where the hump is not existent. This trend explains

why the noise reduction shown in Fig.8 is more effective at

low flow speeds.

For a geometric angle of attack of 20

◦

, the spectral behav-

iour of the base configuration changes. The spectral hump

is broadened compared to the other two angles of attack.

While the lower frequency level and the slope of the spectra

towards higher frequencies stay almost the same for the end-

plate configurations, the baseline spectra are a few dB higher

in the complete frequency range. This leads to a high overall

noise reduction, and once again the difference is larger for

lower Reynolds numbers.

As an example for a fully turbulent case, the spectrum

obtained for the tripped NACA0012 airfoil with attached

end plates is shown at 15

◦

angle of attack in Fig.9. As

observed for the untripped cases, the formation of the spec-

tral hump is suppressed in the tripped cases by applying

end plates as well. For the higher angles of attack where

the noise reduction effect of the end plates is effective, the

boundary layer transition seems to have no great influence

on the far-field noise.

It is noticeable that all the shown sound pressure level

spectra for the end-plate configurations have a similar shape

compared to the far-field noise spectra from a different study

on self-noise of airfoils in stall (Moreau etal. 2009). In that

Fig. 8 Overall sound pressure level difference of integrated spectra

for different tip configuration for frequency range 1 to 20 kHz for

NACA 0012 for different Reynolds numbers and angles of attack,

integration area is whole airfoil, blue indicates a noise reduction and

red a noise increase

Experiments in Fluids (2021) 62:106

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Page 9 of 17 106

paper, NACA0012 airfoils with aspect ratio L/C 1.6 were

attached between two walls and the total pressure was meas-

ured close to the trailing edge to detect the airfoil flow con-

dition. Except for the absolute level, the spectral shape of

the far-field noise for stall conditions resembles that of the

far-field noise observed in the current study for the end-

plate configurations. This indicates that in the current meas-

urements stall occurred for higher angles of attack as well.

The results for the end-plate configurations are close to the

results of the 2D case from Moreau etal. (2009), where also

no broadband frequency hump is apparent.

In the next step, the contribution of each integration area

from Fig.6 to the overall sound pressure level is investi-

gated. Here, the circular end plate is chosen exemplarily,

because it results in the highest noise reduction for most

cases. Qualitatively, these results are similar to those of the

other two configurations, which will not be shown here. In

Fig.10, the overall sound pressure level difference for the

whole airfoil is shown and then the differences in SPL due to

noise sources in the tip region, the leading edge (LE) region

and the trailing edge (TE) region. Thereby, the contribu-

tion of noise reduction from a particular region needs to be

Fig. 9 Power spectral density for NACA0012 airfoil for 10

◦

, 15

◦

, 20

◦

and 15

◦

angle of attack and every second flow speed for all end-plate

configurations and the baseline case (tripped/untripped), the lowest

flow speed belongs to the bottom graphs, between each flow speed

20dB are added to the spectra to avoid overlay (frequency resolution

reduced to 100Hz to increase visibility)

Experiments in Fluids (2021) 62:106

1 3

106 Page 10 of 17

considered in conjunction with the level difference of the

whole airfoil; only if both show a noise reduction, this par-

ticular integration area is relevant. To improve distinctive-

ness of the different cases, the same colour bar as in Fig.8 is

used although the noise reduction reaches maximum values

of up to 16dB for some cases in the tip region.

For the configurations when the end plate is effective in

decreasing noise, it can be seen that the reduction in the tip

region is strongest. For these cases, a slight decrease in SPL

at the leading and trailing edge can be observed. On the

contrary, for lower angles of attack the SPL increases the

strongest at the tip. The noise spectra for these cases show

tonal peaks that get shifted to higher frequencies and higher

amplitudes when the end plate is applied to the airfoil.

For a more detailed look at the noise source locations,

third octave band sound maps for the dominant frequen-

cies from the broadband hump are shown in Fig.11 for the

NACA0012 base configuration and the circular end plate

applied for 15

◦

angle of attack and

Re =125, 000

. Although

the difference of the overall sound pressure level is only

1.3dB for this example, it still gives insight to important

areas of noise generation. For all selected frequency bands,

Fig. 10 Overall sound pressure level difference of integrated spectra for circular end plate for frequency range 1to 20kHz for NACA0012 for

different Reynolds numbers and angles of attack, blue indicates a noise reduction and red a noise increase; for integration sectors see Fig.6

Fig. 11 NACA0012 third octave band sound maps obtained at 15

◦

angle of attack for

125, 000

, top base configuration, bottom circular end

plate, flow from left to right (nozzle exit at –0.386m, black dotted line represents airfoil, grey dashed line represents wall)

Experiments in Fluids (2021) 62:106

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Page 11 of 17 106

the base configuration shows a dominant sound source at the

trailing edge of the wing tip. The sound pressure level of this

source takes values that are between 10dB and 18dB higher

than the strongest source observed for the circular end-plate

configuration. Here, a large number of weaker noise sources

are distributed over the whole span of the trailing edge as

well as at the leading edge near the wall. This shows that for

these frequency bands the most dominant sound source of

the reference airfoil is at the tip and can be associated with

tip vortex formation. By applying the circular end plate, this

sound source is effectively suppressed.

3.2 Beamforming results—NACA4412

In Fig.12, the difference of the overall sound pressure level

between the base configuration and the three end plates is

shown for the NACA4412 airfoil at the same Reynolds

numbers and also additionally for 27.5

◦

and 30

◦

angle of

attack. For this cambered airfoil, the trend is similar for all

three end-plate shapes. For low Reynolds numbers and low

angles of attack, an increase in sound pressure level can be

observed. For angles of attack up to 12.5

◦

and higher Reyn-

olds numbers, the trend rather reveals a noise reduction for

the square and circular end plate, while the trapezoid shows

a slight increase in sound pressure level. For the lower Reyn-

olds number cases between 15

◦

and 20

◦

, a higher reduction

in sound pressure level can be observed than in the cases

with higher flow speed. The maximum noise reduction is

10dB for 20

◦

and

Re =100, 000

. For these cases, the trape-

zoidal end plate shows a slight sound pressure level increase.

A noise reduction can be seen for all flow speeds at angles

of attack above 20

◦

. This effect vanishes again above 25

◦

angle of attack. Compared to the NACA0012 airfoil, a larger

area of the measurement matrix shows favourable behav-

iour in reducing noise. In comparison with the symmetric

NACA0012 airfoil, no strong tonal peaks are apparent at

low angles of attack. This may be due to the fact that flow

separation is less likely to occur for the cambered airfoil due

to the smaller surface pressure gradient (see Fig.1).

The power spectral densities obtained for the NACA4412

airfoil at four angles of attack are shown exemplary in

Fig.13 for every second flow velocity. For 10

◦

angle of

attack, a broadband peak can again be observed in the upper

frequency range. The highest sound pressure levels occur

at a chord-based Strouhal number of around 12 for lower

flow speeds and 15 for higher flow speeds. Again, a clear

dependency of the peak frequency on the flow speed is seen,

although in this case the Strouhal number is lower compared

to that observed for the NACA0012 airfoil. This peak is

not apparent for all end-plate configurations. In comparison

with the sound pressure level of the symmetric NACA0012

at the same angle of attack, the center frequency of the peak

Fig. 12 Overall sound pressure level difference of integrated spec-

tra fordifferent tip configuration for frequency range 1to 20kHz for

NACA 4412 for different Reynolds numbers and angles of attack,

integration area is whole airfoil, blue indicates a noise reduction and

red a noise increase

Experiments in Fluids (2021) 62:106

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106 Page 12 of 17

is lower for the NACA4412 airfoil case. Also, the level

of the peak does not decrease as strongly with increasing

flow speed. For low flow speeds, it even exceeds the low

frequency noise level. Therefore, the end plates are more

effective in reducing the noise level than for the symmetric

airfoil at the same angle of attack and Reynolds number. The

tripped baseline configuration shows a less prominent hump,

which is shifted towards lower frequencies and the mid-fre-

quency levels are also reduced. As seen for the NACA0012,

it can be assumed that the end plates would also completely

suppress the hump in the tripped configuration.

For the NACA4412 at 15

◦

angle of attack not only the

trapezoidal end plate but also the square plate develop a

tonal peak for higher velocities that has a lower frequency

than the broadband peak. It is assumed that vortex shed-

ding from the sharp edges of these configurations causes

this tonal noise contribution. In comparison with the sound

pressure spectra obtained at 10

◦

angle of attack, the hump at

◦

is broadened, which increases the noise reduction effect

of the end plates.

For 20

◦

angle of attack and the two highest flow speeds,

the broadband peak of the baseline configuration is less

Fig. 13 Power spectral density for NACA4412 airfoil for 10

◦

, 15

◦

, and 25

◦

angle of attack and every second flow speed for all end-

plate configurations and the baseline case (tripped/untripped), the

lowest flow speed belongs to the bottom graphs, between each flow

speed 20dB are added to the spectra to avoid overlay (frequency res-

olution reduced to 100Hz to increase visibility)

Experiments in Fluids (2021) 62:106

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Page 13 of 17 106

prominent and has values below the level of the low-fre-

quency broadband noise. Towards lower Reynolds numbers,

the level of the peak rises significantly above the broad-

band level. The sound pressure level spectrum at the lowest

Reynolds number shows a broadened peak that increases the

sound pressure level by 6dB down to 1kHz compared to

the end-plate configurations. Once again the spectra of the

end-plate configurations show a decrease in sound pressure

level with a constant slope with increasing frequency and

their amplitude therefore is below that of the baseline con-

figuration. The only exception is the trapezoidal end plate

that develops a tonal peak with increasing amplitude for

increasing Reynolds number.

The far-field noise spectra observed at 25

◦

angle of attack

show a different spectral behaviour. All four configurations

result in the same spectral shape, but the spectrum obtained

for the baseline configuration has a notably higher level than

those of the end-plate configurations. The difference is in the

order of 6dB at lower frequencies and 3dB at higher fre-

quencies. This effect leads to ahigh overall noise reduction

that is visible in Fig.12 for this angle of attack.

As was done for the NACA0012 airfoil, the noise con-

tributions originating from different regions of the airfoil

are investigated for the NACA4412 airfoil as well, using

the integration sectors from Fig.6. The results are shown

exemplary for the circular end plate in Fig.14, because

maximum noise reduction is achieved for this configura-

tion. In the tip region, a high noise reduction is achieved for

angles of attack between 12.5

◦

and 25

◦

. The sound pressure

level difference compared to the baseline configuration is

above 10dB for almost all operating conditions. Towards

higher Reynolds numbers, the difference becomes smaller.

This effect is also observed in the spectra in Fig.13, where

with increasing flow velocity the level of the broadband peak

becomes smaller in comparison with the rest of the spectra,

and therefore, the noise reduction effect of the end plates

decreases. For angles of attack below 12.5

◦

, the difference

decreases gradually. The 22.5

◦

and 25

◦

cases indicate that

the trailing edge noise has a higher contribution to the over-

all sound pressure level because the high noise reduction

at the tip does not affect the difference for the whole airfoil

strongly. These are the cases where in Fig.13 a different

spectral behaviour is observed, and therefore, different flow

mechanisms on the airfoil can be expected. For all other

angles, no significant difference is observed except for the

lowest Reynolds numbers where the level is increased.

The noise originating from the leading edge region is

increased by a few dB for almost all cases from 0

◦

to 15

◦

angle of attack. For higher angles, the leading edge noise is

not strongly affected by the circular end plate.

Selected third octave band sound maps are shown in

Fig.15 for the NACA4412 airfoil at an angle of attack of

◦

. The chosen frequencies belong to the broadband hump

observed at a Reynolds number of 125,000. Results are

compared for the baseline configuration and the circular end

plate attached. Like the symmetric airfoil, the NACA4412

baseline configuration shows the strongest sources at the

trailing edge at the very tip. The noise reduction of the

strongest source reaches values between 11dB and 19dB

when the circular end plate is attached. The position of

the sources is shifted a few centimeters down the span. In

addition, other sources at the leading edge wall junction

are apparent. Again, the end plate seems to suppress the

strong sound source associated with the tip vortex and a

Fig. 14 Overall sound pressure level difference of integrated spectra for circular end plate for frequency range 1 to 20kHz for NACA4412 for

different Reynolds numbers and angles of attack, blue indicates a noise reduction and red a noise increase; for integration sectors see Fig.6

Experiments in Fluids (2021) 62:106

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106 Page 14 of 17

noise reduction of 3.6dB is achieved for the overall sound

pressure level.

3.3 Hot‑wire measurement results

To improve understanding of the flow field at the airfoil tip,

results from the hot-wire measurements are presented for the

NACA0012 at 15

◦

angle of attack and a Reynolds number of

125,000. For this case, a broadband peak in the sound pres-

sure level spectrum is apparent, and corresponding sound

maps can be seen in Fig.11. The first row in Fig.16 shows

the turbulence intensity (including the streamwise and verti-

cal component) for the baseline configuration (left) and the

configurations with circular (center) and rectangular (right)

end plate. The measurement plane is located 0.1 chord

lengths downstream of the airfoil trailing edge, the location

of which is indicated in the plot. The airfoil ends at (0, 0) so

that only the outermost 20mm of the airfoil span are shown.

The suction side is in the positive z-direction. Between the

measurement points, the values are interpolated.

Two different flow fields can be observed: For the base-

line case with no end plate applied the maximum turbulence

intensity occurs along the trailing edge and around the tip

in shape of a vortex structure that develops from the pres-

sure side of the airfoil around the tip towards the suction

side. The maximum turbulence intensity is around 16%. The

flow field in the case of both end-plate configurations shows

notably different features. A high turbulence intensity can

also be observed at the trailing edge, but additionally in a

large spanwise area, elevated above the trailing edge, another

high turbulence intensity region is present. The maximum

intensity is significantly higher with 27% for the circular

end plate and 24% for the rectangular end plate. A possible

explanation for the different flow fields observed could be

the free end, where the tip vortex is generated. For finite

airfoils, the effective angle of attack decreases towards the

tip (Brooks and Marcolini 1986; Awasthi etal. 2018), and

the tip flow field has influence over the entire span (Moreau

etal. 2018). This implicates that stall occurs at higher angles

of attack compared to a two-dimensional airfoil, because

new fluid from the pressure side can enter the boundary layer

on the suction side and the flow remains attached to the

airfoil surface for a longer downstream distance. This phe-

nomenon was also observed by Genç etal. (2018), where for

a NACA4412 with AR = 3 the separation of the shear layer

was suppressed by the formation of tip vortices, thus causing

a delay of stall for angles of attack up to 40

◦

and Reynolds

number 50,000. For higher Reynolds numbers, the finite

airfoil acted as a 2D airfoil because the effect of viscous

forces and leading edge vortices on the flow decreased. This

could also apply to the study by Riley (1951), where a higher

Reynolds number (

106

) was investigated and where the influ-

ence of the tip was decreased due to the higher aspect ratio

(AR = 4). Therefore, delayed stall cannot be seen in Fig.2a)

where all configurations achieve maximum lift at the same

angle of attack.

In contrast to the baseline case, it seems that for the end-

plate configurations the airfoil is in stall along the whole

part of the span that is shown. The application of end plates

seems to lead to more two-dimensional flow characteristics

Fig. 15 NACA4412 third octave band sound maps obtained at 15

◦

angle of attack for

125, 000

, top base configuration, bottom circular end

plate (nozzle exit at −0.386m, black dotted line represents airfoil, grey dashed line represents wall)

Experiments in Fluids (2021) 62:106

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Page 15 of 17 106

at this angle of attack. The areas of high turbulence intensity

above the trailing edge are assumed to be due to flow separa-

tion along the airfoil. The high turbulence intensity along the

trailing edge could also be the reason for the acoustic source

distribution that is seen in Fig.11. Both end-plate cases also

give reason to believe that some flow is still going around

the top of the end plate. This effect is stronger for the smaller

rectangular plate than for the circular end plate. The interac-

tion of this swept flow with the flow over the airfoil seems

to be small compared to the baseline configuration though,

because the area of high turbulence intensity extends up to

the end plate and is not affected.

The mean streamwise velocity is depicted in the second

row of Fig.16. The baseline configuration shows an area

at the tip where the mean flow velocity is 1.3times higher

than in the undisturbed flow. The location and direction of

this maximum confirm that it belongs to a vortex formed at

the tip. For the two end-plate configurations, a notable area

with low mean velocity is visible, which is stretched along

the span and located just below the high turbulence intensity

region observed in these cases. This is another indicator that

stall occurs, as it shows the remaining flow over the trailing

edge while most of the flow is separated due to stall. This

area is again more pronounced for the larger circular end

plate.

Fig. 16 Turbulence intensity, mean streamwise velocity U and mean

upwash velocity W for NACA 0012 airfoil, angle of attack 15

◦

125, 000

[measurement plane located 0.1chord lengths down-

stream of the airfoil, white dashed line represents the trailing edge,

wall is located at y = –140mm and airfoil tip at y = 0mm, base

configuration (left), circular end plate (center), rectangular end plate

(right)]

Experiments in Fluids (2021) 62:106

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106 Page 16 of 17

Finally, the mean flow velocity in the vertical direction,

perpendicular to the main flow and the trailing edge, is

shown in the bottom row of Fig.16. In the flow field on the

very left, belonging to the baseline configuration, it can be

seen that fluid from the pressure side of the airfoil is going

upwards to the suction side around the airfoil tip (positive

magnitude). Then, roughly 10mm inwards down the span,

the flow is going downwards again (negative magnitude).

This indicates a tip vortex formation from pressure to suc-

tion side. Some upwards flow is also observed around the

end plates as well as a roll-up around the top edge of the

plates and the subsequent forming of a smaller vortex. Based

on the absence of vortex induced noise in the acoustic results

obtained for the end-plate configurations, it appears, though,

that this vortex does not interact with the airfoil trailing edge

due to the increased vertical distance.

The results from the hot-wire measurements clearly show

the absence of an airfoil tip vortex for the end-plate con-

figurations. In the sound maps of Fig.11, no strong sound

sources were found at the tip for these configurations, and

the broadband frequency hump is not apparent in the spectra.

This confirms the tip noise reduction effect of the end plates.

4 Conclusions

This paper presents the experimental investigation of the

noise reduction effect of three different end-plate designs on

symmetric and cambered wall-mounted finite airfoils. The

measurements were taken in a small aeroacoustic wind tun-

nel using microphone array technique and advanced beam-

forming algorithms that enable the determination of both the

location and the strength of the noise sources. In addition,

constant temperature anemometry measurements were taken

on selected configurations to allow for a better understanding

of the flow structures present.

As a first result, beamforming sound maps revealed that

for the baseline configurations at high geometric angles of

attack a dominant sound source is located at the trailing edge

of the tip. These high-frequency broadband sound sources

are greatly reduced when the end plates are applied. Cor-

responding hot-wire measurements downstream of the tip

region showed a significant change of the flow field when

the end plates are applied. The results suggest that the end

plates greatly reduce the transfer of fluid from the pressure

to the suction side and, subsequently, diffuse the vorticity

at the tip. Thus, the influence of flow structures forming at

the free end of the airfoil is reduced and the flow field rather

resembles that of a two-dimensional configuration.

Corresponding sound pressurelevel spectra obtained

from the microphone array measurements further confirm

that the noise at the tip region can be reduced by applying

end plates to the free end of the airfoil. For the cases without

end plates, the spectra feature a notable broadband high-

frequency hump associated with the tip vortex formation.

This hump is not apparent for the end-plate configurations.

Since the hump is more pronounced in the spectra for the

asymmetric NACA4412 airfoil, the overall noise reduction

was found to be greater for this airfoil than for the symmetric

NACA0012 airfoil. The achieved noise reduction at the tip

is highest for higher angles of attack for both airfoils. For

the NACA4412 noise reduction was achieved for almost

all Reynolds numbers at geometric angles of attack from

2.5

◦

to 25

◦

except for the lower Re. The end plates on the

NACA0012 airfoil were effective in reducing noise at geo-

metric angles of attack from 15

◦

to 22.5

◦

. The results showed

that the overall sound pressure level can only be reduced if

the tip region is the dominant source region. Otherwise the

sound pressure levels are not affected or sometimes even

slightly increased by the presence of the end plates. For very

high angles of attack, the end plates lead to a noise reduction

over the whole frequency range.

The exact shape of the end plate did not have a strong

effect on the acoustic results although the circular end plate,

which has the largest area, showed the best results in noise

reduction. This dependency as well as the effects of end

plates on the flow over the airfoil surface and the resulting

aerodynamics is of interest for future research.

Acknowledgements Partial funding for this project was provided by

the German Academic Exchange Service (DAAD) under the Project-ID

57445107, which is gratefully acknowledged. The authors also thank

Tingyi Zhang and Charitha de Silva for their help with the flow visu-

alization experiments.

Funding Open Access funding enabled and organized by Projekt

DEAL.

Open Access This article is licensed under a Creative Commons Attri-

bution 4.0 International License, which permits use, sharing, adapta-

tion, distribution and reproduction in any medium or format, as long

as you give appropriate credit to the original author(s) and the source,

provide a link to the Creative Commons licence, and indicate if changes

were made. The images or other third party material in this article are

included in the article’s Creative Commons licence, unless indicated

otherwise in a credit line to the material. If material is not included in

the article’s Creative Commons licence and your intended use is not

permitted by statutory regulation or exceeds the permitted use, you will

need to obtain permission directly from the copyright holder. To view a

copy of this licence, visit http:// creat iveco mmons. org/ licen ses/ by/4. 0/.

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Authors and Affiliations

ErikSchneehagen1· ThomasF.Geyer2· EnnesSarradj1· DanielleJ.Moreau3

Thomas F. Geyer

thomas.ge[email protected]

Ennes Sarradj

ennes.sar[email protected]

Danielle J. Moreau

d.moreau@unsw.edu.au

1 Technische Universität Berlin, Einsteinufer 25, 10587Berlin,

Germany

2 Brandenburg University ofTechnology Cottbus-Senftenberg,

Siemens-Halske-Ring 15A, 03046Cottbus, Germany

3 University ofNew South Wales, Sydney, NSW2052,

Australia