Combined Particle Image Velocimetry and acoustic measurements on a rod-airfoil configuration: Pilot test of the FLOVIST project

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UNCLASSIFIED

Executive summary

Nationaal Lucht- en Ruimtevaartlaboratorium

National Aerospace Laboratory NLR

Report no. NLR-TR-2008-377 Author(s) V. Lorenzoni M. Tuinstra F. Scarano Report classification UNCLASSIFIED Date September 2008 Knowledge area(s)

Aeroacoustic & Experimental Aerodynamics

Descriptor(s) Rod-Airfoil

Combined Particle Image Velocimetry and acoustic

measurements on a rod-airfoil configuration

Pilot test of the FLOVIST project

Left:Experimental setup

Right: PIV result at one time instant

Problem area

This report documents the results of a pilot test, part of the FLOVIST (Flow Visualization Inspired Aero-acoustics with Time-resolved Tomographic Particle Image Velocimetry) project, conducted by National Aerospace Laboratories (NLR) in collaboration with Delft University of Technology (TU Delft). The aim of the project is to predict the sound emission based on flow measurements.

Description of work

The purpose of the experimental campaign is to obtain the unsteady flow field around an airfoil downstream of a circular section

leading edge. Combined Particle Image Velocimetry (PIV) and acoustic measurements were performed and the results are presented and discussed.

Results and conclusions

A data set, containing both acoustic and PIV measurements, was acquired for a rod-airfoil

configuration at three different free-stream velocities: 10, 15 and 20 m/s.

Applicability

The PIV results will be used to predict the noise emission based solely on velocity measurements. As the sound emission was

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UNCLASSIFIED

Nationaal Lucht- en Ruimtevaartlaboratorium, National Aerospace Laboratory NLR

Combined Particle Image Velocimetry and acoustic measurements on a rod-airfoil configuration

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Nationaal Lucht- en Ruimtevaartlaboratorium

National Aerospace Laboratory NLR

NLR-TR-2008-377

Combined Particle Image Velocimetry and acoustic

measurements on a rod-airfoil configuration

Pilot test of the FLOVIST project

V. Lorenzoni

1

_{, M. Tuinstra and F. Scarano}

₁ 1_{TU Delft}

No part of this report may be reproduced and/or disclosed in any form or by any means without the prior written permission of NLR (and contributing partners).

Customer National Aerospace Laboratory NLR Contract number 6083205

Owner National Aerospace Laboratory NLR and TU Delft Division NLR Aerospace Vehicles

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NLR-TR-2008-377

Summary

This report documents the results of a pilot test, part of the FLOVIST (Flow Visualization Inspired Aero-acoustics with Time-resolved Tomographic Particle Image Velocimetry) project, conducted by National Aerospace Laboratories (NLR) in collaboration with Delft University of Technology (TU Delft). The purpose of the experimental campaign is to obtain the unsteady flow field around an airfoil downstream of a circular section rod, generating a Karman vortex wake impinging onto the airfoil leading edge. Combined PIV and acoustic measurements were performed and the results are presented and discussed. In a later stadium, the time resolved fluid dynamics results will be used to predict the sound emission by means of aero acoustic

analogies. The predicted sound field will then be compared with the results of the acoustic measurements.

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1 Introduction

Noise attenuation and suppression is becoming an increasingly important matter, as noise regulation for public environments as well as for industrial applications is becoming more demanding. Noise emission can be reduced once the sound sources are known and the noise generating mechanisms are understood. Aero acoustics studies the noise generation mechanisms by flows. The sound sources can be determined by means of an acoustic analogy. An analogy is a reformulation of basic conservation equations of fluid dynamics, with the wave equation on the left hand side of the equation and an acoustic source term, build up out of flow parameters, on the right hand side of the equation. Commonly, the source terms are computed by means of CFD techniques. However, in the present study it is attempted to construct the sound source terms by experimentally obtained time-resolved velocity data for a rod-airfoil configuration. For Reynolds numbers ranging from approximately 50 to 200000 the flow behind cylinders is structured as a sequence of counter–rotating vortices known as the Von Karman street. The phenomenology of vortex generation behind round cylinders has been widely described, e.g. by Williamson in [1]. When the shedding frequency is normalized with respect to the free-stream velocity and the cylinder diameter, it gives the Strouhal number. For the considered Reynolds number range the Strouhal number is approximately constant at a value of 0.21. The periodic impingement of vortices on the airfoil, located in the cylinder wake, generates rapid time variations of the surface pressure on the airfoil, which is responsible for the emission of a tone. This report documents the results of a pilot test, part of the FLOVIST (Flow Visualization Inspired Aero-acoustics with Time-resolved Tomographic Particle Image Velocimetry) project, conducted by National Aerospace Laboratories (NLR) in collaboration with Delft University of Technology (TU Delft). The purpose of the experimental campaign is to obtain the unsteady flow field around an airfoil downstream of a circular section rod, generating a Karman vortex wake impinging onto the airfoil leading edge. Combined PIV and acoustic measurements were performed. In a later stadium, the time resolved fluid dynamics results will be used to predict the sound emission by means of aeroacoustic analogies. The predicted sound field will then be compared with the results of the acoustic measurements.

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2 Experimental apparatus

2.1 Experimental facility

The experiment was carried out at NLR-NOP in the small anechoic facility (KAT), configured with a rectangular type nozzle. The nozzle has dimensions of 0.51 x 0.38 m2

. The anechoic

chamber is 5.5 x 5.5 x 2.5 m3 _{large and covered with 0.5 m long foam wedges, yielding 99 %}

acoustic waves absorption above 500 Hz.

A cylindrical rod was vertically mounted at an approximate distance of 22 cm (varying slightly case by case) from the wind tunnel nozzle. A NACA0012 plexiglas airfoil, with a chord of 0.1 m, was vertically placed in the wake of the rod at zero angle of attack. Both rod and airfoil were mounted on a rail system which allowed easy adjustment of the rod position and the rod-airfoil distance. The top edges of rod and rod-airfoil were held by a connection to the frame of the wind tunnel exit nozzle. A generic sketch of the experimental arrangement is shown in the following figure: 0.38 m 0.51 m d θ R = 1.25 m 0.25 m 0.10 m 0.332 m 0.102 m

Fig. 1 Schematic of the rod-airfoil configuration and dimensions for V = 15 m/s and a rod diameter of 6mm

(run 32 FOV 2CAM of the PIV test matrix in App. B)

Two different cylinders of 6 mm and 10 mm diameter were used in order to assess the effect on the sound generation, as a result of the variation of dimension and shedding frequency of the vortices. Each configuration was examined for three different free stream velocities of 10 m/s, 15 m/s and 20 m/s respectively.

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2.2 Acoustic measuring system setup

The first phase of the experiment consisted of a preliminary investigation that was carried out by means of acoustic measurements only. In this first phase a microphone array to conduct beam forming and eight regular far-field microphones were used. The circular microphone-array was placed 0.675m from the hart of the tunnel and was approximately 0.8 m in diameter. Two GBM-Viper systems, yielding a total number of 96 available channels, were used for the acoustic data acquisition. The 80 LinearX, M51 microphones used for the array were positioned in an unstructured manner to avoid spurious source detection. Eight microphones were used to determine the acoustic far-field. These were placed at half-span and perpendicular to the airfoil at a distance of 0.35 , 0.6 m, 0.85 m, 1.1 m, 1.35 m, 1.6 m, 2.1 m and 2.6 m as shown in figure 2.

Fig. 2 Laboratory configuration for acoustic measurements with far-field microphones and planar microphone array

In the second phase, acoustic measurements were carried out simultaneously with PIV

measurements. Four free field microphones were placed at a fixed distance of 1.25 m from the airfoil leading edge, at an angle, with respect to the airfoil chord, of 900_{, 117}0_{, 135}0_{, 143}0_{, as}

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NLR-TR-2008-377 mic 4 1170 R= 1 .25 m 1350 mic 1 1430 mic 2 mic 3 900 θ = 00 ₁₈₀0

Fig. 3 Far field microphones set up (left), top view schematic of microphones disposition (right)

On all array microphones wind shields were placed to prevent the acoustic measurements to be affected by air flow passing over the microphones. The far field microphones were covered with wind shields as well, to ensure that if there were to be an influence of a wind shield on the sound measurements, all microphones would be influenced similarly.

2.3 Flow measuring system setup

Two-component particle image velocimetry (2C-PIV) was used to obtain planar velocity field measurements around the airfoil. By use of a Safex stage smoke generator the flow was seeded with particles of approximately 1 μm diameter. The seeding was introduced at the inlet of the wind tunnel system (see Fig. 4-left), which provided a homogeneously seeded flow at the test section.

Fig. 4 Smoke generator placed at the wind tunnel inlet (left) and the laser and light sheet

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The seeding was illuminated by a Quantronix Darwin Duo Nd-YLF laser. The light sheet was introduced from the side, slightly at an angle upstream with respect to the test section axis as shown in figure 5.

Fig. 5 Overview picture of rod-airfoil / single camera configuration (left), illuminated region and shadow zones behind the airfoil (right)

During the acoustic measurements the rail system, the power box and laser generator were covered by absorbing foam in order to avoid possible acoustic reflections. The light sheet was placed at the middle span section of the airfoil. The transparent material of the airfoil allowed the measurement of the flow-field on the opposite side of the airfoil, with exception of two shadow zones, which were caused by light refraction through the model.

The laser was operated at a frequency of 2700 Hz in double-pulsed mode. The nominal pulse energy is 25 mJ at 1 kHz and is estimated to be 12 mJ at 2700 Hz. The laser pulses have a wavelength of 527 nm (visible green light) and duration of 120 ns. The time interval between the laser pulses was varied depending on the flow velocity in order to obtain a particle image displacement of approximately 10 pixels. The light sheet had a thickness of approximately 2 mm.

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Fig. 6 Laser orientation (left) and camera location (right) for span wise measurements

Two Photron Fast CAM SA1 CMOS cameras (1024x1024 pixels, 12-bit, 20x20 μm2 pixel size)

were used for the image acquisition. The cameras can operate at full frame at a maximum frame rate of 5.4 KHz and have a storage capacity of 8 GB, allowing the acquisition of a maximum of 5400 images (2700 pairs, 1 s) in a single measurement. The acquisition frequency was set to coincide with the frequency of the laser pulse couple in double frame acquisition mode (2700 Hz). Two cameras were employed to increase the spatial resolution. A slight inclination with respect to the vertical plane perpendicular to the airfoil cord (approximately 2o, Fig. 7) of the optical axes of the imaging was necessary in order to obtain overlapping fields of view.

Fig. 7 Two-cameras imaging configuration

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Each camera was equipped with a Nikon objective with a focal length either 105 or 60 mm in order to capture the required field of view. The numerical aperture was set to f# = 2.8. The synchronization between laser and camera was performed by means of a LaVision High-Speed controller and by the DAVIS 7.2 software package. Image processing was done with an iterative multi-grid window deformation technique implemented in the WIDIM software developed at TU Delft.

3 Results and discussion

This section discusses the most relevant results for both the acoustics and flow measurements. The first and the second part deal with acoustics and PIV measurements respectively.

3.1 Acoustic measurements 3.1.1 Far-field check

Eight inline microphones, perpendicular to the airfoil, were used to determine the acoustic far-field. For several measurements it was investigated how the sound pressure level decayed with distance from the source. Sufficiently far from the source, in the geometric far-field, a 1/R decay is expected. The following figure shows the narrow band spectra of the eight microphones, scaled to a reference distance of 1.25 m, for a rod-airfoil configuration with a 6 mm rod at V=20 m/s. [d B] ,Δ f= 1 2 .5 H z 40 60 80 R = 0.35 m R = 0.60m R = 0.85m R = 1.10m R = 1.35m R = 1.60m R = 2.10m R = 2.60m

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The sound pressure level at the peak frequency is the same for all the microphones. It can be seen that the lower frequency part of the spectra (250Hz and below) do not overlap well. This is most probably because at these frequencies background noise emanating from the wind tunnel nozzle exit is the dominant sound source.

Neglecting the acoustic energy contained in the lower frequencies, an over all sound pressure level (OASPL) is calculated at the different microphone positions:

Distance [m] OA S P L [d B ] 0 0.5 1 1.5 2 2.5 3 70 80 90 measured decay 1/R decay

Fig. 9 Theoretical 1/R decay (green) and measured decay (red)

In the figure the theoretical 1/R decay is compared with the calculated OASPL. It is seen that the decay matches well, even up to surprisingly close to the airfoil where the microphone is expected to be located in the acoustic near-field. From this result it is concluded that 1.25 m is sufficiently far to be considered the acoustic far-field.

3.1.2 Background noise

Figure 10 shows the measured background noise spectra for V = 10m/s, 15 m/s and 20 m/s at a distance of 1.35 m, perpendicular to the wind tunnel heart line.

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NLR-TR-2008-377 Frequency [Hz] SP L [dB] ,Δ f= 1 2 .5 H z 102 ₁₀3 ₁₀4 0 20 40 60 80 V = 0 m/s V = 10 m/s V = 15 m/s V = 20 m/s

Fig. 10 Background noise narrow band spectra for V = 0 m/s, 10 m/s, 15 m/s and 20 m/s at R = 1.35 m

At V = 10m/s a tone can be distinguished at approximately 487Hz which is believed to be associated with the wind tunnel fans. The low frequency content of the spectra is emanating from the nozzle, as can be clearly seen in the following source plots.

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NLR-TR-2008-377 X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 0.3 68 67 66 65 64 63 200Hz X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 0.3 66 65 64 63 62 61 250Hz X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 0.3 57 56 55 54 53 52 315Hz X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 0.3 52 51 50 49 48 47 400Hz X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 0.3 45 44 43 42 41 40 500Hz X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 0.3 45 44 43 42 41 40 630Hz X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 0.3 39 38 37 36 35 34 800Hz X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 0.3 37 36 35 34 33 32 1000Hz X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 0.3 39 38 37 36 35 34 1250Hz X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 0.3 34 33 32 31 30 29 1600Hz X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 0.3 35 34 33 32 31 30 2000Hz X [m] Y[ m ] -0.2 0 0.2 -0.2 -0.1 0 0.1 0.2 35 34 33 32 31 30 2500Hz

Fig. 11 Source plot of background noise for V = 15 m/s

It can be clearly seen that at the lowest frequencies the noise is emanated from the wind tunnel nozzle, which is situated at X = -0.35 m. The source spots in the top of the scan plane for frequencies higher than 1000 Hz are believed to be caused by flow interaction with the support structure.

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3.1.3 Airfoil alone noise

Figure 12 shows the airfoil-alone noise spectra for a free stream velocity of 10m/s, 15m/s and 20m/s at R = 1.35. It is noted that the airfoil itself emits tonal noise. This is a well known property of the NACA0012 airfoil which, at a zero degrees angle of attack forms a Von Karman vortex street behind the airfoil due to the hydrodynamic instabilities associated with a laminar separation bubble near the trailing edge. Interaction of the periodically shedding vortices with the trailing edge causes the emissions of a tone. A study on trailing edge noise prediction using time resolved PIV data have been recently published by the group of Schröder in Göttingen [3].

Frequency [Hz] SP L [d B ], Δf = 12. 5 H z 102 ₁₀3 ₁₀4 0 20 40 60 80 V = 10 m/s V = 15 m/s V = 20 m/s

Fig. 12 Airfoil alone noise narrow band spectra at V = 10, 15 and 20 m/s, R = 1.25 m

This interpretation of the spectrum is confirmed by the source plots of figure13 the airfoil alone noise measurement at V=15 m/s.

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Fig.13 Source plot of airfoil-alone noise at V=15 m/s

The 29th_{1/3-octave band, containing the primary peak at 760Hz shows a maximum noise level}

of 90dB and emanates from the trailing edge region. The second harmonic at 1530 Hz is distinguished as well and is also emanating from the trailing edge region; in fact at all 1/3-octave bands the dominant sound source is located at the trailing edge. It is noted that only at high frequencies above 4000 Hz, the part of the airfoil protruding through the wind tunnel shear layer causes a disturbance source of same order of the sound levels due to the airfoil noise. The relative circular shape of the source spots is caused by the fact that the line source on the trailing

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edge possesses a high degree of coherence. The high frequency content sources beyond 2000 Hz are slightly shifted downstream of the trailing edge by approximately 2.5 cm. This could be caused by either a measurement inaccuracy in the beam forming technique or the actual measurement of the dimensions of the experimental setup.

In order to ensure that the sound emanating from the rod-airfoil configuration is caused primarily as a result of the vortices impinging onto the leading edge, the airfoil was tripped. If the trailing edge is tripped the boundary layer is forced into turbulent state and no laminar separation occurs. As a consequence, the formation of vortices at the trailing edge is prevented and the tonal peaks of the acoustic spectrum disappear as shown in the plots of figure 14.

Frequency [Hz] SPL [d B], Δ f= 1 2 .5 H z 102 ₁₀3 ₁₀4 0 20 40 60 80 100 V = 20 m/s, laminar V = 20 m/s, tripped V = 15 m/s, laminar V = 15 m/s, tripped

Fig. 14 Rod-airfoil noise narrow band spectra at V = 15 and 20 m/s with tripped and laminar boundary layer, R =1.35 m

This effect of the trip wire on the source identification at 15 m/s free stream velocity is shown in the following source plot:

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Fig. 15 Source plot of airfoil-alone noise at V=15 m/s, boundary layer tripped by 0.4 mm trip wire at 10 % cord

It is seen that noise levels have been reduced significantly. The highest noise levels are achieved at low frequencies and are caused by wind tunnel background noise. At higher frequencies, ranging between 800 Hz and 2500 Hz the turbulent airfoil boundary layer is the dominant source of noise, generating sound as the turbulence eddies are conveyed past the trailing edge. However, it never reaches a sound pressure level higher than 44 dB, which is appreciably lower than the laminar boundary layer case. It is remarked that unlike the source plots of the tonal case, the source spot is line shaped as there is little span-wise correlation of the sound sources.

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3.1.4 Rod alone noise

A cylindrical rod placed in a flow is known to posses a shedding frequency related to a Strouhal number of 0.21. From this relation an expected shedding frequency is calculated:

Table 1 Expected rod-shedding frequency at St = 0.21

V \ d 0.004 0.006 0.01 10 525 350 210 15 750 525 315 20 1050 700 420 Frequency [Hz] SPL [d B ], Δf = 12. 5 H z 102 ₁₀3 ₁₀4 0 20 40 60 80 V = 10 m/s V = 15 m/s V = 20 m/s

Fig. 16 Rod-alone noise narrow band spectra for V = 10, 15 and 20 m/s, rod diameter = 6 mm, R = 1.25 m

Figure 16 shows good agreement with the expected Strouhal number. At 10 m/s, 15 m/s and 20 m/s a measured Strouhal number of 0.211, 0.205 and 0.202 is found. The first harmonic peak, corresponding to the Strouhal frequency of 350 Hz, for the 10 m/s case is not clearly visible but

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NLR-TR-2008-377 3.1.5 Rod-Airfoil configuration Frequency [Hz] SP L [d B ], Δf = 12. 5 H z 102 ₁₀3 ₁₀4 0 20 40 60 80 V = 10 m/s V = 15 m/s V = 20 m/s

Fig. 17 Rod-airfoil noise narrow band spectra for V = 10, 15 and 20 m/s, rod diameter = 6 mm distance rod-airfoil = 10.2 cm, R = 1.25

The noise produced by the airfoil in a rod airfoil configuration is mainly due to the impingement of the vortices, shed by the cylinder, on the airfoil leading edge. The noise peak frequency is expected to correspond to the rod shedding frequency. This is confirmed by the plot of figure 17. The first harmonic at 10 m/s, unlike figure 16, is now clearly visible. The peaks for all the velocities are shifted to slightly lower frequencies with respect to the rod alone case of figure 16. A possible explanation, proposed by Casalino in [5], is a hydrodynamic-acoustic feedback of the airfoil upon the cylinder shedding mechanism. The presence of the airfoil increases the intensity of the noise emitted by the rod of over 10 dB, as featured in figure 18. A shift of the peak frequency from 525 Hz to 500 Hz and broadening of the noise spectrum around it is noted.

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NLR-TR-2008-377 Frequency [Hz] SP L [d B ], Δf = 12. 5 H z 102 ₁₀3 ₁₀4 0 20 40 60 80 Rod-airfoil Rod alone

Fig. 18 Rod-alone versus rod-airfoil noise spectra for V = 15 m/s, rod diameter = 6 mm, distance rod-airfoil = 8.5 cm, R = 1.35 m

In figure 19 the source plots of the rod-airfoil configuration at V=15 m/s is shown. The maximum noise emission of about 78 dB occurs around 500 Hz and remarkably, appears to be generated in the airfoil trailing edge region as opposed to the expected leading edge region. No explanation for this was found as of yet. For higher frequencies the main source location is found at centre of the airfoil. From 3000 Hz the interaction of the model support with the tunnel shear layer becomes a dominant noise source. However, the attained noise levels are far below peak level.

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NLR-TR-2008-377 Frequency [Hz] SP L [d B ], Δf = 12. 5 H z 102 ₁₀3 ₁₀4 0 20 40 60 80 Laminar Tripped

Fig. 20 Rod-airfoil noise spectrum with tripped (green line) and non-tripped (red) airfoil at V = 15 m/s, 6 mm cylinder diameter. Rod airfoil distance 9.0 cm for tripped case and 10.2 cm for smooth airfoil; R = 1.35 m

Tripping the boundary layer has little effect on the spectrum shape of the rod-airfoil configuration. A difference is found in peak sound pressure level (~3 dB), however, this is attributed to the fact that the rod-airfoil distance has varied so the configuration is slightly altered. As the vortices have to travel further in the laminar case, they lose coherence and hence the sound production is expected to be less when impinging on the airfoil. The source plots (Fig. 21) also show qualitatively the same image and hence for the remaining experiments the trip was removed as it was deemed unnecessary.

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3.2 PIV measurements

3.2.1 Large field of view: overall flow survey

Field of View A (FoV A, Fig. 22) includes the rod and part of the airfoil. It gives an overall visualization of the vortex wake development and interaction with the airfoil. The field of view has dimensions of 18 x 18 cm2_{and the velocity measurement was performed with a spatial}

resolution of 1.4 mm (1.4 % chord), an overlap factor of 75 % and an integration window size of 31x31 px2. 18 cm 8.5 c m 8.5 cm 8 cm 1.5 cm

Fig. 22 Schematic of the overall field of view FOV A including cylinder and airfoil LE (left). Calibration image (right)

The following figure shows an instantaneous velocity field obtained with this FoV for V = 10 m/s and a rod diameter of 6 mm:

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The rod and the airfoil sections are indicated by black surfaces. The velocity measurement was not available in the patched regions because the flow could not be illuminated at this location. The origin of the reference frame has been chosen to coincide with the airfoil leading edge. The flow velocity increases along the sides of the cylinder and around the airfoil leading edge. The cylinder wake is visualized as a region of lower velocity. The instantaneous velocity vectors indicate a sinuous pattern for the instantaneous streamlines. In the next figure the vorticity field time evolution is shown for this case:

Fig. 24 Vorticity field time evolution. Recordings are separated in time by roughly 40 % of the shedding period; Sequence order: top-bottom left column-top bottom right column for increasing t*; V = 10 m/s, d = 6 mm

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The instantaneous vorticity sequence of figure 24 shows more clearly the presence of alternating vortices in the cylinder wake. t* is the recording time normalized with respect to the shedding period. Red (counter-clockwise vorticity) and blue (clockwise vorticity) blobs indicate the counter-rotating vortices. Vortices are shed by the cylinder at a rate of 350 Hz and are conveyed with a mean spacing of 2.5 cm and travelling with a velocity of approximately 7 m/s (70 % of the free stream velocity). The peak vorticity measured in this configuration is certainly

underestimated due to insufficient spatial resolution and therefore, the value of 2000 s-1_should

only be taken as a reference value. It is however, in good agreement with the reference vorticity obtained from dimensional analysis ωref = V∞/d = 1666 Hz. The vorticity significantly decays

downstream due to three-dimensional effects acting early at vortex formation as well as later due to interaction between neighbouring vortices. As a result one can observe that the vorticity blobs loose coherence while approaching the airfoil. Finally, when the vortices impact on the airfoil LE they are rapidly displaced along the airfoil contour, which causes further

degeneration.

3.2.2 View around the airfoil by two cameras

The FOV 2CAM view allows the visualization of the flow field all around the airfoil at high spatial resolution (Fig. 25 and Fig. 26). Vectors are spaced with space and time resolution of respectively: 0.68 mm (0.7 % cord) and 3.7e-4 s (20 % shedding period at 15 m/s). An overlap of 75 % and an integration window of 31 x 31 px2 were used.

8.3 c m 3.9 cm 0.9 cm 1.8 cm 3.8 cm 10 cm 8.3 cm

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Fig. 26 Calibration images camera A (left) camera B (right)

The camera views overlap by 9 mm as is shown in the figures above. The simultaneous images from the two-cameras have been merged together before PIV processing. The composed images have dimensions of 1939 x 1024 pixels, corresponding to a field of view of 16.39 x 8.3 cm2_.

3.2.2.1 RUN 32: Velocity = 15 m/s, cylinder = 6 mm

The airfoil was placed in the wake of the 6 mm diameter rod at a distance of 10.2 cm. The average flow velocity obtained from an ensemble of 1000 realizations corresponding to 185 shedding cycles is shown in figure 27. The stream wise velocity component and velocity vector profiles plotted in figure 27-top show a velocity deficit downstream of the cylinder. The cylinder wake appears to further expand due to the interaction with the airfoil upstream of the TE. A slight flow acceleration is visible around 30 % chord and the growth of a rather thick wake occurs downstream of the TE. The development of the airfoil boundary layer cannot be measured in the present configuration due to limits in the spatial resolution. The contours of vertical velocity component (Fig. 27-bottom) illustrate a strong flow deflection at the LE and a weaker one at the TE.

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Fig. 27 Mean velocity field. Velocity vector profiles and horizontal component contours (top). Vertical component and streamlines (bottom); V=15 m/s, d = 6 mm, distance rod- airfoil = 10.2 cm

The dynamical behaviour of the Karman wake interaction with the airfoil is described in figure 28 by a sequence of instantaneous velocity fields separated by 2*Δt = 740 μs (0.37 shedding cycle T) in which Δt is the images acquisition time Δt =1/2700.

The vortex street impinging on the airfoil LE is visualized by the contours of vertical velocity component, which exhibits patches of positive and negative values. The vortices however appear significantly less coherent than in the previous case as a mere result of the higher resolution of the measurement. It is well known that the Karman vortices are composed of smaller structures originating from the instability of the shear layer separating from the cylinder

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description of the vorticity field approaching the boundary layer. The values indicated in the figure 29 should be taken as reference.

Fig. 28 Velocity field time evolution. Velocity vector profiles and vertical component contours. t* recording separation time normalized with respect to the shedding time

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Fig. 29 Instantaneous vorticity contour (V = 15 m/s, rod diameter = 0.6 cm, distance rod-airfoil=8.5 cm)

Despite the three-dimensional degeneration of the vortices in the Karman wake a clear periodicity is observed just upstream of the LE. Close alternating upward (red) followed by a downward (blue) velocity indicate a clockwise rotating vortex. The vertical velocity fluctuations in the Karman vortices approaching the LE have a magnitude of approximately 4 m/s (27 % V∞), whereas after the interaction such fluctuation levels are shown to decrease to less than

15 % with respect to the free-stream velocity.

The interaction may be described as follows: counter-clockwise rotating vortices approaching the LE appear already slightly shifted towards the lower side of the wake axis. Conversely clockwise rotating vortices are slightly on the upper side. By approaching the LE the vortex is accelerated towards one of the two sides. In most cases counter-clockwise vortices shed from the lower side of the cylinder surface are passing on the lower side and the opposite occurs for the clockwise ones. As the flow is accelerated along the first 20 % of the airfoil chord the vortices further degenerate. However, in particular over the aft part of the airfoil, where an adverse pressure gradient is prevalent, a strong coherence loss occurs.

The vorticity pattern does not show presence of any Karman shedding emanating from the trailing edge, as expected to exist on this airfoil without rod interaction at zero incidence, see [2],[4]. This result is in agreement with the acoustic measurements, where the TE tonal noise

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NLR-TR-2008-377 20 40 60 80 100 120 140 160 180 200 220 20 40 60 80 100 120 B A C

Fig. 30: Points for correlation coefficients estimation: A in the cylinder wake, B outside the wake, C on airfoil lower side at 40 % chord

The normalized time correlation of the horizontal (Fig. 31-left) and vertical (Fig. 31-right) velocity components show an identical pattern. Point A has the highest correlation level. The amplitude of the correlation in point B, slightly outside the wake, is reduced by a factor 6 but still features the same periodic behaviour as point A. The velocity fluctuations loose coherence after interaction with the airfoil as indicated by autocorrelation of the signal at point C.

0 0.5 1 1.5 2 2.5 3 -1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1 Δ t [ms] Rxx h o riz o n ta l v e lo ci ty [m 2 /s 2 ] point A point B point C 0 0.5 1 1.5 2 2.5 3 -1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1 Δ t [ms] R xx ver tica l ve lo ci ty [ m 2/s 2] point A point B point C

Fig. 31 Time correlation function of the horizontal (left) and vertical (right) velocities components at point locations: A, B and C

The power spectrum in figure 32 shows a peak frequency of 498 Hz, which coincides with the measured acoustic peak frequency of the same rod-airfoil configuration at 15 m/s free stream as shown in figure 33.

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NLR-TR-2008-377 0 0.2 0.4 0.6 0.8 1 1.2 -35 -30 -25 -20 -15 -10 frequency (kHz) nor m al iz ed PS D ( dB /H z)

Fig. 32 Power spectrum point A

The sound spectra corresponding to the rod-airfoil configuration with 6 mm rod at 15 m/s are shown in the following figure 33. Microphone 1, 2, 3 and 4 correspond to angles of respectively 90 0_{, 117}0_{, 135} 0_{and 143}0_{with respect to the airfoil chord at 1.25 m distance, as indicated in}

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3.2.2.2 RUN 34: Velocity=10 m/s, cylinder = 10 mm

A 10 mm cylinder was placed at a distance of 10.2 cm in front of the airfoil. The vertical velocity contours and velocity vectors for a free stream velocity of 10 m/s are shown in figure 34.

Fig. 34 Instantaneous vertical velocity field, V= 10 m/s, cylinder diameter = 1 cm, distance rod-airfoil = 10.2 cm

Figure 34 shows the instantaneous vertical velocity field around the airfoil. The incoming vortex structures are visibly bigger than the 6 mm rod case. A quite high level of coherence of the vortices is kept until 60 % chord. An instantaneous vorticity contour is shown in the next figure:

Fig. 35: Instantaneous vorticity field. V= 10 m/s, cylinder diameter = 1 cm, distance rod-airfoil = 10.2 cm

The vorticity field shows higher degree of coherence with respect to the 6 mm cylinder case both before the impingement and during convection along the airfoil sides.

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Figure 36 shows the influence of the cylinder dimensions on the noise spectrum in a rod-airfoil configuration. The curves are obtained for a free stream velocity of 20 m/s for 6 mm and 10 mm rod diameters.

Fig. 36 Rod airfoil noise spectra for 6 mm and 10 mm diameter rods at V=20 m/s, distance rod airfoil = 9.7 for 10 mm rod and 10.2 for 6 mm rod

The peak frequency is shifted to a lower value as the diameter of the cylinder increases. The sound pressure level at peak frequency is believed to be higher for the 10 mm case because of the smaller distance between rod and airfoil and the higher degree of coherence and size of the vortices.

3.2.3 Span wise visualization

Span wise velocity measurements have been performed in order to estimate the span wise phase coherence of the line vortices impinging on the airfoil leading edge. The free stream velocity and cylinder diameter are the same as for the two-camera visualization, 15 m/s and 6 mm respectively. The distance between the rod and the airfoil is 10.8 cm. The field of view (FOV

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NLR-TR-2008-377 1.3 cm _{10.8 cm} 1.2 cm rod airfo il 13. 3 cm

Fig. 37 Sketch of FOV S-W.V (left) and calibration image (right)

The following plots show on the left hand side the horizontal and the right hand side the vertical instantaneous velocity component.

Fig. 38 Contours of the instantaneous horizontal (left) and vertical (right) velocity components

The cylinder and the airfoil are patched black. The vertical velocity corresponds to the out of plane motion of the two camera field of view. Its maximum value is about 25 % of the free stream velocity. The general flow motion in the cylinder wake is shown to be three dimensional. In the horizontal velocity contour span wise coherent structures are shown. The vertical velocity contour features alternating upward and downward displacement spots that increase in

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with the in-plane vertical velocity as seen in figure 28. To emphasise span wise coherence a sequence of the horizontal velocity is shown in figure 39.

Fig. 39 Time sequence of the horizontal velocity along the span wise direction at V =15 m/s with 6 mm cylinder diameter and distance rod-airfoil = 10.8 cm. t* is the recording time normalized with the shedding period

The above figure allows a visualization of the phase coherence of the vortex shedding along the span wise direction. The horizontal velocity contour indicates the presence of correlated

structures in the span wise plane. The images are phase locked at consecutive shedding periods in the limits of the given time resolution of the PIV recording. The red stripes of iso-velocity represent vortex lines at a fix phase angle. These appear to be slightly slanted with respect to the

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3.3 Trailing Edge Zoom

In addition to the rod-airfoil case, the airfoil alone case has been analyzed through a zoomed view around the trailing edge (Fig. 40). This allows for visualization of the Karman vortices generation-release process responsible for the tonal noise emission. The field of view

dimensions are 3.5 x 3.5 cm2_{. A spatial resolution of 0.26 mm (0.26 % cord) permits a partial}

visualization of the boundary layer which is estimated to be of the order of 1 mm at the trailing edge. An overlap of 75 % and an integration window size of 31x31 px2_{were used for the PIV}

processing. The nominal free stream velocity is 15 m/s. 3.5 cm 1.5 cm 1.5 c m 3.5 cm

Fig. 40 Sketch of FoV TE-ZOOM (left) and calibration image (right)

The following figures show the horizontal and vertical velocity component around the trailing edge. In these figures the airfoil is marked black, whereas the shadow zones and boundary regions are patched white.

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The vertical velocity on the right hand side shows that counter rotating vortices are shed at the airfoil trailing edge alternately from the upper and lower surfaces. Close to the lower surface at about 95 % of the chord, reversal flow can be noticed. This indicates a laminar separation bubble. It is remarked that this laminar separation bubble is only found at the bottom side of the airfoil. Therefore, it is concluded that only one side of the airfoil is required to have a laminar separation bubble in order for a Von Karman vortex street to form.

A time visualization of the development a vortex is given in figure 42.

Fig. 42 Time evolution of the vertical velocity around the trailing edge at V =15 m/s. Time interval between images t* is normalized with the vortex shedding period. Sequence: for increasing t*, left top-bottom, centre top-bottom, right top-bottom

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Fig. 43 Mean horizontal (left) and mean vertical (right) velocities for free stream velocity of 15 m/s

Vorticity is generated along the airfoil sides and agglomerate at the trailing edge into two well defined spots. The periodic vortex shedding at the trailing edge causes a tone to be emitted by the airfoil trailing edge (see Fig. 12).

Fig. 44 Instantaneous vorticity contour (left) and airfoil alone noise peak at V = 15 m/s (right) The width of vortex shedding region at the trailing edge is about 4 mm. An estimate of the sound emission frequency using a Strouhal of 0.08, in the range proposed by Lele in [2], scaled with the boundary layer thickness at the TE (half the shedding region width) gives a value of about 600 Hz. This is in fair good agreement with the recorded tone frequency (Fig. 44-right). The next sequence gives a close insight on the vorticity development and shedding process:

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Fig. 45 Vorticity field time evolution at V = 15 m/s. Time spacing t* is non-dimensionalized with

the shedding period

3.4 Trailing Edge Span

Span wise imaging has been performed at the trailing edge in order to visualize the coherence of the vortex release at the trailing edge in the airfoil alone configuration . The field of view (FoV TE-span, Fig. 46) is 12.8 X 12.8 cm2_{wide and allows for a spatial resolution of 1.03 mm (1 %}

chord). An overlap of 75 % and an integration window of 31x31px2_{are applied and the nominal}

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NLR-TR-2008-377 1.6 cm 11.2 cm ai rf oi l 12.8 cm

Fig. 46 Sketch of FOV TE-SPAN (left) and calibration image (right)

Figure 47 shows the vertical velocity component of the span wise TE-visualization.

Fig. 47 Vertical velocity contour (left) and vorticity contour (right)

The vertical velocity component is weaker in comparison to the cylinder shedding (seen in Fig. 38). This indicates a higher two-dimensional character of the trailing edge shedding vortices. Spots of opposite velocity are organized along vertical lines parallel to the trailing edge. These smear out with the distance from the airfoil without appreciable distortion up to 11 cm distance from the trailing edge.

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The horizontal velocity sequence of figure 48 allows for a direct comparison of span wise phase release between the trailing edge in the airfoil alone case and the rod shedding in the rod airfoil configuration, visualized in figure 39 for FOV-SPAN.

Fig. 48 Span wise visualization of the horizontal velocity behind the trailing edge in airfoil alone configuration at V =15 m/s. The contours for increasing t* are phase locked at the shedding period

Figure 48 shows a sequence of horizontal velocity contours at consecutive shedding periods. The horizontal velocity exhibits higher level of coherence in the vertical direction aligned with the trailing edge span compared to the cylinder shedding of figure 39. This suggests that Karman vortices release takes place, approximately, at the same distance from the trailing edge (see Fig. 47) all along the span. It is concluded that trailing edge vortex shedding exhibits higher

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4 Conclusions

Combined acoustic and PIV measurements were performed on a rod-airfoil configuration for three different free-stream velocities of respectively: 10, 15 and 20 m/s. The acquired velocity fields will be implemented as source terms in an model based on an acoustic analogy in order to predict the noise emission from PIV data.

The preliminary acoustic measurements, conducted with 8 in-line microphones, showed that a distance of 1.25 m from the airfoil was sufficiently far to be considered in the acoustic far field. The airfoil alone emits a tone due to the vortices released from the trailing edge as a

consequence of boundary layer near the trailing edge. The peak frequency can be related to the thickness of the detached region at the trailing edge through a Strouhal number approximately 0.08 as found in the literature. The trailing edge shedding mechanism showed a very high degree of coherence over the span of the airfoil.

Trailing edge tonal noise disappeared when the cylinder was placed in front of the airfoil. The presence of the cylinder caused the airfoil boundary layer to transit to a turbulent state, precluding laminar separation, therefore interrupting the tone generation mechanism. For the rod-airfoil configuration, the main process of noise generation is the impingement of vortices onto the airfoil leading edge. The frequency of the recorded sound coincided with the peak frequency of the horizontal and vertical velocity spectra in front of the airfoil. However, the source plot suggested the source is mainly located close to the trailing edge, instead of the leading edge as would be expected. It is believed this an unwanted effect caused by the relative low frequencies investigated in comparison with the array diameter. Another reason could lie in the fact that the vortex impingement noise is coherent with noise emission caused by the shedding of vortices from the cylinder which could cause an artificial shift of source location. The presence of the airfoil slightly reduced the frequency of the recorded sound with respect to the rod alone case, as also experienced by Casalino in his hybrid CAA investigation over the same configuration. A possible explanation can rely on a hydrodynamic-acoustic interference of the airfoil on the rod shedding mechanism itself. The noise level increased with more than 10 dB and the sound spectrum broadened around the peak in presence of the airfoil. The effect of tripping the boundary layer has been shown to be negligible in the case of rod airfoil configuration, whereas it completely changed the sound spectrum in the airfoil alone

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case. In the latter case the tonal peak and all the harmonics disappeared because the laminar hydrodynamic instability, which lies at the basis of this noise generation mechanism, is disrupted.

Phase locked sequence of horizontal velocity contours showed a slight variation of the shedding phase along the rod span. Iso-velocity lines assumed a tilde-like shape which further deform downstream towards the airfoil.

References

[1] Williamson, C. H. K.; Vortex Dynamics in the Cylinder Wake, Annual review of fluid mechanics, Vol. 28, 1996.

[2] Tim Colonius; Sanjiva K.Lele; Computational aeroacoustics: progress on nonlinear problems of sound generation, Progress in AEROSPACE SCIENCE, ELSEVIER, 2004. [3] Schröder, A.; Dierksheide, U.; Wolf, J.; Herr, M.; Kompenhans, J.; Investigation of

trailing-edge noise sources by means of high-speed PIV, 12th_{international Symposium on}

Application of Laser Techniques to Fluid Mechanics,Lisbon,2004

[4] Brooks, T.F., Hodgson, T.H.; Trailing edge noise prediction from measured surface pressures”, Journal of Sound and Vibration, 1981.

[5] Damiano Casalino Vortex Analytical and numerical methods in vortex-body aeroacoustics,

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Appendix A

Test matrix acoustic measurements

Measurement Test plan #1 Remarks

01 35 Silence measurement 02 14 Rod 6 + airfoil, b = 14.2cm 03 14 Rod 6 + airfoil, b = 9cm 04 14 Rod 6 + airfoil, b = 7.4cm 05 14 Rod 4 + airfoil, b = 8.2cm 06 36 Airfoil alone, V=10m/s 07 36 Airfoil alone, V=15m/s 08 36 Airfoil alone ,V=20m/s

09 36 Trip placed on airfoil at 20% chord, 10 m/s, 0.2mm 10 36 Trip placed on airfoil at 20% chord, 15 m/s, 0.2mm 11 36 Trip placed on airfoil at 20% chord, 20 m/s, 0.2mm 12 36 Trip placed on airfoil at 10% chord, 10 m/s, 0.4mm 13 36 Trip placed on airfoil at 10% chord, 15 m/s, 0.4mm 14 36 Trip placed on airfoil at 10% chord, 20 m/s, 0.4mm

15 14 Rod6 + tripped airfoil, b=9cm, V=10 m/s

18 14 Rod6 + tripped airfoil, b=9cm, V=45.7 m/s 19 14 Rod4 + tripped airfoil, b=9.8cm, V=20 m/s 20 14 Rod4 + tripped airfoil, b=9.8cm, V=46.1m/s

21 11 Rod4 alone, b=9.8cm, V=10m/s 22 12 Rod4 alone, b=9.8cm, V=15m/s 23 13 Rod4 alone, b=9.8cm, V=20m/s 24 05 Rod6 alone, b=9cm, V=10m/s 25 06 Rod6 alone, b=9cm, V=15m/s 26 07 Rod6 alone, b=9cm, V=20m/s 27 08 Rod10 alone, b=10.5, V= 10m/s

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32 04 Empty wind tunnel, V=20m/s

33 01 Empty wind tunnel, V=0m/s

34 01 Fixed gain setting, PIV_fixed_noff_4

35 01 Fixed gain setting, PIV_fixed_noff_5, 100x gain

36 01 Silence measurement, inc. PIV equipment

37 23 Rod6/Airfoil, V=10m/s , PIV run 24, 2 mics, steel plate to protect array

38 19 Rod6/Airfoil, V=20 m/s, PIV run 25, 2 mics

39 23 Rod6/Airfoil, V=10m/s, Seeding on, No PIV, without steel plate to check the influence on far field measurements

40 19 Rod6/Airfoil, V=20m/s

41 19 Rod6/Airfoil, V=10m/s, Seeding off, check if seeding has influence on the acoustic measurements

42 23 Rod6/Airfoil, V=20m/s

43 22 Rod6/Airfoil, V=15m/s, PIV run 26, 2 mics, no plate 44 20 Rod10/Airfoil, V=20 m/s, PIV run 28, mics

45 31 Background measurements

49 23 Rod6/Airfoil, V=10m/s, two camera setup, one camera not functioning

50 23 Rod6/Airfoil, V=10m/s, PIV run 31

53 20 Rod10/Airfoil, V=20m/s, one camera out of focus

55 26 Airfoil alone, V=15m/s

56 27 Airfoil alone, V=15m/s, with brush, seeding not on 57 27 Airfoil alone, V=15m/s, with brush, PIV run 36 58 26 Airfoil alone, V=15m/s, no brush, PIV run 37

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Appendix B

Test matrix PIV measurements

Run nr. Date Time Temp. Pressure Case Velocity Cyl. Diameter Rod-airfoil

(C) (Pa) (m/s) (mm) dist. (cm) 7 6/3/2008 14:27 13.6 102030 rod(6)-airfoil 10 6 7.9 8 6/3/2008 15:00 13.6 102030 rod(6)-airfoil 10 6 7.9 9 6/3/2008 16:35 13.6 102030 rod(6)-airfoil 20 6 7.9 10 6/3/2008 17:28 13.6 102030 rod(6)-airfoil 20 6 7.9 11 7/3/2008 10:10 12.6 101440 rod(6)-airfoil 10 6 7.9 12 7/3/2008 10:36 12.6 101440 rod(6)-airfoil 10 6 7.9 13 7/3/2008 15:18 12.6 101440 rod(6)-airfoil 10 6 7.9 14 7/3/2008 15:36 12.6 101440 rod(6)-airfoil 10 6 7.9 15 7/3/2008 16:36 12.6 101440 rod(6)-airfoil 10 6 7.9 16 7/3/2008 17:36 12.6 101440 rod(6)-airfoil 10 6 7.9 17 7/3/2008 18:36 12.6 101440 rod(6)-airfoil 10 6 7.9 18 8/3/2008 15:39 12.6 101440 rod(6)-airfoil 10 6 7.9 19 8/3/2008 15:40 12.6 101440 rod(6)-airfoil 10 6 7.9 20 8/3/2008 15:40 12.6 101440 rod(6)-airfoil 10 6 7.9 21 8/3/2008 16:03 12.6 101440 rod(6)-airfoil 20 6 7.9 22 10/3/2008 10:52 12.6 101440 rod(6)-airfoil 10 6 7.9 23 10/3/2008 11:00 12.6 101440 rod(6)-airfoil 20 6 7.9 24 10/3/2008 14:36 11 98200 rod-airfoil+2 mic 10 6 7.9 25 10/3/2008 14:38 11 98200 rod-airfoil+2mic 20 6 7.9 26 10/3/2008 14:40 11 98200 rod-airfoil+2mic 15 6 7.9 27 10/3/2008 17:28 11 98200 rod-airfoil+2mic 20 10 9.7 28 10/3/2008 17:35 11 98200 rod-airfoil+2mic 20 10 9.7 29 11/1/2008 15:13 11 98200 rod-airfoil+2cam 10 6 10.2 30 11/2/2008 16:13 11 98200 rod-airfoil+2cam 20 6 10.2 31 12/3/2008 15:30 12.3 98710 rod-airfoil+2cam 10 6 10.2 32 12/3/2008 16:40 12.3 98710 rod-airfoil+2cam 15 6 10.2 33 12/3/2008 16:52 12.3 98710 rod-airfoil+2cam 20 6 10.2 34 12/4/2008 17:32 12.3 98710 rod-airfoil+2cam 10 10 10.2

35 12/5/2008 18:03 12.3 98710 airfoil alone 15 N/A N/A

36 12/6/2008 18:03 12.3 98710 airfoil+brush 15 N/A N/A

37 12/7/2008 19:18 12.3 98710 airfoil 15 N/A N/A

38 13/3/2008 16:50 12.3 98710 zoom T.E. 15 N/A N/A

39 13/3/2008 17:00 12.3 98710 zoom T.E. 10 N/A N/A

40 14/3/2008 10:40 12.7 100930 span-visualization 10 6 10.8 41 14/3/2008 10:45 12.7 100930 span-visualization 10 6 10.8 42 14/3/2008 10:51 12.7 100930 span-visualization 15 6 10.8 43 14/3/2008 10:57 12.7 100930 span-visualization 20 6 10.8 44 14/3/2008 11:20 12.7 100930 span-visualization 10 10 10 45 14/3/2008 11:25 12.7 100930 span-visualization 10 10 10 46 14/3/2008 11:32 12.7 100930 span-visualization 10 10 10 47 14/3/2008 11:35 12.7 100930 span-visualization 15 10 10 48 14/3/2008 11:40 12.7 100930 span-visualization 20 10 10

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Run nr. Laser power Pulse separation FoV Specifications Acquisition Nr. Images

(A) ( μs) freq. (Hz) 7 25 120 A 2700 2700 8 25 100 A 2700 2700 9 25 60 A 2700 2700 10 25 60 A more focused 2700 2700 11 25 100 B 2700 2700 12 25 100 B laser closer 2700 1000 13 25 60 B 2700 1000 14 25 100 B laser thicker 2700 2000 15 25 60 B 2700 2000 16 25 60 B 2700 2000 17 25 100 B 2700 2000 18 25 60 B 2700 2000 19 25 80 B laser thin (1mm) 2700 1000 20 25 80 B 2700 1000 21 25 40 B 2700 1000 22 25 80 C 2700 1000 23 25 40 C 2700 1000 24 25 80 C 2700 2748 25 25 40 C 2700 2748 26 25 60 C 2700 2748 27 25 40 C 2700 1000 28 25 40 C 2700 2748

29 25 80 2CAM run..-A /run.._B 2700 1000

30 25 40 2CAM run..-A /run.._B 2700 1000

31 25 70 2CAM run..-A /run.._B 2700 2700

32 25 50 2CAM run..-A /run.._B 2700 2700

33 25 30 2CAM run..-A /run.._B 2700 2700

34 25 70 2CAM run..-A /run.._B 2700 2700

35 25 50 2700 2700 36 25 50 2700 2700 37 25 50 2700 2700 38 25 20 2700 1000 39 25 30 2700 1000 40 25 80 SP-VIS 1 camera 2700 1000 41 25 120 SP-VIS 2700 1000 42 25 80 SP-VIS 2700 1000 43 25 60 SP-VIS 2700 1000 44 25 120 SP-VIS 2700 1000 45 25 80 SP-VIS 2700 1000

46 25 100 SP-VIS laser thicker 2700 1000

47 25 80 SP-VIS 2700 1000 48 25 60 SP-VIS 2700 1000 49 25 60 SP-VIS 2700 1000 50 25 40 SP-VIS 2700 1000 51 25 20 SP-VIS 2700 1000 52 25 60 SP-VIS 2700 1000 53 25 40 SP-VIS 2700 1000 54 25 20 SP-VIS 2700 1000 55 25 60 ZOOM-TE 2700 1000 56 25 40 ZOOM-TE 2700 1000 57 25 20 ZOOM-TE 2700 1000

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Appendix C

Reference test plan

Phase2:

Nr Test Model Measurement D [m] V [m/s] Objective/description

19 Base

configuration

Rod6 +

Airfoil PIV/Acoustic 0.006 20

Obtain database for the base configuration 20 Airfoil/Rod1 0 Rod10 + Airfoil

PIV/Acoustic 0.010 20 Obtain database for d =

10 mm

21 Airfoil/Rod4 Rod4 +

Obtain database for d = 4 mm

22 Free Stream

Velocity 1

Rod6 +

Obtain database for V=15 m/s

23 Free Stream

Velocity 2

Rod6 +

Obtain database for V=10 m/s Phase3a:

24 Tonal Noise

AoA Airfoil Acoustic N/A 20

Determine AoA at which a tone is emitted

25 Tonal Noise

PIV Airfoil PIV N/A 20

Adjust PIV configuration for TE

26 Tonal Noise Airfoil PIV/Acoustic N/A 20 Obtain database tonal

noise case

27 Tonal Noise

+brush

Airfoil

+ brush PIV/Acoustic N/A 20

Assess the effect of brush Phase3b:

28 Rod Noise 1 Rod4 Acoustic/PIV 0.004 20 Obtain database

rod-alone case, D = 0.004 Obtain database

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Final background noise check:

31 Empty wind

tunnel None Acoustic N/A 0

Register wind tunnel noise

32 Empty wind

33 Empty wind

34 Empty wind

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Appendix D

PIV fields of view

4 cm

6.8 c

m

3 c

m

Field of view B sketch and calibration

5.5

cm

12 cm

0.5 cm 1.5 cm

Field of view C sketch and calibration

The images have been rotated before processing as indicated in the sketches.

The results of Fields of view B and C have not been presented since they provide no additional information to the already mentioned cases.

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The following table summarizes the PIV processing settings for the main fields of view. Table D.1 PIV processing settings

2 cameras F.O.V. Span Visualization Trailing edge zoom Fov: A Ensemble size 1000 1000 500 500 Overlap [%] 75 75 75 75

Final integration window [pixel] 31 x 31 31 x 31 31 x 31 31 x 31 Grid spacing [mm] (% chord) 0.68 (0.7 %) 1.31 ( 1.3 %) 0.26 (0.26 %) 1.45 (1.4 %)

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Appendix E

Microphone setup and location

Array microphones:

Channel nr. Id name Serial nr. x [m] y[m] z[m] Sensitivity [Pa/V]

5 AK01 162076 0.0167 0.675 -0.0167 56.30 6 AK02 162077 -0.0167 0.675 0.0500 84.43 7 AK03 162078 -0.0500 0.675 -0.0167 63.75 8 AK04 162079 -0.0167 0.675 -0.0500 53.52 9 AK05 162080 0.0500 0.675 -0.0167 61.09 10 AK06 162081 0.0500 0.675 0.0500 64.49 11 AK07 162082 -0.0500 0.675 0.1167 63.39 12 AK08 162083 -0.1167 0.675 0.0500 72.28 13 AK09 162084 -0.1167 0.675 -0.0167 60.88 14 AK10 162085 -0.1167 0.675 -0.0833 69.26 15 AK11 162086 -0.0500 0.675 -0.1167 70.55 16 AK12 162087 0.0167 0.675 -0.1167 69.34 17 AK13 162088 0.0833 0.675 -0.0833 62.81 18 AK14 162089 0.1167 0.675 -0.0167 76.21 19 AK15 162090 0.1167 0.675 0.0500 97.95 20 AK16 162091 0.0833 0.675 0.0833 76.47 21 AK17 162092 0.0167 0.675 0.1167 62.45 22 AK18 162093 -0.0167 0.675 0.2167 64.64 23 AK19 162094 -0.0833 0.675 0.1833 76.82 24 AK20 162095 -0.1500 0.675 0.1500 63.53 25 AK21 162096 -0.1833 0.675 0.0833 65.99 26 AK22 162097 -0.2167 0.675 -0.0167 64.05 27 AK23 162098 -0.1833 0.675 -0.0833 73.11 28 AK24 162099 -0.1500 0.675 -0.1500 63.10 29 AK25 162100 -0.0833 0.675 -0.1833 73.20 30 AK26 162101 0.0167 0.675 -0.2167 90.05 31 AK27 162102 0.0833 0.675 -0.1833 62.02 32 AK28 162103 0.1500 0.675 -0.1500 64.27 33 AK29 162104 0.1833 0.675 -0.0833 64.86 34 AK30 162105 0.2167 0.675 -0.0167 73.37 35 AK31 162106 0.1833 0.675 0.0833 74.99 36 AK32 162107 0.1500 0.675 0.1500 64.49 37 AK33 162108 0.0833 0.675 0.1833 59.63 38 AK34 162109 -0.0500 0.675 0.3167 67.14 39 AK35 162110 -0.1167 0.675 0.2833 62.52 40 AK36 162111 -0.1833 0.675 0.2167 67.45 41 AK37 162112 -0.2500 0.675 0.1500 87.30

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51 AK47 162122 0.2167 0.675 -0.1833 91.94 52 AK48 162123 0.2833 0.675 -0.1167 72.28 53 AK49 162124 0.3167 0.675 -0.0167 59.70 54 AK50 162125 0.2833 0.675 0.0500 64.64 55 AK51 162126 0.2833 0.675 0.1500 110.03 56 AK52 162127 0.2167 0.675 0.2167 72.61 57 AK53 162128 0.1500 0.675 0.2833 59.98 58 AK54 162129 0.0500 0.675 0.2833 55.34 59 AK55 162130 -0.0833 0.675 0.3833 58.08 60 AK56 162131 -0.1833 0.675 0.3500 62.52 61 AK57 162132 -0.2500 0.675 0.3167 61.38 62 AK58 162133 -0.3167 0.675 0.2500 56.49 63 AK59 162134 -0.3833 0.675 0.1500 66.53 64 AK60 162135 -0.3833 0.675 0.0500 61.94 65 AK61 162136 -0.3833 0.675 -0.0167 63.39 66 AK62 162137 -0.3833 0.675 -0.1167 59.57 67 AK63 162138 -0.3500 0.675 -0.2167 59.91 68 AK64 162139 -0.2833 0.675 -0.2833 78.07 69 AK65 162140 -0.2167 0.675 -0.3500 61.73 70 AK66 162141 -0.1167 0.675 -0.3833 58.21 71 AK67 162142 -0.0167 0.675 -0.3833 69.02 72 AK68 162143 0.0833 0.675 -0.3833 89.74 73 AK69 162144 0.1833 0.675 -0.3500 64.64 74 AK70 162145 0.2500 0.675 -0.3167 61.45 75 AK71 162146 0.3167 0.675 -0.2500 60.46 76 AK72 162147 0.3833 0.675 -0.1500 64.94 77 AK73 162148 0.3833 0.675 -0.0500 69.90 78 AK74 162149 0.3833 0.675 0.0167 67.07 79 AK75 162150 0.3833 0.675 0.1167 71.37 80 AK76 162151 0.3500 0.675 0.2167 67.14 81 AK77 162152 0.2833 0.675 0.2833 63.31 82 AK78 162153 0.2167 0.675 0.3500 62.37 83 AK79 162154 0.1167 0.675 0.3833 67.14 84 AK80 162155 0.0167 0.675 0.3833 61.09

Line array microphones:

85 AK84 162159 0.32 -0.35 0.0000 54.39 86 AK85 162160 0.32 -0.6 0.0000 59.43 87 AK86 162161 0.32 -0.85 0.0000 81.47 88 AK87 162162 0.32 -1.1 0.0000 69.26 89 AK88 162163 0.32 -1.35 0.0000 56.17 90 AK91 162166 0.32 -1.6 0.0000 66.22 91 AK92 162167 0.32 -2.1 0.0000 74.22 92 AK95 162170 0.32 -2.6 0.0000 69.50

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Far field microphones:

Channel nr. Id name Serial nr. x [m] y[m] z[m] Theta [º] Sensitivity [Pa/V]

85 AK84 162159 0.332 1.250 0 90 54.39

86 AK85 162160 0.899 1.114 0 117 59.43

87 AK86 162161 1.216 0.884 0 135 81.47

Combined Particle Image Velocimetry and acoustic measurements on a rod-airfoil configuration: Pilot test of the FLOVIST project

Executive summary

Nationaal Lucht- en Ruimtevaartlaboratorium

National Aerospace Laboratory NLR

Combined Particle Image Velocimetry and acoustic

measurements on a rod-airfoil configuration

Pilot test of the FLOVIST project

Nationaal Lucht- en Ruimtevaartlaboratorium

National Aerospace Laboratory NLR

NLR-TR-2008-377

Combined Particle Image Velocimetry and acoustic

measurements on a rod-airfoil configuration

Pilot test of the FLOVIST project

V. Lorenzoni

, M. Tuinstra and F. Scarano

Summary

Contents

1

Introduction

2

Experimental apparatus

3

Results and discussion

4

Conclusions

References

Appendix A

Test matrix acoustic measurements

Appendix B

Test matrix PIV measurements

Appendix C

Reference test plan

Appendix D

PIV fields of view

Appendix E

Microphone setup and location

Array microphones:

Line array microphones:

Far field microphones:

_{, M. Tuinstra and F. Scarano}