Prediction of low-speed fan trailing-edge noise based on RANS and on scale resolved simulations

Transcription

1 Prediction of low-speed fan trailing-edge noise based on and on scale resolved simulations Gabriele Grasso Julien Christophe Christophe Schram von Karman Institute for Fluid Dynamics, Rhode-Saint-Genèse, 164, Belgium This work compares the prediction of the broadband trailing-edge noise emitted by a four-bladed low-speed ventilation fan obtained with different computational approaches. The objective of each approach is to compute the spectrum of the wall pressure fluctuations close to the trailing-edge, which is an input to Amiet s theory for the computation of the far-field noise spectrum. The wall pressure spectrum can be computed by applying Panton and Linebarger s theory to boundary layer data extracted from a time-averaged simulation, or it can be obtained directly by frequency analysis of a scale resolved simulation. The results of both kinds of computation are compared with the measurements taken on the same fan. A better correspondence with the experimental data has been obtained by using a Scale Adaptive Simulation (SAS) rather than a flow solution. I. Introduction The trailing-edge noise is the minimum broadband sound that a lifting surface would produce in absence of other sound mechanisms as turbulence interaction at the leading-edge or tip-vortex interaction. This source of sound, caused by the scattering of the boundary-layer disturbances into acoustic waves, as illustrated in Fig. 1, remains the only broadband noise contributor for subsonic fans operating in homogeneous stationary flows, in absence of any upstream, downstream and tip interaction. The present study aims at predicting the trailing-edge noise of a four-bladed low-speed axial fan on which an experimental campaign 1 has been conducted at CETIAT. The available experimental database is shared in a collaborative project including VKI, CETIAT 2 and CETIM, 3 aimed at the development and validation of numerical flow and noise prediction methods. In an earlier stage of this research, 4 the flow and acoustic prediction have been based on a method using stationary flow computation and semi-analytical Amiet s theory. 5 The flow was first solved using OpenFOAMR, whereby a 3D steady computation was performed using a Multiple Reference Frame method. The wall-pressure spectrum upstream the trailing-edge required in the trailing-edge noise theory was then obtained from a wall-pressure reconstruction model of Panton and Linebarger 6 using boundarylayer information. Furthermore, the statistical reconstruction of the wall-pressure spectrum and the noise computation have been validated with respect to experimental data on a flat plate boundary layer and on a controlled diffusion airfoil. 7 Finally, the far-field sound pressure spectrum was computed by means of Amiet s theory. This method keeps balance between cost and accuracy and for this reason it has already been implemented in an optimization procedure. 8 The purpose of this study is to perform a Scale Adaptive Simulation (SAS) of the ventilator in order to compute directly the trailing-edge wall-pressure spectrum from the time-resolved CFD data. The SAS model is based on unsteady but it can resolve turbulence structures with LES quality, providing a steady solution in stable flow regions while resolving turbulence in transient instabilities. 9 The results thereby obtained are compared with those previously obtained from data for what concerns the boundary layer characteristics, the wall pressure spectrum and the predicted radiated noise. Furthermore, a comparison of the last two with experimental data is included for validation of both the CFD approaches. PhD Candidate, Environmental and Applied Fluid Dynamics Department, AIAA Student Member. Corresponding author: 1 of 15

2 Boundary layer Sound scattering Turbulent eddies passing the trailing-edge Figure 1: Trailing-edge noise generation mechanism. II. Methodology for noise prediction The broadband trailing-edge noise of a stationary airfoil can be computed using Amiet s theory. 5 It assumes the airfoil to be infinitively thin, without camber or angle of attack, and in uniform flow conditions. Under such assumptions, the airfoil is approximated by a flat plate for the analytical derivation of the aeroacoustic transfer function. The main trailing-edge scattering obtained by Amiet 5 has been corrected by a leading-edge back-scattering contribution which accounts for the finite chord length. 1 The radiated sound field is calculated by integrating the induced surface sources on the actual chord length, c and the airfoil span, d, assuming convection of frozen turbulent boundary-layer eddies past the trailing-edge. In the following formulas, the subscripts 1, 2 and 3 indicate the stream-wise, cross-wise and span-wise directions, respectively. For large aspect ratio (d/c), the power spectral density of the sound pressure in the far field is obtained using: ( ) 2 ( ) ωcx3 d S pp (x, ω) = ω 2πc S 2 2 L kx2 2 Φ pp (ω)l y (ω), (1) U c S where ω is the angular frequency, x is the position of the listener, S is the distance from the source, L = L 1 + L 2 is the aeroacoustic transfer function, with L 1 the main contributing term from the trailingedge and L 2 the back-scattering term from the leading-edge. Equation (1) derives from the assumption of large-span airfoils, 11 implying that for each listener position there will be only one radiating span-wise aerodynamic wave number k y. Φ pp (ω) is the wall-pressure spectrum upstream the trailing-edge while l y (ω) is the corresponding span-wise correlation length. This quantity is modelled in this work by using Corcos formula, 12 ω/(bu c ) l y (ω) = ky 2 + (ω/(bu c )) 2 (2) where b is a model constant and U c is the convective velocity of turbulent structures at the trailing-edge. Φ pp (ω) is the main input to Amiet s theory and can be obtained either directly from a scale resolved simulation or from with a reconstruction model. In the second case, using the model developed by Panton and Linebarger, 6 Remmler et al. 7 implemented the following expression for the wall-pressure spectrum Φ pp (ω) = 8 ρ 2 k 1 (ω) 2 exp k(ω)(x2+ˆx2) k(ω) 2 S 22 (x 2, ˆx 2, ω) U 1 U 1 dx 2 dˆx 2 dk 3 (3) x 2 ˆx 2 where ρ is the reference density and x 2 is the wall normal distance. The energy spectrum of the vertical velocity fluctuations, S 22, is expressed as: S 22 (x 2, ˆx 2, ω) = ū 2 (x 2) ū 2 (ˆx 2) π 2 Λ 2 R 22 cos(α k 1 (ω) r 1 ) cos(α k 3 r 3 ) d r 1 d r 3 (4) gabriele.grasso@vki.ac.be Senior Research Engineer, Environmental and Applied Fluid Dynamics Department, AIAA Member. Associate Professor, Environmental and Applied Fluid Dynamics Department & Aeronautics and Aerospace Department, AIAA Member. 2 of 15

3 The integration coordinates and the wave numbers were non-dimensionalized with the turbulence integral length scale, Λ. The model therefore uses the stream-wise mean velocity profile U 1 (x 2 ) and the cross-wise velocity fluctuation profile u 2(x 2 ). Both velocities and the turbulence integral length scale Λ are calculated from the result. The velocity correlation coefficient R 22 and the scale anisotropy factor α need to be modelled. In particular, the cross-wise velocity correlation coefficient is modelled according to Panton and Linebarger as [ ] r 2 R 22 = 1 2 r e r 2 +ỹ 2 (5) 2 + ỹ 2 where r = r 1 + r 3 and ỹ = (x 2 ˆx 2 ) /Λ. This is a symmetric function, thereby implying local homogeneity of turbulence. No quadratures are used to calculate the quintuple integral in Eq. (3) as they would require prohibitive memory. The integration is performed with a Monte Carlo method using importance sampling to enhance convergence. In case of rotation, the far-field noise PSD of a low solidity fan with B independent blades is given by an integration over all possible azimuthal positions, ψ, of the stationary airfoil formulation: 13 S pp (X, ω) = B 2π 2π ω e (Ψ) ω SΨ pp(x, ω e )dψ (6) where X is the position of the listener in the absolute frame of reference and x is its position in the rotating frame attached to the airfoil trailing-edge. The factor ω e (Ψ)/ω accounts for Doppler effects due to the rotation, being ω e the emitted frequency and ω the observed frequency. In order to take into account the variation of the flow along the airfoil span, a strip theory is used, splitting the blade in a given number of segments. A local Cartesian frame of reference is defined for each strip close to the trailing-edge, as represented in Fig. 2(a), and boundary layer informations are extracted along the axis normal to the airfoil surface. Then, each strip is approximated by a flat rectangular surface (Fig. 2(b)) in order to comply with the hypothesis of Amiet s theory. A certain overlap can be observed between the surfaces approximating the lowest blade strip and the above strip respectively. This is due to the fact that these surfaces follow the shape of the trailing-edge of the blade which becomes more concave close to the hub. Finally, the total radiated sound is the summation of sound emitted by each rectangular surface. In the present case, the blade is split in 5 strips. III. Test case: four-bladed low-speed ventilation fan III.A. Experimental set-up The ventilator, shown in Fig. 3, has.8 m diameter and hub ratio of.34. It has four equally spaced blades, two of which are equipped with unsteady pressure sensors in order to measure the wall-pressure spectrum, Φ pp (ω), the span-wise correlation length spectrum, l y (ω) and the convective velocity of turbulent eddies U c. The chord at tip is of.135 m and the blade s maximum thickness is 4 mm, while the mean tip gap is 5 mm. The shroud trailing-edge is aligned with that of the blade when the stagger angle of the blade is set to 3. The shroud is mounted on the wall of the reverberant chamber, which is upstream of the ventilator. A rotating microphone performs the acoustic measurements in the reverberant chamber, while three fixed microphones are placed downstream of the ventilator, allowing to make a spatial average of the acoustic data. 1 III.B. III.B.1. Numerical set-up computation A three-dimensional steady computation of the flow around the low-speed ventilation fan has been performed by means of OpenFOAMR 2.3. The flow has been solved in the best efficiency condition for 6 RPM rotational speed. Since the relative Mach number at tip is approximately equal to.3, the incompressible solver simplefoam has been selected. 14 Figure 4 depicts the computational domain with the related boundary conditions. The computational domain is divided in three main regions: a stationary region upstream of the ventilator, containing the inlet 3 of 15

4 (a) Figure 2: Division of the blade surface in 5 radial strips: (a) definition of the local frame of reference for the extraction of boundary layer data; (b) definition of planes for the computation of radiated noise by means of Amiet s theory. (b) 4 of 15

5 Figure 3: View of the ventilator from the aspiration. The unsteady pressure sensors at mid-span and close to the tip are visible on the lower and upper blade, respectively. (Courtesy of CETIAT). boundary, a moving reference frame (MRF) region surrounding the ventilator and, finally, another stationary region downstream of the ventilator, containing the outlet boundary. Non-conformal interfaces (defined as cyclicami in OpenFOAMR) link the MRF with the stationary regions. Periodic boundary conditions have been applied to the lateral patches of the domain, allowing to simulate only one quarter of the fan geometry. The volumetric flow rate (3.5 m 3 s 1 ) has been imposed at the inlet a 1% turbulence intensity and.1 m turbulent length scale, while a fixed pressure has been imposed at the outlet. The selected turbulence model is k ω SST 15 with enhanced wall treatment. Finally, the stationary walls upstream and downstream of the ventilation fan, the hub, the blade surface and the shroud have been defined as no-slip wall. Inside the moving frame of reference, the hub and the blade are defined as stationary walls, while the shroud is counter-rotating with the speed of -6 RPM. The computational grid has been produced by means of the open source tool snappyhexmesh, which can generate three-dimensional meshes containing hexahedra and split-hexahedra automatically from triangulated surface geometries. SnappyHexMesh also allows to add a boundary layer mesh corresponding to selected surfaces, an option that has been applied to the blade/hub wall and to the shroud for the meshes used in the present study. A grid independence study has been performed in a previous work 4 by producing four different meshes having 1.5, 3.4, 7.2 and 13.9 millions of cells. As can be seen in Fig. 5a, the grid of 7.2 M cells is refined enough to have the non-dimensional wall distance y+ < 1 at the first cell on the blade surface. Figure 5b, on the other hand, shows that a smoother pressure coefficient profile can be obtained with the 13.9 M cells mesh. For these reasons, the results presented hereafter are obtained from the most refined mesh. III.B.2. Scale Adaptive Simulation The SAS method is a hybrid -LES approach which has been developed by Menter et al. 16 It is based on the k ω SST turbulence model with the addition of a production term in the ω equation which increases when the flow solution starts to go unsteady. The SAS production term is based on the ratio of the modelled turbulent length scale L = Cµ 1/4 k 1/2 ω, C µ =.9 (7) 5 of 15

6 Figure 4: Computational domain with boundary conditions. and the von Kármán length scale, which is in turn proportional to the ratio of first and second flow velocity gradients. The classical boundary layer formulation of the von Kármán length scale L vk,1d = κ U/ y U 2 / y 2 (8) is generalized for three-dimensional flows as L vk,3d = κ U U. (9) Since L vk,3d is smaller in unsteady flow region, when the flow equations resolve unsteadiness the SAS term is activated and increases the production of ω. This causes a decrease of the turbulent viscosity, in which expression ω appears in the denominator, and an increase in the magnitude of the destruction term in the modelled turbulent kinetic energy equation. For this reason, the solver switches to LES mode in transient unsteadiness. This approach is implemented in OpenFOAMR in the komegasstsas LES turbulence model. Since the MRF method adopted in the previous simulation is not suitable for time dependent cases, for the SAS simulation all the computational domain is included in a single rotating frame of reference (SRF). Consequently, the SRFPimpleFoam incompressible transient solver (operating in PISO mode) has been used. In order to ensure the stability of the numerical solution, the time step has been allowed to vary with the constraint of keeping the maximum CFL number below 1. The overall simulation time was of.12 s, being.1 s the duration of one blade rotation. The transient part of the simulation lasted approximately.29 s, after which the wall pressure has been sampled at a frequency of 57 khz, collecting 52 samples in total. The SAS data have been averaged in time over a period of.5 s. This averaging window corresponds to the time it takes for a fluid parcel to travel along the chord of the blade, from leading to trailing-edge. Finally, this computation has been performed on a cluster of 128 processors for approximately 24 hours. 6 of 15

7 y Pressure coefficient (- C p ) M 3.4 M 7.2 M 13.9 M Number of cells 1 6 Normalized arc length (a) y+ values vs. number of grid cells (b) Pressure coefficient for different grids at mid-span Figure 5: Assessment of the grid independence. 4 III.B.3. Flow field comparison Figure 6 represents the relative velocity field at three equally spaced radial sections close to the hub (6(a), 6(b)), at mid-span (6(c), 6(d)) and in the vicinity of the tip (6(e), 6(f)). This figure compares the results, on the left column, with SAS results, on the right column. A separation zone on the pressure side of the blade at the lower radius can be seen in both Figs. 6(a) and 6(b), which is the effect of the small axial extent of the hub. The relative velocity contours of Figs. 6(e), 6(f) suggest that SAS predicts earlier, with respect to, the separation of the flow towards the trailing-edge on the suction side of the tip section. Also the separation bubble appearing upstream, close to the leading-edge, is more important in the SAS prediction than in the. Figures 7(a) and 7(b) depict the trend of the wall pressure coefficient ( C p ) at the same radial positions as in Figs. 6(c), 6(d) and Figs. 6(e), 6(f), respectively. While at mid-span (7(a)), the onset of an adverse pressure gradient on the suction side is predicted at approximately 6% of the arc length by both methods, the prediction is rather different towards the tip (7(b)). While the simulation exhibits a favourable pressure gradient on the blade surface from 1% to 8% of the arc length, followed by a steep adverse pressure gradient unto the trailing-edge, the SAS simulation predicts a continuous decay of the pressure coefficient on all the suction side. This explains the bigger extension of the separated flow areas close to both the leading and the trailing-edge seen in Fig. 6(f). Finally, it is necessary to remind that the tip region is place of considerable three-dimensional effect related to the formation of a tip-clearance vortex and to its interaction with the blade and with the shroud s termination, which are arguably better detected by the SAS simulation. IV. Comparison of boundary layer characteristics extracted from and time averaged SAS As explained in Sec. II, boundary layer data are extracted, for each blade strip, on a line normal to the blade surface close to the trailing-edge in order to reconstruct the wall pressure spectrum by means of Panton and Linebarger s model. In summary, the three inputs to the model that are extracted from CFD are: the mean stream-wise velocity, U 1 (x 2 ); the averaged cross-wise velocity fluctuation, ū 2 (x 2); the turbulent length scale, Λ. In this section, the boundary layer data extracted from the solution will be compared with those extracted from the time averaged SAS computation. This comparison is done at two radial positions, namely 27 mm and 357 mm, which correspond to the pressure sensors visible in Fig of 15

8 (a) r = 162 mm. (b) r = 162 mm (c) r = 267 mm (d) r = 267 mm (e) r = 372 mm (f) r = 372 mm Figure 6: Relative velocity magnitude visualization at three equally spaced iso-radius planes. Left column: computation. Right column: time averaged SAS computation. IV.A. Mean stream-wise velocity The stream-wise mean velocity profile depicted in Fig. 8 provides the boundary layer thickness, δ, and the undisturbed external velocity, Ue. According to the values reported in Table 1, underestimates δ with respect to SAS. At the mid-span section, Ue is captured with little difference by the two methods, while close to the tip the external velocity predicted by SAS is significantly smaller, due to the complexity of the three-dimensional flow in that region. The presence of a recirculation zone at both radii is revealed by the plots of Fig. 8. While at mid-span the thickness of the recirculating flow region is predicted with a small discrepancy by the two methods, at tip the extension and entity of the recirculation are bigger in the SAS simulation. As in Eq. (3), the trailing-edge wall pressure spectrum depends on the derivative of the mean stream-wise velocity, which is visibly different in the two simulations. This accounts partially for the difference of the corresponding modelled wall pressure spectra, presented in Sec. V. 8 of 15

9 1 r = 267 mm 2.5 r = 372 mm.8 2 Pressure coefficient (-C p ) Pressure coefficient (-C p ) Normalized arc length (a) Normalized arc length Figure 7: Pressure coefficient ( C p ) over the blade sections corresponding to Figs. 6(c), 6(d) and Figs. 6(e), 6(f), respectively. The arc length is normalized with respect to the chord-wise coordinate. In the scale of the abscissa, represents the leading-edge and 1 represents the trailing-edge. (b) r = 27 mm r = 357 mm SAS SAS δ [m] U e [ms 1 ] Table 1: Boundary layer thickness, δ, and external flow velocity, U e. IV.B. Averaged cross-wise velocity fluctuation Φ pp is also a function of the energy spectrum of the cross-wise velocity fluctuations S 22, defined in Eq. (4), which is in its turn function of the distribution of averaged cross-wise velocity fluctuation, ū 2 (x 2). This quantity is not provided in the simulation by the k ω SST turbulence model, which computes instead the turbulent kinetic energy k =.5 ū i2. On the contrary, from a time averaged SAS computation it is possible to extract all the components of the Reynolds stress tensor, whose trace is proportional to the turbulent kinetic energy. In order to compute ū 2 (x 2) from the results, the turbulent kinetic energy is multiplied by an anisotropy factor defined as β 2 = ū 22 /(2k) and dependent on the distance from the wall. The β 2 factors used in this work were calculated by Remmler et al 7 from a computation of a flat plate boundary layer, using a Reynolds stress transport turbulence model (RSTM). First, a comparison can be made regarding the distribution of k in the cross-wise direction: as can be seen in Figs. 9(a) and 9(b), the values predicted in the simulation are one order of magnitude smaller that those predicted by the SAS method. Then, a comparison is made in Figs. 9(c) and 9(d) between the flat plate vertical anisotropy factor with the one computed directly from the Reynolds stress tensor given by the averaged SAS solution. In the boundary layer on the fan blade s suction side the turbulence is more anisotropic than in a zero pressure gradient boundary layer, as it was assumed in first approximation. Even though the β 2 factor used in the post-processing of the data is overestimated, the resulting ū 2 (x 2) is smaller than the corresponding value given directly by the SAS computation due to the different order of magnitude of the turbulent kinetic energy. 9 of 15

10 .25 r = 27 mm.25 r = 357 mm x 2 /C [-].1 x 2 /C [-] U [m/s] (a) U [m/s] Figure 8: Comparison of the stream-wise velocity profile extracted from and time averaged SAS computations at the positions of the experimental pressure sensors. (b) IV.C. Turbulent length scale The third quantity that is extracted from CFD simulations and given as an input to Panton and Linebarger s model is the turbulence integral length scale Λ. When inputs are used, this is simply considered proportional to the turbulent length scale defined in Eq. (7). When, on the other hand, averaged SAS inputs are used, following Panton and Linebarger Λ could be considered proportional to Prandtl s mixing length, l m, and computed directly with the expression u 1 Λ = 1.5 l m = 1.5 u 2 U 1 / x 2 ( U 1 / x 2 ). (1) However, Eq. (1) does not hold in presence of flow recirculation, which is found in this application. For this reason, it was decided to follow in first approximation Prandtl s simple rule for which κ x 2, for x 2 <.22 δ, l m = (11).9 δ, for x 2 >.22 δ. even though it is known to represent poorly boundary layers different from zero pressure gradient. V. Trailing-edge wall pressure spectra and far-field noise results The trailing-edge wall pressure spectrum has been computed, at two points on the blade surface corresponding to the position of the pressure sensors in the experimental set-up, by using different data sources. The Φ pp reconstruction method described in Sec. II has been applied to the boundary layer data, extracted from the and averaged SAS simulation, presented in the preceding section. Furthermore, Φ pp has been computed directly by frequency analysis of the SAS data. As it can be seen from Figs. 1(a) and 1(b), the based reconstruction fails to match the experimental curve by 2 to 4 db, which is due to the underestimation of the turbulent kinetic energy in the boundary layer. On the contrary, the data produce a spectrum that has a good match with the experiments up to 1.2 khz, but then decreases at higher frequencies. The reason for this decrease is still under investigation, but it can be reminded that the higher frequencies of the spectrum produced by Panton and Linebarger s method are influenced by the part of the boundary layer profile close to the wall. In this case, it is possible that the high frequency behaviour of the spectrum depends on the negative U 1 / x 2 values computed in the recirculating flow area close to the blade surface. 1 of 15

11 .25.2 r = 27 mm.25.2 r = 357 mm x 2 /C [-].1 x 2 /C [-] k [m 2 /s 2 ] (a) k [m 2 /s 2 ] (b).25.2 (flat plate) r = 27 mm.25.2 (flat plate) r = 357 mm x 2 /C [-].1 x 2 /C [-] β 2 β 2 (c) (d).25.2 r = 27 mm.25.2 r = 357 mm y/c [-].1 y/c [-] u 2 [m/s] u 2 [m/s] (e) Figure 9: Comparison of the turbulent kinetic energy (a,b), of the β 2 factors (c,d) and of the cross-wise average velocity fluctuation (e,f) extracted from and time averaged SAS computations at the positions of the experimental pressure sensors. (f) 11 of 15

12 The wall pressure spectra have been also computed by extracting the pressure fluctuations at the same location on the blade surface than where the boundary layer has been extracted. The power spectral density curves presented in Figs. 1(a) and 1(b) are obtained with Welch s averaging technique, dividing the sampled data in four non-overlapping segments and applying a Hanning window to each segment.in both plots Φ pp is overestimated at low frequencies and underestimated above 5-6 Hz. The accuracy of the prediction at low frequency can improve by adopting a longer sampling time. Finally, the experimental far-field sound power level (SWL) can be compared with the results obtained by applying Amiet s theory (Eqs. (1) and (6)), in Fig. 11. In order to compute the spectrum of the span-wise correlation length of the pressure fluctuations, l y (ω), with Corcos model (Eq. (2)), it is necessary to provide the value of the convective velocity of turbulent structures, U c, and of the constant b at each radius. For this, we used the results of the experimental campaign conducted by Rozenberg 17 on an experimental set-up made of the two blades equipped with pressure sensors, but without shroud. The ratio between the external and the convective velocity, α, with the coefficient b computed by Rozenberg are reported in Table 2. It is notable that the experimental wall pressure spectrum used in Amiet s theory produces a noise spectrum that corresponds to the experimental result only below 4 Hz, while above it decreases significantly. For this reason, we are led to assume the effect of other noise sources at high frequency, due for instance to the tip-clearance vortex. As for the results obtained with the data directly extracted from the SAS computation, the tendency to overestimation of the wall pressure spectrum below 5 Hz and to underestimation above, seen in the previous plots, is reflected in the noise computation. The best correspondence to the experimental data at low frequency is given by the Φ pp computed from averaged SAS results. Not surprisingly, the noise prediction based on results is far below the experimental curve, since it is directly linked to the corresponding wall pressure spectra. α = U e /U c r = 27 mm r = 357 mm.93.9 b Table 2: Parameters used in wall pressure spectra and noise computation: ratio between the external and the turbulent convection velocity, α, and Corcos model constant, b. 12 of 15

13 r = 27 mm SAS Experimental + PL + PL Φ pp (db/hz) Frequency (Hz) (a) r =357 mm SAS Experimental + PL + PL Φ pp (db/hz) Frequency (Hz) (b) Figure 1: Trailing-edge wall pressure spectra computed at mid-span and tip. 13 of 15

14 W] SWL 3 rd Octave [db ref SAS φ pp + AMIET Experimental SWL Experimental φ pp + AMIET Frequency [Hz] Figure 11: Far-field sound power level spectra, third octave average. VI. Conclusions and future work The Scale Adaptive Simulation performed in this study has helped understanding the causes of the offset found between the experimental wall pressure and noise spectra and the prediction based on a simulation of the flow field in the ventilation fan. In fact, the latter was not able to capture the magnitude of the turbulent kinetic energy in the boundary layer on the suction side of the blade, thus leading to an underestimation of the wall-normal velocity fluctuations. Furthermore, the statistical model for the reconstruction of Φ pp was developed for the case of attached boundary layers and has proved its effectiveness, for instance, when applied to a controlled diffusion airfoil. 7 However the related hypothesis are weak in the case of separated boundary layers, which are often found on ventilation fan blades. For this reason they will be reconsidered in order to better extend the wall pressure model to this kind of applications. So far only the wall pressure spectrum has been extracted from the time resolved SAS results. The convective velocity of turbulent structures and the span-wise correlation length could also be extracted by cross-correlating the pressure signals taken at different points of the blade surface. However, this would have required a much longer sampling period, thus increasing considerably the computational effort. Anyway, the computation of these data will be attempted in the successive stage of this work, which also involves a better understanding of the interaction of the tip-clearance vortex with the blade as a noise generating mechanism. Acknowledgements The Authors gratefully acknowledge the support of the EC under the project FlowAirS (Grant Agreement no ) and Alain Guédel and Mirela Robitu of CETIAT, Lyon, for providing the geometry of the fourbladed ventilator with the measured wall-pressure and noise spectra. References 1 Guédel, A., Robitu, M., and Dhaussy, C., Prévision du bruit de bord de fuite d un ventilateur hélicoïde à quatre pales, Tech. rep., CETIAT, Guédel, A. and Robitu, M., Prediction of the broadband noise of a low-speed axial fan by CFD simulations, International Conference on Fan Noise, Technology and Numerical Methods, Lyon, France, April Legros, M., Goth, Y., and Guédel, A., Fan noise prediction from local experimental source term and numerical sound propagation, International Conference on Fan Noise, Technology and Numerical Methods, Lyon, France, April Grasso, G., Christophe, J., and Schram, C., Broadband Trailing-Edge Noise Prediction of a Four-Bladed Axial Fan Using a Semi-Analytical Method, International Conference on Fan Noise, Technology and Numerical Methods, Lyon, France, April of 15

15 5 Amiet, R. K., Noise due to Turbulent Flow past a Trailing Edge, Journal of Sound and Vibration, Vol. 47, No. 3, 1976, pp Panton, R. L. and Linebarger, J. H., Wall Pressure Spectra Calculations for Equilibrium Boundary Layers, J. Fluid Mech., Vol. 65, No. 2, 1974, pp Remmler, S., Christophe, J., Anthoine, J., and Moreau, S., Computation of Wall-Pressure Spectra from Steady Flow Data for Noise Prediction, AIAA Journal, Vol. 48, No. 9, September 21, pp Grasso, G., Christophe, J., Schram, C., and Verstraete, T., Influence of the noise prediction model on the aeroacoustic optimization of a contra-rotating fan, 2th AIAA/CEAS Aeroacoustics Conference, Davidson, L., Evaluation of the SST-SAS model: channel flow, axysimmetric diffuser and axisymmetric hill, European Conference on Computational Fluid Dynamics - ECCOMAS CFD 26, TU Delft, The Netherlands. 1 Roger, M. and Moreau, S., Back-Scattering Correction and Further Extensions of Amiet s Trailing-Edge Noise Model. Part 1: Theory, Journal of Sound and Vibration, Vol. 286, 25, pp Christophe, J., Application of Hybrid Methods to High Frequency Aeroacoustics, Ph.D. thesis, Université Libre de Bruxelles - von Karman Institute for Fluid Dynamics, Corcos, G. M., The structure of the turbulent pressure field in boundary-layer flows, Journal of Fluid Mechanics, Vol. 18, 1964, pp Christophe, J., Sanjosé, M., and Moreau, S., Uncertainty Quantification of Low-Speed Axial Fan Self Noise, 14th International Symposium on Transport Phenomena and Dynamics of Rotating Machinery, OpenFOAM Foundation, OpenFOAM - The open source CFD toolbox. User Guide, version ed., December Menter, F. R., Two-Equation Eddy-Viscosity Turbulence Models for Engineering Applications, AIAA Journal, Vol. 32, No. 8, 1994, pp Menter, F. R. and Egorov, Y., The Scale-Adaptive simulation method for unsteady turbulent flow predictions. Part 1: theory and model description, Flow, Turbulence and Combustion,, No. 85, Rozenberg, Y., Modélisation analytique du bruit aérodynamique à large bande des machines tournantes: utilisation de calculs moyennés de mécanique des fluides, Ph.D. thesis, École Centrale de Lyon, of 15