The Effect of Blade Thickness and Angle of Attack on Broadband Fan Noise

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1 The Effect of Blade Thickness and Angle of Attack on Broadband Fan Noise Stewart Glegg Florida Atlantic University and William Devenport Virginia Tech Work Supported by ONR, Prog. Manager: Dr. Ron Joslin Authors would like to thank Ed Envia for providing additional slides 1

2 Sources of Broadband Fan Noise Rotor Alone Rotor/Stator Interaction Trailing Edge Noise Blade Tip Gap Noise Duct Wall BL Inflow Noise Inflow Noise Rotor Turbulence Rotor Tip Vortices Duct Wall BL Self Noise AIAA Papers No: ,

3 Comparison of Predicted R/S Interaction Noise and Measurements for the SDT Fan Rig (Envia et al, 2008) Envia et al AIAA paper no

4 Prediction of R/S Interaction Noise with Rotor Alone Noise Added for the SDT Fan Rig (Envia et al, 2008) Envia et al AIAA paper no

5 Outline of Talk Rotor Alone Noise Predicting Trailing Edge Noise Sources Blade Surface Pressure Spectra and AoA Recent Developments Rotor Stator Interaction Noise Review of Existing Studies Physics of Leading Edge Noise AoA and Thickness Effects Wind Tunnel Measurements Of Leading Edge Noise Conclusions 5

6 Rotor Alone Noise 6

7 Sources of Rotor Alone Noise Tip Flows may dominate for low speed fans with large tip gaps and low AR Wall BL Inflow Noise -typically BL is very thin and source is not important Trailing Edge Noise -assumed to dominate for moderate AR blades with small tip gaps AIAA Paper No:

8 Predicting TE Noise 8

9 Trailing Edge Noise Single Blade BL Turbulence Sound Waves Sharp Trailing Edge Scattering of BL pressure fluctuations causes propagating acoustic waves Kutta condition is assumed at the trailing edge (Amiet (1977), Brooks(1984), Howe (1998) 9

10 Trailing Edge Noise Blade Row Sound Waves BL Turbulence In a blade row there is acoustic scattering from multiple blades AIAA Jnl Vol 36(9)

11 TE Noise Prediction Blade is broken down into sections Surface pressure wavenumber spectrum defined at each blade section Coupled Duct/ 3D TE Scattering Response defined for each blade section Spanwise wavenumber varies with radius Assumes small spanwise correlation lengthscale Blade Surface Pressure Blade Row Scattering Duct Mode Amplitudes AIAA Jnl Vol 36(9)

12 Properties of the Surface Pressure Assume that the boundary layer on a single blade has the same surface pressure spectrum as a blade in a cascade if U and δ* are the same AIAA Jnl Vol 36(9)

13 Deducing the Surface Pressure Spectrum for a given U and δ* Single Blade Blade Surface Pressure * TE Scattering = Far Field Sound Blade Row Blade Surface Pressure * Blade Row = Scattering Duct Mode Amplitudes AIAA Jnl Vol 36(9)

14 Source Spectra at Different Radii Source Power Spectrum db re 20e-6Pa/sqrt(Hz) Constant AoA of 7 deg blade radius frequency khz AIAA Jnl Vol 36(9)

15 Source Spectra at Different Radii Source Power Spectrum db re 20e-6Pa/sqrt(Hz) blade radius frequency khz AoA is 8 deg at hub and 4 deg at tip AIAA Jnl Vol 36(9)

16 Effect of Increasing AoA on Sound Power Output Sound Power Spectrum db re 1e-12W/Hz Increasing AoA frequency khz AoA constant from hub to tip AIAA Jnl Vol 36(9)

17 Effect of AoA on Total Sound Power Sound Power Output angle of attack re case A and B Sound power increases as ~2.4 db per degree AIAA Jnl Vol 36(9)

18 Recent Developments In collaboration with Bruce Morin, Oliver Attasi and Ramon Reba AIAA paper no

19 Using RANS to get surface pressure spectrum Surface pressure spectrum Both models give tke Local Linearized Unsteady Flow RANS Mean Flow Dissipation Scale AIAA paper no

20 SDT Test Comparison Rotor Alone Noise SDT Test Case Nominal VAN Dnstrm Data Dnstrm Pred RMG Dnstrm Pred Brooks Brooks Input 105 PWL (db) Data RANS Input F (Hz) The downstream sound power radiated from the fan blade versus frequency. The blue curve corresponds to the measured data, the green curve is the prediction based on the current model and the red curve is the prediction using the Brooks et al (1989) self noise database. AIAA paper no

21 Rotor Stator Interaction Noise 21

22 Rotor/Stator Interaction Noise Hanson and Horan (AIAA ) Glegg and Walker (AIAA ) Evers and Peake (JFM v.43, 2002) Nallasamy and Envia (JSV v. 283, 2005) Cheong et al (JASA 119, 2006) Attasi and Vinogradov (AIAA ) u.n U s b 22

23 Rotor/Stator Interaction Noise Effect of AoA Evers and Peake (JFM v.43, 2002) b Gave a high frequency analysis of loaded cascade, Showed effects of AoA for tonal noise from blade wakes was large ~10dB u.n U s However only a small effect ~2dB was found on broadband noise for a homogeneous turbulent inflow 23

24 Rotor/Stator Interaction Noise Fully Loaded Rotor Attasi and Vinogradov (AIAA ) Gave a complete analysis of loaded set of stator vanes in a circular duct Showed up to 3-4dB of real blade effects on broadband noise Predictions were similar to flat plate results at high frequencies Verified predictions using experimental data Concluded that real blade effects were small, but what is the physics?? 24

25 Leading Edge Noise 25

26 Phase of a Vortical Gust AIAA paper no

27 Streamwise Component of Goldstein s Gust AIAA paper no

28 Time Domain Methods Calculate the unsteady loading from a blade vortex interaction and hence the noise Arbitrary gust is modeled by multiple vortices Point vortex Works well for problems which can be reduced to 2D Glegg and Devenport JSV

29 Unsteady Loading from a BVI unsteady lift Unsteady Lift 10 deg 0 deg 5 deg Ut/a 0.5 Unsteady Drag unsteady drag 0 0 deg 5 deg 10 deg Ut/a Time Glegg and Devenport JSV

30 Response to Step Upwash Gust F(t) U Step Upwash Gust Glegg and Devenport JSV

31 Unsteady Loading from a Step Gust for 0,5,10 deg AoA 5 Net Unsteady Loading Unsteady Loading unsteady lift Ut/a Time drag 5 Glegg and Devenport JSV

32 Unsteady Loading from a Step Gust for thickness to chord ratios 0.001,0.06,0.15 Net Unsteady Loading 5 Unsteady Loading unsteady lift Ut/a Time 5 32 ag Glegg and Devenport JSV 2008

33 Conformal Mapping (a) Streamlines at an AoA α deg (b) Streamlines at zero AoA (a) Flow at an angle of attack For 2D incompressible flow the airfoil can be mapped onto a circle on the distance of the (b) Flow at zero a Unsteady loading depends vortex from the singular 1 point Location of the singular point depends Overlay of on (b) airfoil -1 onto thickness (a) rotated clockwise by 2α deg Glegg and Devenport JSV 2008

34 Conformal Mapping (a) Streamlines (a) Flow at at an an angle AoAof α attack(b) deg (b) Flow Streamlines at zero angle at zero of attack AoA After rotation streamlines are overlaid Increased levels from vortices in upper plane cancels reduced level from vortices in lower plane Overlay of (b) onto (a) rotated clockwise by 2α deg Glegg and Devenport JSV

35 Experimental Verification Experiments carried out at Virginia Tech in collaboration with William Devenport and Joshua Staubs 35

36 VT Stability Wind Tunnel Air exchange tower 0.5MW Drive and Fan AIAA Removeable Test section 6'x6'x24' test section 15' dia., 600 h.p. fan produces flow speeds to 80m/s (Re > per meter) Control room Anechoic system installed in control room area 36

37 AIAA paper no From Inside Loss (db) Frequency (Hz) 37

Turbulence Grid AIAA paper no 2008-3018 5cm bars on 30cm centers 69.4% open area ratio In contraction (area 32% larger than test section) 19.

38 Turbulence Grid AIAA paper no cm bars on 30cm centers 69.4% open area ratio In contraction (area 32% larger than test section) 19.5 mesh sizes ahead of airfoils 10 0 Longitudinal spectrum functions, large grid, 30m/s lf= Transverse corrrelation functions, large grid, 30m/s R uu (y) R ww (y) R uu (z) R vv (z) von Karman u'/u = 3.8% v'/u = 3.9% w'/u = 4.0% L = 0.082m E uu E uu (single wire) von Karman modified von Karman Separation (inches) k/k e 38

39 Measurement locations and conditions c α o NACA m -12,-8,-4,0,4,8,12 1.8m Kevlar acoustic window NACA m 0,2,4,6,8,10,12 NACA m 0,2,4,6,8,10,12 α 1.83 DU96 0.9m 0,2,4,6,8,10 S m 0,2,4,6,8,10 Foam transition All measurements at 30m/s (additional at 40m/s not shown) AIAA paper no

40 WJD1 0.2-m NACA 0012 Effects of Angle of Attack SPL 1/3 (db) L SPL 1/3 (db) o 4 o 8 o 12 o Frequency (Hz) Reduced Frequency ω r 21 AIAA paper no

41 Diapositiva 40 WJD1 Sync animation change Gershfeld to Howe William Devenport; 06/05/2008

42 0.2-m NACA 0012 AIAA paper no Spp (db) α eff 0 o 1.8 o 3.6 o 5.4 o Reduced Frequency Moreau et al. (2005) NACA 0012 h/c=1.3 (jet) L=6.6%c u'/u=5% U=40m/s SPL (db) α eff 0 o 12 o Reduced Frequency NACA 0012, 0.2-m h/c=9 L=40%c u'/u=3.9% U=30m/s 0 o 4 o 8 o 12 o

43 Noise Calculations Glegg, Devenport and Staubs (2008) Theory. Von Karman S pp 2 2 πω UR3 Lˆ( ω, k2) = Ω33( ω / U, k 2 2 2,0) 2 r c0 2R 3 ( ω) dk Sound from a flat plate Panel Method k Drift coordinates ω Wavenumber space SPL1/3(dB) Amiet Analytical Solution Reduced Frequency ω r

44 WJD2 0.2-m NACA 0012 Effects of Angle of Attack with predictions Flat plate L SPL 1/3 (db) o 4 o 8 o 12 o Frequency (Hz) Reduced Frequency ω r 21 AIAA paper no

45 Diapositiva 43 WJD2 Sync animation change Gershfeld to Howe William Devenport; 06/05/2008

46 0.6-m NACA 0015 Effects of Angle of Attack m 0012, 12 o 0 o 2 o 4 o 70 6 o o 8 L SPL 1/3 (db) m 0012, 0 o 10 o o Frequency (Hz) Reduced Frequency ω r 64 AIAA paper no NACA 0015, 0.6-m h/c=3 L=13%c u'/u=3.9% U=30m/s 44

47 0.6-m NACA 0015 Effects of Angle of Attack with predictions Flat Plate 0 o 2 o 4 o L SPL 1/3 (db) db 6 o o 8 10 o o Frequency (Hz) Reduced Frequency ω r 64 AIAA paper no

48 Conclusions Rotor alone noise and rotor stator interaction noise are equally important Trailing edge noise can be predicted from calculations of the blade surface pressure spectrum Blade stall dramatically increases trailing edge noise Angle of attack effects on blades in isotropic turbulence are small (but not necessarily for anisotropic turbulence and asymmetric blades) Blade thickness can significantly reduce high frequency noise (but the details of the blade shape are important) 46

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