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Development of Simulation Methodologies for Forced Mixers
Anastasios LyrintzisSchool of Aeronautics &
AstronauticsPurdue University
Acknowledgements
• Indiana 21st Century Research and Technology Fund
• Prof. Gregory Blaisdell • Rolls-Royce, Indianapolis (W. Dalton, Shaym
Neerarambam) • L. Garrison, C. Wright, A. Uzun, P-T. Lew
Motivation
• Airport noise regulations are becoming stricter.
• Jet exhaust noise is a major component of aircraft engine noise
• Lobe mixer geometry has an effect on the jet noise that needs to be investigated.
Methodology
• 3-D Large Eddy Simulation for Jet Aeroacoustics
• RANS for Forced Mixers• Coupling between LES and RANS
solutions• (Semi-empirical method)
3-D Large Eddy Simulation for Jet Aeroacoustics
Objective• Development and full validation of a
Computational Aeroacoustics (CAA) methodology for jet noise prediction using: A 3-D Large Eddy Simulation (LES) code
working on generalized curvilinear grids that provides time-accurate unsteady flow field data
A surface integral acoustics method using LES data for far-field noise computations
Numerical Methods for LES• 3-D Navier-Stokes equations• 6th-order accurate compact differencing scheme
for spatial derivatives• 6th-order spatial filtering for eliminating
instabilities from unresolved scales and mesh non-uniformities
• 4th-order Runge-Kutta time integration• Localized dynamic Smagorinsky subgrid-scale
(SGS) model for unresolved scales
Tam & Dong' s radiation boundary conditions
Tam & Dong' s radiation boundary conditions
Tam & Dong' soutflow boundaryconditions
Sponge zone
Tam &Dong' sradiationbcs
Vortex ring forcing
Computational Jet Noise Research• Some of the biggest jet noise computations:
Freund’s DNS for ReD = 3600, Mach 0.9 cold jet using 25.6 million grid points (1999)
Bogey and Bailly’s LES for ReD = 400,000, Mach 0.9 isothermal jets using 12.5 and 16.6 million grid points (2002, 2003)
• We studied a Mach 0.9 turbulent isothermal round jet at a Reynolds number of 100,000
• 12 million grid points used in our LES
Computation Details• Physical domain length of 60ro in streamwise
direction• Domain width and height are 40ro • 470x160x160 (12 million) grid points• Coarsest grid resolution: 170 times the local
Kolmogorov length scale• One month of run time on an IBM-SP using 160
processors to run 170,000 time steps• Can do the same simulation on the Compaq
Alphaserver Cluster at Pittsburgh Supercomputing Center in 10 days
x / ro
y/r
o
0 10 20 30 40 50 60 70-20
-10
0
10
20
30
40
z / ro
y/r
0
-20 -10 0 10 20-20
-15
-10
-5
0
5
10
15
20
x = 5ro
z / ro
y/r
0
-20 -10 0 10 20-20
-15
-10
-5
0
5
10
15
20
x = 15ro
z / ro
y/r
0
-20 -10 0 10 20-20
-15
-10
-5
0
5
10
15
20
x = 35ro
Mean Flow Results• Our mean flow results are compared with:
Experiments of Zaman for initially compressible jets (1998)
Experiment of Hussein et al. (1994) Incompressible round jet at ReD = 95,500 Experiment of Panchapakesan et al. (1993)
Incompressible round jet at ReD = 11,000
x / Dj
Uj/U
c(x)
0 10 20 300
0.5
1
1.5
2
2.5
3
3.5
4
4.5
5
slope = 0.161
From Zaman' sexperiments (1998):slope 0.155 for Mj = 0.9
Jet Mean Centerline Velocity Decay
x / Dj
Q(x
)/Q
e
10 15 20 25 304
5
6
7
8
9
10
11
slope = 0.267
From Zaman' sexperiments (1998):slope 0.26 for Mj = 0.9
Streamwise Mass Flux
slope = A = 0.092
experimental valuesof A : 0.086 - 0.096
x / ro
r 1/2(x
)/r o
0 5 10 15 20 25 30 35 40 45 50 55 600
0.5
1
1.5
2
2.5
3
3.5
4
4.5
5
5.5
6
Jet Half-Velocity Radius Growth
r / r1/2
u/U
c
0 0.5 1 1.5 2 2.50
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1x = 45rox = 50rox = 55roexp. data of Hussein et. al.exp. data of Panchapakesan et. al.
Mean Streamwise Velocity Profiles
r / r1/2
rx
0 0.5 1 1.5 2 2.50
0.005
0.01
0.015
0.02
0.025x = 45rox = 50rox = 55roexp. data of Hussein et. al.exp. data of Panchapakesan et. al.
rx = vx' vr' / Uc2
Reynolds Shear Stress Profiles
k1
E u(1)(k
1)
5 10 15 2010-7
10-6
10-5
10-4
10-3
10-2
10-1
100
k1-5/3
Grid cutoff
One-dimensional spectrum Eu(1) (k1) of vx'
at x = 20ro on the jet centerline
Jet Aeroacoustics• Noise sources located at the end of potential core• Far field noise is estimated by coupling near field
LES data with the Ffowcs Williams–Hawkings (FWH) method
• Overall sound pressure level values are computed along an arc located at 60ro from the jet nozzle
• Cut-off Strouhal number based on grid resolution is around 1.0
X
Y
Z
Control Surface
Control Surface
Jet Flow
x = 35 ro x = 45 ro x = 60 ro
30 ro
x / ro
y/r
o
0 10 20-5
0
5
10
15
R
• OASPL results are compared with: Experiment of Mollo-Christensen et al. (1964) Mach 0.9 round jet at ReD = 540,000 (cold jet) Experiment of Lush (1971)
Mach 0.88 round jet at ReD = 500,000 (cold jet) Experiment of Stromberg et al. (1980)
Mach 0.9 round jet at ReD =3,600 (cold jet) SAE ARP 876C database
Jet Aeroacoustics (continued)
(deg)
OAS
PL(d
B)
10 20 30 40 50 60 70 80 90 100 110 120100
102
104
106
108
110
112
114
116
118
120
LES + FWH (isothermal jet)SAE ARP 876C predictionexp. of Mollo-Christensen et al. (cold jet)exp. of Lush (cold jet)exp. of Stromberg et al. (cold jet)
St = f Dj / Uj
SPL
(dB
/St)
0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2 2.2 2.490
100
110
120
130
Our spectrum at x = 29ro and r = 12roBogey and Bailly' s spectrum at x = 29ro and r = 12ro
St = f Dj / Uj
SPL
(dB
/St)
0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2 2.2 2.490
100
110
120
130 Our spectrum at x = 11ro and r = 15roBogey and Bailly' s spectrum at x = 11ro and r = 15ro
Conclusions• Localized dynamic SGS model stable and
robust for the jet flows we are studying• Very good comparison of mean flow results
with experiments• Aeroacoustics results are encouraging• Valuable evidence towards the full
validation of our CAA methodology has been obtained
Near Future Work• Simulate Bogey and Bailly’s ReD = 400,000
jet test case using 16 million grid points 100,000 time steps to run About 150 hours of run time on the
Pittsburgh cluster using 200 processors• Compare results with those of Bogey and
Bailly to fully validate CAA methodology• Do a more detailed study of surface integral
acoustics methods
Can a realistic LES be done for ReD = 1,000,000 ?
• Assuming 50 million grid points provide sufficient resolution:
• 200,000 time steps to run• 30 days of computing time on the
Pittsburgh cluster using 256 processors• Only 3 days on a near-future computer that
is 10 times faster than the Pittsburgh cluster
Future Work
• Extend methodology to handle:– Noise from unresolved scales– Supersonic flow– Solid boundaries (lips)– Complicated (mixer) geometries
multi-block code
RANS for Forced Mixers
Objective
• Use RANS to study flow characteristics of various flow shapes
What is a Lobe Mixer?
Internally Forced Mixed Jet
Bypass Flow
Mixer
Core Flow
Nozzle
Tail Cone
Exhaust Flow
Exhaust / Ambient Mixing Layer
Lobed Mixer Mixing Layer
Forced Mixer
H
Lobe Penetration (Lobe Height)
H:
3-D Mesh
WIND Code options• 2nd order upwind scheme• 1.7 million/7 million grid points• 8-16 zones• 8-16 LINUX processors• Spalart-Allmaras/ SST turbulence model• Wall functions
Grid Dependence
Density Contours1.7 million grid points
Density Contours7 million grid points
Grid Dependence1.7 million grid points 7 million grid points
Density
VorticityMagnitude
Spalart-Allmaras and Menter SST Turbulence Models
Spalart-Allmaras
Menter SST
Spalart-Allmaras and and Menter SST at Nozzle Exit Plane
Spalart SST
Density
VorticityMagnitude
Mean Axial Velocity at x = 2.88”(High Penetration)
¼ Scale Spalartat x = 2.88/4”
experiment Spalart Allmaras
Mean Axial Velocity at x = 2.88”(High Penetration)
¼ Scale Menter SSTat x = 2.88/4”
experiment Menter SST
Spalart-Allmaras vs. Menter SST
• The Spalart-Allmaras model appears to be less dissipative. The vortex structure is sharper and the vorticity magnitude is higher at the nozzle exit.
• The Menter SST model appears to match experiments better, but the experimental grid is rather coarse and some of the finer flow structure may have been effectively filtered out.
• Still unclear which model is superior. No need to make a firm decision until several additional geometries are obtained.
Geometry at Mixer ExitLow Penetration Mid Penetration High Penetration
DENSITY CONTOURS (¼ Scale)
Low Penetration
Mid Penetration
Vorticity Magnitude at Nozzle Exit(¼ Scale Geometry)
Low Penetration Mid Penetration High Penetration
Turbulent Kinetic Energy at Nozzle Exit(¼ Scale Geometry)
Low Penetration Mid Penetration High Penetration
Preliminary Conclusions
• 1.7 million grid is adequate• Further work is needed comparing the
turbulence models and results for different penetration lengths
Future Work
• Analyze the flow fields and compare to experimental acoustic and flow-field data for additional mixer geometries.
• Further compare the two turbulence models.• If possible, develop qualitative relationship
between mean flow characteristics and acoustic performance.
Implementing RANS Inflow Boundary Conditions for 3-D
LES Jet Aeroacoustics
Objectives• Implement RANS solution and onto
3-D LES inflow BCs as initial conditions.
• Investigate the effect of RANS inflow conditions on:– Reynolds Stresses– Far-field sound generated
Implementation Method
• RANS grid too fine for LES grid to match.
• Since RANS grid has high resolution, linear interpolation will be used.
LES
RANS
Issues and Challenges• Accurate resolution of outgoing
vortex with LES grid.• Accurate resolution of shear layer
near nozzle lip.• May need to use an intermediate
Reynolds number eg. Re = 400,000
Final Conclusion
• Methodologies (LES, RANS, coupling) are being developed to study noise from forced mixers