bcfd predictions for the aiaa drag prediction …...title results_pw.ppt author joseph morrison...
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![Page 1: BCFD Predictions for the AIAA Drag Prediction …...Title Results_PW.ppt Author Joseph Morrison Created Date 6/27/2006 3:27:20 PM](https://reader034.vdocuments.site/reader034/viewer/2022050307/5f6fb338b1979211ac3f2608/html5/thumbnails/1.jpg)
BT_PW_no-icon_simple.ppt | 8/4/2006BOEING is a trademark of Boeing Management Company.Copyright © 2006 Boeing. All rights reserved.
BCFD Predictions for the3rd AIAA Drag Prediction Workshop(DPW3)Chad Winkler, Andy Dorgan, Mark Fisher, Mori ManiThe Boeing CompanySt. Louis, MO
S. P. VankaUniversity of Illinois at Urbana-ChampaignUrbana, IL
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BCFD Code Details
•Cell-centered, finite-volume approach
•HLLE flux calculation with second-order spatial reconstruction
•Linear preserving gradient calculation
•Fully implicit time integration
•Turbulence models
•Spallart-Allmaras
•SST
•Additional capabilities: Time accurate LES, real gas effects,hybrid structured/unstructured solver, additional fluxformulations available
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Grid Details
• Unstructured grids
•Mixed tetrahedra and prisms (boundary layer)
•Surface grids generated with MADCAP
•Volume grids generated with AFLR3
•Available on NASA FTP site
•Running on 64 bit Linux clusters
•Typical execution time : 24 hours on fine grid (33M cells)running on 33 processors
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F6 Wing/Body Grids
Coarse (~4M cells) Medium (~8M cells)
Fine (~33M cells)
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F6 Wing Root Region Grid
Coarse (~4M cells)
Medium (~8M cells)
Fine (~33M cells)
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F6 + FX2B Wing Root Region Grid
Coarse (~4M cells)
Medium (~8M cells)
Fine (~33M cells)
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Wing Root RegionSurface flow – Fine grid CL=0.5
•No separation seen on theF6+FX2B geometry wing root
•Separation seen on the F6geometry wing root
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CD - CL^2/(PI*AR)
CL
0.017 0.018 0.019 0.02 0.021 0.022 0.023 0.024 0.0250.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
F6+FX2B S-A
F6+FX2B SST
F6 S-A
F6 SST
Drag Polars
•Error bars represent magnitude ofoscillations of CL in the F6 solution
•F6+FX2B solutions saw littleoscillation
•SST model seen to predict ~10counts less drag than the S-A model
•FX2B fairing seen to reduce dragregardless of turbulence model
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Skin friction behavior
CD_Skin Friction
CL
0.01 0.011 0.012 0.013 0.014 0.0150.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
F6+FX2B S-A
F6+FX2B SST
F6 S-A
F6 SST
•Lower SST drag comes fromreduced viscous dragcontribution
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Grid convergence study
GRIDFAC = 1/(GRIDSIZE)^2/3
CD-CL^2/(PI*AR)
0 1E-05 2E-05 3E-05 4E-05 5E-050.017
0.018
0.019
0.02
0.021
0.022
0.023
0.024
F6+FX2B S-A
F6+FX2B SST
F6 S-A
•SST results seen to extrapolate toa lower drag value when comparedto S-A for the FX2B configuration
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Crinkle cut, F6+FX2B , S-A , Mach contours atBL=200mm
Isotropic tetrahedraquickly dissipate wake
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Wing Cp contours, F6+FX2B, S-A model
AoA = -3 -1 0 1.5
Top View
Bottom View
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Comparison of Cp between turbulence modelsCL=0.5, fine grid
(Cp_Spalart – Cp_SST)
Cp cut at BL=240.37mm
x/c
Cp
0 0.2 0.4 0.6 0.8 1
-1
-0.5
0
0.5
F6+FX2B, S-A
F6+FX2B, SST
Shock moves forwardin S-A solution
αS-A=0.119°αSST=0.166°
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Comparison of skin friction between turbulence models
(Cf_Spalart / Cf_SST)
CL=0.5, fine grid
•Localized regions of higher skin friction using S-A when compared to SST
αS-A=0.119°αSST=0.166°
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Summary
•Strong need for best-practices in unstructured grid generation – bothsurface and volume gridding
•Refine wake region using localized source nodes in volume gridgeneration
•Difficulty converging F6 cases (without fairing) for both turbulence models
•Turbulence model + grid dependencies•~10 counts drag difference predicted between S-A and SST models
•Refine grid further to remove any grid dependency on turbulence model
•Future plans•Alternate grids – highly resolved and selectively resolved grids, other DPW3grids
•Unsteady simulations
•Cross-code solution comparisons
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BT_PW_no-icon_simple.ppt | 16Copyright © 2006 Boeing. All rights reserved.