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BOEING is a trademark of Boeing Management Company. Copyright © 2014 Boeing. All rights reserved. Numerical Modeling Approach for SWBLI in a High Aspect.

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Presentation on theme: "BOEING is a trademark of Boeing Management Company. Copyright © 2014 Boeing. All rights reserved. Numerical Modeling Approach for SWBLI in a High Aspect."— Presentation transcript:

1 BOEING is a trademark of Boeing Management Company. Copyright © 2014 Boeing. All rights reserved. Numerical Modeling Approach for SWBLI in a High Aspect Ratio Tunnel Matthew Lakebrink Mori Mani Boeing Research & Technology 7 th Annual Shock Wave/Boundary Layer Interaction Workshop May 6-7, 2014

2 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Project Objectives 1.To quantify, through numerical studies, the impact of various modeling approaches on CFD-predicted SWBLIs  Grid resolution  Computational domain similarity (Plenum vs. No Plenum)  Turbulence model (SST vs. SST-QCR)  Structured vs. Unstructured | 2 2.Validate numerical predictions against high quality experimental data  Schlieren imaging  Boundary layer profiles  Wall pressure distributions 3.Apply validated numerical approach to a parametric flow control investigation

3 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Experimental Test Setup | 3 Test Section Apparatus Figures from AIAA Test Section Design Experimental Facility Layout  Test section layout similar to that used by Bruce et al. 1  Width of 245mm was chosen to create a high aspect ratio (3.57) channel in order to isolate centerline SWBLI from corner effects M 1.6 -> 1.5

4 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. | 4 Computational Domain Symmetry boundary at spanwise centerline (full span is pictured) Initial wall normal spacing is everywhere 7e-5 inches Inlet (42”) Nozzle (30”) Test Section Choke Flap Shock Holding Plate LE Inflow 2 Separate Outflow Planes

5 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. General Simulation Parameters  Flow Solver: All simulations were conducted using BCFD (Boeing proprietary CFD code, structured and unstructured capability)  Roe scheme with 2 nd order spatial accuracy  Koren flux limiter  Turbulence Modeling: SST both with and without the QCR  Run Conditions:  Pt Inflow = 40.6 psi ; Tt = 358 R  M TestSection = 1.5  Unit Re TestSection ~ 1.85e6/in  Boundary Conditions  Walls modeled as no-slip and adiabatic  Zeroth order extrapolation used on outflow boundaries  Total pressure and total temperature held constant at inflow boundary | 5

6 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Grid Resolution Study | 6 Coarse: ~2.25e6 points Medium: ~9e6 points Fine: ~36e6 points Three images on right are axial station cuts 2.25” upstream of the shock holding plate  A nested family of three grids was used to assess solution sensitivity to grid resolution  Axial spacing was held constant for the study  All solutions were converged to similar levels based on Navier-Stokes and SST residuals, and integrated pressure and viscous loads Centerline cut depicting grid resolution in the vicinity of the shock holding plate

7 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Grid Resolution Study - Boundary Layer | 7  Noticeable differences in the predicted boundary layer exist between the coarse and medium grids  Difference in predicted boundary layer between medium and fine grids is very slight across the span

8 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Grid Resolution Study – Wall Pressure | 8  Upstream of plate…  The flow is characterized by isentropic compression and is largely insensitive to grid resolution  Downstream of plate…  The flow is characterized by regular oblique shock reflections and the wall pressure varies largely between coarse and medium grids, but is very similar between medium and fine grids Contours of |grad( ρ)| Holding Plate Leading Edge Wall pressure interrogation path is located on lower wall centerline. Domain Outflow Plate Leading Edge Plan View of Test Section

9 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Iso-Surface of u=-10ft/s Plenum/Inlet Interface Impact of Modeling the Supply Plenum Supply Plenum Flow Direction Centerline Cuts Flow Direction Upstream Downstream  The presence of the plenum significantly influences flow in the region upstream of the nozzle  Downstream of the nozzle, the choice to model the supply plenum is of no consequence | 9 With Plenum Without Plenum Inflow Plane

10 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Choke Flap 0° - SST  Oil flow is uniform across span  No detrimental corner flow  SHP generates oblique shock on top, and very weak oblique shock on bottom | 10 Cuts visible where Pt/Pt0 <= 0.98 Centerline Pseudo-Schlieren Bottom Wall Oil Flow x=0 is at the start of the test section Boundary Layer Visualization

11 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Choke Flap 0° - SSTQCR  Oil flow is uniform across span  No detrimental corner flow  SHP generates oblique shock on top, and very weak oblique shock on bottom  No significant difference from SST | 11 Centerline Pseudo-Schlieren Bottom Wall Oil Flow x=0 is at the start of the test section Boundary Layer Visualization Cuts visible where Pt/Pt0 <= 0.98

12 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Choke Flap 0° - Corner Flow Comparison | 12 x=0 is at the start of the test section SSTSSTQCR Turbulence induced secondary flow is predicted in a region tightly confined to the corner when the QCR is employed

13 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Choke Flap 4.5° - SST | 13 x=0 is at the start of the test section Boundary Layer Visualization Centerline Pseudo-Schlieren Cuts visible where Pt/Pt0 <= 0.98 Bottom Wall Oil Flow  Deflecting the choke flap results in the development of a lambda shock in the lower channel  Small recirculation region develops at the centerline  Large corner separation develops which influences centerline SWBLI

14 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Choke Flap 4.5° - SSTQCR x=0 is at the start of the test section Boundary Layer Visualization Bottom Wall Oil Flow Centerline Pseudo-Schlieren Cuts visible where Pt/Pt0 <= 0.98 | 14  Deflecting the choke flap results in the development of a lambda shock in the lower channel  Small recirculation region develops at the centerline  Large corner separation develops which influences centerline SWBLI  No significant difference from SST

15 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Structured VS Unstructured (SST) | 15 STRUCTURED Flow Direction UNSTRUCTURED

16 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Conclusions | 16 1.The flow approaching the SWBLI is sufficiently resolved on the fine grid 2.For choke flap angles of 0° and 4.5° there is little difference between SST and SST with the QCR 3.Solutions from the structured and unstructured solvers are very similar when using SST 4.Modeling the supply plenum is unnecessary if there is no interest in the flow field upstream of the nozzle 5.The simulations indicate that for the present tunnel width and boundary layer thickness, the shock induced corner separation influences the centerline SWBLI

17 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Future Work  Validate computational predictions  Wall pressure distributions  Schlieren  Characterize flow field as a function of choke-flap angle from fully supersonic to detached normal shock  Assess different methods for controlling shock induced separation at the centerline, and in the corners  Bleed  Micro-ramp VG | 17 Flow Direction Shock Holding Plate Side View Plan View of Bottom Wall Experimental Pressure Tap Locations Choking Flap

18 BOEING is a trademark of Boeing Management Company. Copyright © 2014 Boeing. All rights reserved. BACKUP

19 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. References 1.Bruce, P. J. K., Burton, D. M. F., Titchener, N. A., and Babinsky, H., “Corner Effect and Separation in Transonic Channel Flows,” Journal of Fluid Mechanics, Vol. 679, pp , | 19

20 Engineering, Operations & Technology Copyright © 2014 Boeing. All rights reserved. Choke Flap 4.5° - SSTQCR (continued) x=0 is at the start of the test section Δx 1 = 11.75” W = 9.65” δ 1 ~ 0.25” at x=4.5” (0.5” upstream of shock holding plate) This plot was modified from: Benek, Suchyta, and Babinsky, “The Effect of Wind Tunnel Size on Incident Shock Boundary Layer Interaction Experiments,” 6 th Annual NASA/USAF SWBLI Technical Interchange Meeting, April 2013 (0.026, 47) using Δx 1, δ 1 BCFD SST-QCR centerline sep. 47 δ 2 ~ 0.18” at x=0.25” (4.75” upstream of shock holding plate) (0.019, 65) using Δx 1, δ 2 65 Δx 2 ~ 1.2” is the centerline separation length (0.026, 4.8) using Δx 2, δ 1 (0.019, 6.7) using Δx 2, δ 2 BCFD SST-QCR corner sep. θ=11 β=50 θ=13 β=31 θ=8 β=26 | 20


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