DeBonis, Jim

Malik, Mujeeb

Rumsey, Chris

NASA Technical Reports Server

Jim DeBonis
Inlets and Nozzles Branch
NASA Glenn Research Center
Cleveland, Ohio
216‐433‐6581
james.r.debonis@nasa.gov
Test Cases for NASA’s Revolutionary 
Computational Aerosciences 
Technical Challenge
Part 1: Standard Test Cases for Separated Flows
Part 2: Standard Test Cases for Propulsion
AIAA SciTech, Kissimmee, FL, January 2015
Chris Rumsey & Mujeeb Malik
Computational AeroSciences Branch
NASA Langley Research Center
Hampton, Virginia
757‐864‐2165
c.l.rumsey@nasa.gov
m.r.malik@nasa.gov
https://ntrs.nasa.gov/search.jsp?R=20160006856 2019-08-31T02:41:15+00:00Z
2Introduction
• Revolutionary Computational Aerociences (RCA) Technical 
Challenge*:
– Identify and down‐select critical turbulence, transition, and numerical 
method technologies for 40% reduction in predictive error against 
standard test cases for turbulent separated flows, evolution of free shear 
flows and shock‐boundary layer interactions on state‐of‐the‐art high 
performance computing hardware.
– Timeframe: by 2017
• Discussions held with AIAA Turbulence Modeling Benchmark 
Working Group (TMBWG) and others
– Ideas for good (representative) cases
• Simple enough to be useful (avoid complex geometries that introduce 
uncertainties)
• Possess the relevant flow physics
– Ideas for evaluation metrics
– By defining “common” cases and metrics, unbiased evaluation of future 
model improvements will be easier
* see: https://www.aeronautics.nasa.gov/fap/aeronautical_sciences.html
3Introduction, cont’d
• Issue 1:
– There are too many cases out there, and everyone has their own favorite
• Solution 1:
– Highlight top 2‐3 representative “primary” cases for each category, but list many 
others as optional
• Issue 2:
– Many cases exist for which RANS “fails” but LES/DNS/hybrid “succeeds”
• Solution 2:
– Interpret differently: For RANS, look for 40% improvement in results; for 
LES/DNS/hybrid, look for 40% improvement in time‐to‐solution for those cases 
where the prediction is already good
• Issue 3:
– Impossible to agree on metrics; they are either too simplistic or too difficult to 
define
• Solution 3:
– Define some relevant metrics for the primary cases, but allow some leeway and 
use judgment
4Separated Flow Cases
2‐D NASA Hump
Greenblatt et al
• Rationale for: excellent high‐quality reference experimental 
data set; good 2‐D characteristics; includes both baseline and 
flow control; RANS known to do poorly; eddy‐resolving methods 
have been shown to do well; well‐vetted in previous workshop
• Rationale against: endplates introduced some blockage
5
M=0.1
Rec=0.936 million
- Greenblatt, D., Paschal, K. B., Yao, C.-S., Harris, J., Schaeffler, N. W., Washburn, A. E., 
“Experimental Investigation of Separation Control Part 1: Baseline and Steady Suction,” AIAA 
Journal, vol. 44, no. 12, pp. 2820-2830, 2006.
- Rumsey, C. L., “Turbulence Modeling Resource,” http://turbmodels.larc.nasa.gov (data posted 
online), and “CFD Validation of Synthetic Jets and Turbulent Separation Control,”
http://cfdval2004.larc.nasa.gov (data posted online).
2‐D NASA Hump (cont’d)
6
Baseline case: Typical RANS 35% error in bubble size (overprediction)
2‐D NASA Hump (cont’d)
7
Baseline case: Typical RANS turbulent shear stress too small in magnitude
in separated shear layer of bubble
2‐D NASA Hump (cont’d)
• General validation metrics:
– Separation and reattachment locations
– Turbulent shear stress profiles
– Velocity profiles
– Surface pressure and skin friction coefficients
• Specific validation metrics:
– From first two, can find bubble size relative to experiment
– Current typical RANS for baseline case:
8
((x / c)sep 0.665) / 0.665
((x / c)reattach 1.1) /1.1
 ( u v /U 2ref )min@x/c0.8  0.020  / 0.020
 ( u v /U 2ref )min@x/c0.8 0.020  / 0.020  45%
(x / c)reattach  (x / c)sep  0.435  / 0.435  35%
Axisymmetric Transonic Bump
Bachalo & Johnson
• Rationale for: includes shock‐induced separation; widely‐used 
dataset for many years; axisymmetry removes 2‐D questions; 
RANS OK for some aspects but poor for others
• Rationale against: old experiment; no experience yet with 
eddy‐resolving methods
9
M=0.875
Rec=2.763 million
- Bachalo, W. D., Johnson, D. A., “Transonic, Turbulent Boundary-Layer Separation Generated on 
an Axisymmetric Flow Model,” AIAA Journal, vol. 24, no. 3, pp. 437-443, 1986.
- Rumsey, C. L., “Turbulence Modeling Resource,” http://turbmodels.larc.nasa.gov (data posted 
online).
Axisymmetric Transonic Bump (cont’d)
10
Typical RANS:  Cp can be reasonable, but 20-30% error in bubble size 
(overprediction)
Axisymmetric Transonic Bump (cont’d)
11
Typical RANS: turbulent shear stress too small in magnitude
in separated shear layer of bubble
Axisymmetric Transonic Bump (cont’d)
• General validation metrics:
– Separation and reattachment locations
– Turbulent shear stress profiles
– Velocity profiles
– Surface pressure coefficients
• Specific validation metrics:
– From first two, can find bubble size relative to experiment
– Current typical RANS:
12
((x / c)sep 0.7) / 0.7
((x / c)reattach 1.1) /1.1
 ( u v /U 2ref )min@x/c1.0 0.019  / 0.019
 ( u v /U 2ref )min@x/c1.0  0.019  / 0.019  50%
(x / c)reattach  (x / c)sep  0.4  / 0.4  25%
Additional Optional Separated Cases
• 2‐D Wake flow in APG (Driver & Mateer, 2002)
• 2‐D Transonic Diffuser (Sajben, 1983)
• 2‐D Ramp flow (Cuvier et al, 2014)
• 2‐D Periodic Hill (Almeida et al, 1993; Frohlich et al, 2005)
• 2‐D Planar Asymmetric Diffuser (Buice & Eaton, 2000)
• 2‐D NACA 4412 at high AoA (Coles & Wadcock, 1979 & 1987)
• 2‐D Curved Backward‐Facing Step (Bentaleb et al, 2012)
• 2‐D Separated Channel flow (Marquillie et al, 2008)
13
Additional Optional Separated Cases
• 3‐D Axisymmetric Hill (Byun & Simpson, 1996)
• 3‐D FAITH Hill (Bell et al, 2006)
• 3‐D Cherry Diffuser (Cherry et al, 2008)
• 3‐D ONERA M6 Wing (Schmitt & Charpin, 1979)
• 3‐D NASA Trapezoidal Wing (Johnson et al, 2000)
• 3‐D Common Research Model (Rivers et al, 2012)
• 3‐D Wing‐body‐juncture (Devenport & Simpson, 1992)
• 3‐D Prolate Spheroid (Chesnakas & Simpson, 1996)
14
15
Proposed Fundamental Experiments Investigating Separation 
Onset and Progression
Plots showing effect of top wall 
shape on separation bubble 
behind smooth bump
-5 deg wall
5 deg wall
12.5 deg wall
no separation
small separation
large separation
Surface skin friction
Turbulent shear stress
Axisymmetric Bump 3-D Juncture Flow
Vary alpha to achieve attached, 
incipient, and separated side‐of‐
body corner flow
16
Propulsion Cases
Compressible Mixing Layer
17
• Goebel, Dutton, & Gruber 
Univ. of Illinois
• Two convective Mach 
numbers
– Case 2, Mc = 0.46
– Case 4, Mc = 0.86
- Goebel, S.G. and Dutton, J.C., “Experimental Study of Compressible Turbulent Mixing Layers,” AIAA 
Journal, vol. 29, no. 4, pp. 538-546, April, 1991
- *Goebel, S.G. “An Experimental Investigation of Compressible Turbulent Mixing Layers,” Ph.D. 
Thesis, Dept. of Mech. and Ind. Eng., Univ. of Illinois., Urbana, Ill., 1990.
- Gruber, M.R. and Dutton, J.C., “Three-Dimensional Velocity Measurements in a Turbulent 
Compressible Mixing Layer,” AIAA Paper 92-3544, July 1992.
*primary data source
• Data
– LDV
• Mean velocities
• Reynolds stresses
– Growth rates
– Schlieren
Compressible Mixing Layer
• Objective: Improve the prediction of shear layers, including 
the developing region, and the effect of compressibility
• Metrics
– Growth rate
– Mean velocity fields
– Reynolds stress fields
• Rationale for
– Comprehensive data set for Mc’s from 0.2 to 1.0
– 2‐component LDV, with some 3‐component LDV and PIV
– Data compares well against historical measurements
• Rationale against
– High Mc case difficult to converge
– Tunnel wall effects, blockage and shock reflections are present
• Need new experiment?
18
Compressible Mixing Layer
19
Growth Rate
Round Jet
• Bridges & Wernet, NASA Glenn
• Three test conditions
– Set point 3, Mach 0.5 cold
– Set point 23, Mach 0.5 hot
– Set point 7, Mach 0.9 cold
• Data
– Temporally resolved PIV data
• Mean velocities
• Reynolds stresses
– Farfield noise
20
- Brown, C.A., and Bridges, J., “Small Hot Jet Acoustic Rig Validation,” NASA TM-214234, April, 2006. 
- *Bridges, J., Wernet, M.P., “The NASA Subsonic Jet Particle Image Velocimetry (PIV) Dataset,” 
NASA TM 2011-216807, Nov. 2011
- Bridges, J. and Wernet, M.P., “Validating Large-Eddy Simulation for Jet Aeroacoustics,” Journal of 
Propulsion and Power, vol. 28, no. 2, pp. 226-234, March-April, 2012.
*primary data source
Round Jet
• Objective: Improvement the prediction of jets, including the developing 
region, the effect of compressibility, and the effect of temperature.
• Metrics
– Location of the end of the inviscid core flow (the point where u/Ujet = 0.98 )
– Value and location of the peak turbulent kinetic energy on the jet centerline.
– Mean velocity fields
– Reynolds stress fields
• Rationale for
– Data already widely used for validation (de facto standard jet data set)
– Data compares against historical measurements
• Rationale against
– Nozzle boundary layer and/or initial jet profile not defined 
21
Round Jet
22
SP 3: Mach 0.5, cold
Set
Point Expt. SA %err SST %err
3 6.3 6.73 6.83 8.67 37.6
7 7.8 6.84 ‐12.3 9.01 15.5
23 5.0 6.00 20.0 7.65 53.0
End of invisicid core, x/Dj
Centerline Mean Velocities
SP 7: Mach 0.9, cold
SP 23: Mach 0.5, hot
Shock Wave Boundary Layer Interaction
• J. Brown et al, NASA Ames
• Axisymmetric compression corner
• Mach 2.85
• 30 deg. conical flare
• Data
– LDV
• Mean velocities
• Reynolds stresses
– Surface static pressures
– Interferometry
– Schlieren
23
- Dunagan, S.E., Brown, J.L. and Miles, J.B. ,” Interferometric Data for a Shock/Wave Boundary-Layer 
Interaction,” NASA TM 88227, Sept. 1986.
- Brown, J.D., Brown, J.L. and Kussoy, M.I., “A Documentations of Two- and Three-Dimensional 
Shock-Separated Turbulent Boundary Layers,” NASA TM 101008, July, 1988.
- *Settles, G.S., and Dodson, L.J., “Hypersonic Shock/Boundary-Layer Interaction Database NASA CR 
177577, April 1991
- Wideman, J., Brown, J., Miles, J., and Ozcan, O., “Surface Documentation of a 3-D Supersonic 
Shock-Wave/Boundary-Layer Interaction,” NASA TM 108824, 1994
*primary data source
Shock Wave Boundary Layer Interaction
• Objective: Improve the prediction of shock boundary layer interaction 
including the extent of separation and the Reynolds stresses
• Metrics
– Separation and reattachment locations, and separation length 
– Surface pressure coefficient
– Velocity profiles
– Reynolds stress profiles
• Rationale for
– Simple geometry
– Absence of corner flows found in many SWBLI experiments
– Reynolds stresses available
• Rationale against
– Separation onset and pressure distributions are already well predicted by 
some RANS models
– Cusped nose of model not defined; effect is assumed negligible
24
Shock Wave Boundary Layer Interaction
25
Mach number contours
Surface static pressure
Separation
pt. (cm)
Reattach
pt. (cm)
Length
(cm)
Expt. ‐2.73 0.97 3.70
BSL ‐2.50 1.38 3.88
SST ‐4.57 2.54 7.11
k‐ε ‐2.21 1.13 3.33
SA ‐2.68 2.27 4.92
ε
∞
Shock Wave Boundary Layer Interaction
26
Shear Stress Profiles
ε
∞
Additional/Optional Cases
• Shear flows
– Numerous shear layer cases
– Seiner’s Supersonic jet, Mj = 2 (Seiner et al, 1992)
– Egger’s Supersonic jet (Eggers, 1996)
• Shock‐wave/boundary‐layer interaction
– Mach 2.25 Impinging SWBLI, UFAST (DuPont et al, 2008)
– Mach 5 Impinging SWBLI (Schulein, 2004)
– Mach 7 Axisymmetric Compression Corner (Kussoy & Horstman, 1989)
– NASA GRC Mach 2.5 Axisymmetric Impinging SWBLI (Davis, TBD)
27
Turbulence CFD Validation Experiments (TCFDVE)
PROBLEM
Very few Shock Wave/Boundary-Layer Interaction (SWBLI) experiments reported in the open literature meet the rigorous criteria required 
to be considered as a CFD validation dataset. This is particularly true for experiments with detailed turbulence measurements.
OBJECTIVES
Obtain mean and turbulence quantities through a M=2.5 SWBLI of sufficient quantity and quality to be considered as a CFD validation 
dataset. Initial efforts will focus on a Mach 2.5 2-D (in the mean) interaction with follow-on efforts investigating 3-D interactions. Both 
attached and separated interactions will be considered.
APPROACH
A new M=2.5 17cm axisymmetric facility is being constructed to investigate SWBLIs. The facility will be located in Test Cell W6B at NASA 
GRC. The SWBLI is generated by a cone located on the centerline of the facility. The strength of the interaction is varied by changing the 
cone angle. The measurement region of interest is where the conical shock interacts with the naturally occurring facility boundary-layer and 
is highlighted by the box shown in Figure 1. The new facility will be instrumented with conventional pressure instrumentation as well as hot-
wire anemometry for measurement of turbulence quantities. Non-intrusive optical techniques such as PIV will be incorporated in the future. 
Test are also planned with dynamic surface shear film and fast response Pressure Sensitive Paint (PSP) in collaboration with Innovative 
Scientific Solutions, Incorporated (ISSI).
POC: David O. Davis (GRC)  
Figure 1. 17cm Axisymmetric Supersonic Wind Tunnel
RESULTS
The new facility design is complete and 
delivery of the hardware is expected by the 
end of December 2014. Calibration of the 
facility is expected to commence in 
December. RANS and LES simulations of 
the facility are also underway at GRC.
SIGNIFICANCE
The data to be generated has been 
previously unavailable. Further, 
development of an in-house capability to 
investigate SWBLIs will allow CFD code 
developers and turbulence modelers to 
have direct input into the experiment. It will 
also allow the ability to revisit 
measurements if deemed necessary.
Future Validation Needs
• Most of the primary test cases are old datasets
– There have been significant advances measurement technology
– New scale‐resolving simulations require much more detailed boundary 
information
– We have gained a better understanding of the requirements for 
validation experiments
• We believe there is a need for new validation data sets
– What flows should be considered?
– What quantities should be measured?
– What are we missing?
29
Contact information
• Pleases contact us if you have any comments, or information 
on possible additional test cases
• Chris Rumsey, NASA Langley
– c.l.rumsey@nasa.gov
– 757‐864‐2165
• Jim DeBonis, NASA Glenn
– james.r.debonis@nasa.gov
– 216‐433‐6581
30
This Powerpoint presentation as well as additional write-ups on the 
recommended test cases will be posted to the 
http://turbmodels.larc.nasa.gov website


Test Cases for NASA's Revolutionary Computational Aerosciences Technical Challenge

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server