tbcc dual-inlet mode transition - shanti pages · – previous imx cfd analysis (2008) and...
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TBCC Dual-Inlet Mode Transition TBCC Inlet Analysis & Modeling
Kevin Bowcutt, Matt Sexton, Deric Babcock, Marty Bradley
The Boeing Company
Dave Saunders, John Slater
NASA Glenn Research Center
Jack Edwards, Santanu Ghosh
North Carolina State University
National Center for Hypersonic Combined Cycle Propulsion
June 16, 2011
Status Presentation
National Center for Hypersonic Combined Cycle Propulsion
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2
Goals and Approach
National Center for Hypersonic Combined Cycle Propulsion
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TBCC Dual-Inlet Mode Transition
Goals and Approach
Goals
Generation I – Assess current CFD technology
Generation II – Improve Gen I codes with Gen
II technology as required to
improve accuracy
– Retain Gen I efficiency for CFD
use in vehicle design and
development
Approach
Generation I – Bleed Modeling Methods
– Backpressure Modeling Methods
– Turbulence models (Year 1)
Generation II – Immersed Boundary (IB) bleed
modeling
– Hybrid RANS-LES* simulation
*RANS = Reynolds-Averaged Navier-Stokes
LES = Large Eddy Simulation
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4
Project Overview
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Gen I / Gen II TBCC Inlet Task Flow
IMX Testing
& CFD IMX
2007 2008 2009 2010 2011 2012 2013 2014
LIMX
IMX
LIMX
RALV Flowpath
Optimization
Gen I LIMX Analysis
Gen I IMX Analysis
Gen II
NC
HC
CP
Du
al
Inle
t A
ctivitie
s
LIMX Testing,
Pre-LIMX CFD, and Post-LIMX CFD
NA
SA
Hypers
on
ic
Co
mb
ine
d C
ycle
En
gin
e
LIMX CFD
IMX
CFD
RALV Dual Inlet
Flowpath
Optimization
Pre-LIMX CFD
Baseline for Gen II Post-LIMX CFD
Enhanced Meth.
IMX CFD Compare &
Enhanced Methods
IMX Bleed &
IB Enhance
IMX/LIMX RANS/LES
Enhancements
NC State Boeing NASA Legend
Bo
ein
g N
RA
Gen II
Devel.
Post-LIMX CFD
Enhanced Meth.
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Features Modeled in CFD Analysis
• Geometry Translation/Rotation
• Vortex Generators
• 9 Bleed Surfaces
• Multiple Bleed Patterns
IMX (1’x1’) and LIMX (10’x10’) Geometries
Low-speed ramp
Flow Direction
High-speed cowl
Low-speed cowl
High-speed flowpath
Low-speed flowpath
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IMX and LIMX Features of Interest
R1
R2 R3
R4
SW1
SW2
SW2
C1
C2
HS Throat
HS AIP
LS AIP
LS Throat
LS Throat Exit
Vortex Generators (3)
9 Bleed Surfaces (R=Ramp, C=Cowl, SW=Sidewall)
5 Comparison Planes
3 Vortex Generators
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LIMX Increases Available Data: IMX (9 Probe) and LIMX (40 primary Probe) Rakes
-1
-0.8
-0.6
-0.4
-0.2
0
0.2
0.4
0.6
0.8
1
-1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1
LIMX: Primary 40-tube
rake in black circles IMX: Primary 9-tube
rake in blue squares
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10
Bleed Modeling Approaches
0.75
0.78
0.81
0.84
0.87
0.90
0.93
0.84 0.86 0.88 0.90 0.92 0.94
Engine Flow Ratio
Tota
l P
ressure
Recovery
knee
Supercritical (terminal shock downstream
of throat bleed regions)
Stability
Standard Bleed Model
• Constant mass flux per area
• Does not depend on local flow
solution
Improved Bleed Model
• Based model by John W.
Slater of NASA GRC
• Based on plenum pressure
and local flow solution
• Allows for shocks across a
bleed surface
exitm
pplenum pt plenum
Aexit
M
M 0
shock
Qsonic
CD
pt Tt
pexit
Wbleed =Wholes = Wexit
Wholes
Wexit
Tt plenum
Aregion
Ableed
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Backpressure Approaches Generation I Methods: • Constant pressure outflow boundary condition
– Initially problematic, but developed approach that works and is now routinely used
• Converging-diverging nozzle – Time and labor prohibitive for an unstructured grid (no direct cell to cell mapping as in structured
grid)
• Mass injection at 90º angle to surface – As injected mass flow is increased, flow constricts, simulating a converging-diverging nozzle
– Observed large pockets of reverse flow extending from throat to mass injection
• Mass injection at 45º angle to surface (baseline Boeing approach) – This approach was initially the most stable
– Results compared favorably to other methods and to NASA CFD and wind-tunnel tests
Generation II: Model mass-flow plug using Immersed Boundary method
Mass is injected at a 45º
angle to the surface
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Comparison of Supercritical and
Backpressure Simulations x =
11
0 in
.
x =
12
0 in
.
x =
13
0 in
.
x =
14
0 in
.
x =
15
0 in
.
x =
16
0 in
.
x =
17
0 in
.
x =
18
0 in
.
x =
19
0 in
.
x =
20
0 in
.
x =
21
0 in
.
x =
22
0 in
.
x =
23
0 in
.
x =
24
0 in
.
x =
25
0 in
.
x =
26
0 in
.
x =
27
0 in
.
x =
28
0 in
.
x =
29
0 in
.
Supercritical Simulation Mach Number
Mach Number
Recovery
Recovery
Backpressured Simulation
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LIMX Boeing CFD
0.62
0.59 0.61
S-A (p0 = 35.216 psi)
LIMX NASA CFD
Comparison of Engine Face Flow Distortion:
Test vs. CFD Data
LIMX Test Data (TBD)
IMX Boeing CFD IMX NASA CFD IMX Test Data
RECave = 56.7%
DIST = 6.5% RECave = 55.4% (SA)
DIST = 16.7% (SA)
Rake
Measurements
Differences: • Unstructured grid • Bleed modeling (minor) • Grid density • Backpressure method
Differences: • Scale 6.6X • Bleed pattern • Geometry (minor)
Pattern
“flips”
Large scale validation data
needed
Differences: • Scale 6.6X • Bleed pattern • Geometry (minor)
Differences: • Structured grid • Bleed modeling (minor) • Grid density • Backpressure method
RECave = 60.1%* (SA)
DIST = 7.1% (SA)
Predicted flow distortion sensitive to modeling differences
Contours
Interpolated
RECave = 56.6% (SST)
DIST = 16.8% (SST)
* Includes adjustment for M=4
6.5 deg shock
RECave = 60.4% (SA)
DIST = 7.06% (SA)
* Includes adjustment for M=4
6.5 deg shock
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NASA LIMX Test Status
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LIMX Installed in the 10x10 SWT
17
Low-Speed
Cowl / Splitter
High-Speed
Cowl
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CCE LIMX Test Status
Low-Speed Inlet Mass-Flow Ratio, W2LS / WHS
0.00
0.10
0.20
0.30
0.40
0.50
0.60
0.00 0.10 0.20 0.30 0.40 0.50 0.60
0.70
Lower operability
possible inlet
unstart/buzz
Higher
distortion
Simulated
Inlet Mode
Transitioning
ScheduleT
ota
l P
ress
ure
Rec
ov
ery,
pt2
/ p
t0
• Phase I testing commenced 7 March 2011
• 10’X10’ SWT started and operated
nominally with large LIMX model
• Over one thousand data points obtained
to date - being analyzed
• Series of “cane curves” generated that
outline operating limits and quantify
performance
• Flow distortion data at the aerodynamic
interface plane (AIP) obtained
• Data being compared to performance
goals:
• Recovery: 64% (minimum of 57%)
• Distortion: less than 10%
• Phase I testing resumed in June after repairs on
ramp hydraulic leak and edge seals
• Phase 1 testing should be complete in
August-September
• Data will be made available after NASA
quality review
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Boeing CFD Analysis Status
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Boeing Activities
• Work to Date – Previous IMX CFD analysis (2008) and comparisons to NASA
CFD and test data
– LIMX pre-test CFD analysis (2008-2011) to investigate backpressure approaches, turbulence models, and numerical methods
– Worked with NC State to define geometry, assumptions, and approach for Gen II IMX analysis
• Working to identify flow field discrepancies and ways to improve solution
• Next Steps – Compare IMX CFD solutions using standard bleed model with
those from NC State using IB bleed model
– Begin running LIMX (ITAR configuration) CFD solutions using wind-tunnel conditions and compare results with test data
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LIMX CFD Analysis Effort
• Different bleed and backpressure methods investigated
• Compared LIMX and IMX solutions
• Performed detailed investigation of flow features such as
separation and recirculation
• Identified correlation between terminal shock location
and flow features that is largely independent of
backpressure and bleed methods employed
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LIMX Low-Speed (LS) Duct Features
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Highlighted LIMX LS Duct Features
Cowl (Green)
Ramp (Red)
Sidewall (Grey)
FLOW
Cowl (Green)
Sidewall (Grey)
Ramp (Red)
Ramp (Red)
Cowl (Green)
Sidewall (Grey)
Vortex Generators
(Black Circles)
Throat Region
(Blue Squares)
Looking Downstream Into Inlet Looking Outward From Centerline
Between Throat and Vortex Generators
Looking Upstream From Vortex
Generators to Throat Region
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LIMX LS Duct – Flow Features of Interest
Blue: Pt/Pt0 = 0.4 Shows boundary layers and regions
of high total pressure loss
Green: Mach = 1 Shows throat and sonic locations
Pink: Velocity = 0 Shows flow separation regions
Location of
Separation Region
Location of
Terminal Shock
Location of
Boundary Layer/
Vortical Flow Region
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LIMX LS Duct – Terminal Shock Location
As a Function of Backpressure
Constant BP 21.7 Constant BP 22.3
Blowing BP 22 Constant BP 22.0
Constant BP 22.2
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Modeling Methods Investigated
• Backpressure methods: – Blowing = Blowing at 45deg pointed downstream
• Pressures quoted are psi total of injection plenum
– Const BP = Imposing a constant backpressure BC at exit plane of low speed duct and increasing level as solution proceeds
• Pressures quoted are psi static at outflow plane
• Grids: – Original = Standard-quality grid of ordinary density typical of automated
unstructured grid process
– Refined = High-quality grid of higher density generated with manual effort
• Bleed models: – Specified mass flux through bleed zone (“fixed bleed” model)
– Bleed mass flux a function of specified plenum pressure and computed surface static pressure (Slater bleed model)
Identified correlation between terminal shock location and flow features that
is largely independent of backpressure and bleed methods employed
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LIMX LS Duct – Constant BP, Original Grid
Blue: Pt/Pt0 = 0.4 Shows boundary layers and regions
of total pressure loss
Green: Mach = 1 Shows throat and sonic locations
Pink: Velocity = 0 Shows flow separation regions
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LIMX LS Duct – Blowing, Original Grid
Blowing BP Method is at a
slightly higher backpressure
Blue: Pt/Pt0 = 0.4 Shows boundary layers and regions
of total pressure loss
Green: Mach = 1 Shows throat and sonic locations
Pink: Velocity = 0 Shows flow separation regions
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LIMX LS Duct – Blowing, Refined Grid
Shock is aft for refined grid More
blowing required to achieve
equivalent backpressure
Blue: Pt/Pt0 = 0.4 Shows boundary layers and regions
of total pressure loss
Green: Mach = 1 Shows throat and sonic locations
Pink: Velocity = 0 Shows flow separation regions
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LIMX LS – Recovery Comparisons at AIP For
Different Back Pressure Methods and Grids
Constant BP
Original Grid
Recovery: 0.628
Distortion: 0.098
Blowing BP
Original Grid
Recovery: 0.624
Distortion: 0.097
Blowing BP
Refined Grid
Recovery: 0.624
Distortion: 0.092
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LIMX LS Duct – Fixed Bleed, Original Grid
Blue: Pt/Pt0 = 0.4 Shows boundary layers and regions
of total pressure loss
Green: Mach = 1 Shows throat and sonic locations
Pink: Velocity = 0 Shows flow separation regions
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LIMX LS Duct – Slater Bleed, Original Grid
Shock forward compared
to fixed bleed case
Blue: Pt/Pt0 = 0.4 Shows boundary layers and regions
of total pressure loss
Green: Mach = 1 Shows throat and sonic locations
Pink: Velocity = 0 Shows flow separation regions
National Center for Hypersonic Combined Cycle Propulsion
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LIMX LS Duct – Slater Bleed, Refined Grid
Refined grid again requires
more blowing to achieve
equivalent backpressure
Blue: Pt/Pt0 = 0.4 Shows boundary layers and regions
of total pressure loss
Green: Mach = 1 Shows throat and sonic locations
Pink: Velocity = 0 Shows flow separation regions
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LIMX LS – Recovery Comparisons at AIP
For Different Bleed Models and Grids
Fixed Bleed
Original Grid
Recovery: 0.624
Distortion: 0.097
Slater Bleed
Original Grid
Recovery: 0.621
Distortion: 0.106
Slater Bleed
Refined Grid
Recovery: 0.624
Distortion: 0.082
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IMX CFD Analysis Observations
• IMX analysis exhibits greater sensitivity to backpressure – Nonlinear variation of terminal shock location with backpressure
• Separation/recirculation pattern flips from ramp (LIMX) to cowl (IMX)
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IMX LS Duct – Terminal Shock Location
As a Function of Backpressure
Constant BP 85, Fixed Bleed Blowing BP 88, Slater and Fixed Bleed
Constant BP 90, Fixed Bleed Constant BP 90, Slater Bleed
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IMX LS Duct – Constant BP 85, Fixed Bleed
Terminal shock aft
Blue: Pt/Pt0 = 0.4 Shows boundary layers and regions
of total pressure loss
Green: Mach = 1 Shows throat and sonic locations
Pink: Velocity = 0 Shows flow separation regions
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IMX LS Duct – Constant BP 90, Fixed Bleed
Terminal shock moves forward
Blue: Pt/Pt0 = 0.4 Shows boundary layers and regions
of total pressure loss
Green: Mach = 1 Shows throat and sonic locations
Pink: Velocity = 0 Shows flow separation regions
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IMX LS – Recovery Comparisons at AIP
For Different Back Pressures
Constant BP 90 psi
Recovery: 0.574
Distortion: 0.102
Constant BP 85 psi
Recovery: 0.539
Distortion: 0.122
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IMX LS Duct – Blowing BP 88, Fixed Bleed
Terminal shock location closely matches
Slater Bleed case on following chart
Blue: Pt/Pt0 = 0.4 Shows boundary layers and regions
of total pressure loss
Green: Mach = 1 Shows throat and sonic locations
Pink: Velocity = 0 Shows flow separation regions
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IMX LS Duct – Blowing BP 88, Slater Bleed
Blue: Pt/Pt0 = 0.4 Shows boundary layers and regions
of total pressure loss
Green: Mach = 1 Shows throat and sonic locations
Pink: Velocity = 0 Shows flow separation regions
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Blowing 88 psi
Fixed Bleed
Recovery: 0.554
Distortion: 0.167
Blowing 88 psi
Slater Bleed
Recovery: 0.552
Distortion: 0.165
IMX LS – Recovery Comparisons at AIP
For Different Bleed Models
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LIMX LS Duct – Time-Accurate BCFD Solution
Delayed-DES (DDES) starting from
Unsteady RANS (URANS) solution
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45
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NC State Immersed Boundary
Bleed Modeling Status
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Immersed Boundary (IB) Methodology
• Immersed objects rendered as stereo-lithography (STL) files
• Flow domain divided into three categories of cells based on signed
distance function Φ
– Field Cells
– Band Cells
– Interior Cells
Field points Band points Interior
Immersed body surface
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plenums for ramp
and cowl bleed
regions
inflow
turbine flow path
scramjet
flow path
sidewall
plenums for R4
and C2
Computational domain for the IMX inlet Bleed regions for the IMX inlet
Overview of IB Generated Grid
• Immersed Boundary (IB) approach distinguishes between flow and solid regions
from overlay of CAD model on mesh
• Cells ~ 35 M spread over 128 or 144 processors
• Bleed regions R1, R2, R3, R4, C1 and C2 connected to modeled plenums; SW1,
SW2 and SW3 discharge to infinite plenum
• Patched-mesh methodology provides ~ 60 cells / bleed hole
(10 streamwise x 6 spanwise)
inflow
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Modeling Issues Being Worked
• CAD file/STL file surface roughness (fixed)
• Error in communication between zones having patched interfaces found (fixed)
• Blunt sidewall leading edge (fixed)
• Bleed rates achieved lower than NASA/Boeing CFD and IMX wind-tunnel test (in-work)
• Grid resolution and/or grid discrepancies (in-work)
• Bleed zone interactions (in-work)
• Transient flow start-up (0 velocity, 0 bleed) simulation not working, switching to steady state and over-bled (in-work)
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• Regenerated STL files to be flush with the surface
• Local roughness effects removed •
49
49
CAD File Surface Roughness
STL rendition of C1 bleed;
(NEW), (OLD)
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Bleed Mass Flow Differences
iteration
ma
ss
flo
w(k
g/s
)
0 5000 100000.0E+00
5.0E-03
1.0E-02
1.5E-02
2.0E-02
2.5E-02
R1-in
R1-out
R1-Boeing
iteration
ma
ss
flo
w(k
g/s
)
0 5000 100000.0E+00
5.0E-03
1.0E-02
1.5E-02
2.0E-02
2.5E-02
R2-in
R2-out
R2-Boeing
iteration
ma
ss
flo
w(k
g/s
)
0 5000 100000.0E+00
5.0E-03
1.0E-02
1.5E-02
2.0E-02
2.5E-02
R3-in
R3-out
R3-Boeing
iteration
ma
ss
flo
w(k
g/s
)0 5000 10000
0.0E+00
5.0E-03
1.0E-02
1.5E-02
2.0E-02
2.5E-02
R4-in
R4-out
R4-Boeing
iteration
ma
ss
flo
w(k
g/s
)
0 5000 100000.0E+00
5.0E-03
1.0E-02
1.5E-02
2.0E-02
2.5E-02
C1-in
C1-out
C1-Boeing
iteration
ma
ss
flo
w(k
g/s
)
0 5000 100000.0E+00
5.0E-03
1.0E-02
1.5E-02
2.0E-02
2.5E-02
C2-in
C2-out
C2-Boeing
iteration
ma
ss
flo
w(k
g/s
)
0 5000 100000.0E+00
5.0E-03
1.0E-02
1.5E-02
2.0E-02
2.5E-02
SW-1
SW-1 Boeing
iteration
ma
ss
flo
w(k
g/s
)
0 5000 100000.0E+00
5.0E-03
1.0E-02
1.5E-02
2.0E-02
2.5E-02
SW-2
SW-2 Boeing
iteration
ma
ss
flo
w(k
g/s
)
0 5000 100000.0E+00
5.0E-03
1.0E-02
1.5E-02
2.0E-02
2.5E-02
SW-3
SW-2 Boeing
Ramp bleed (red)
Cowl bleed (blue)
Sidewall bleed (green)
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Spurious Flow Feature Observed In
Supercritical Solution
Recirculation
Wall pressure along centerline:
line – NCSU, squares – Slater
(AIAA 2009-2349)
• Strong recirculation observed in
region between bleed stations
R3/C1 and R4/C2
• Leads eventually to formation of
a normal shock at this location
• Supercritical simulation should
have no terminal shock
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Different Sidewall Bleed Methods Investigated
Mach Contours
recirculation
Sidewall with IB resolved bleed Sidewall with bleed BC
Response essentially identical for both bleed models –
growth of separation region near throat led to eventual unstart
recirculation
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NC State Gen II Next Steps
– Plan:
• Apply Boeing fixed bleed model to all bleed surfaces to check
if a converged solution can be achieved
• Replace fixed bleed model with IB resolved bleed one bleed
zone at a time
• Backpressure IB solution to wind-tunnel level
– Potential sources of problem
• Interaction between multiple IB bleed locations at a given
streamwise location
• Detailed modeling of localized bleed zones – in place of
corner to corner bleed – may not yield desired improvement
in boundary layer health in RANS simulations
• Unphysical start-up transients due to uniform supersonic flow
and plenum pressure initial conditions
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Summary
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Importance and Impact of Effort
• Accurate analysis of complex supersonic inlets essential for design and optimization of TBCC flowpaths and for developing turbine-to-scramjet transition strategies
• NASA experimental data and NASA, Boeing, and NCSU CFD work being leveraged to improve TBCC inlet analysis methods
• Gen I CFD methods will be validated and/or inaccuracies identified
• Gen II methods will be developed and applied to improve solution accuracy and to better understand sources of Gen I inaccuracy
• Gen II methods will be incorporated in Gen I codes and/or knowledge gained in their use will help improve modeling employed in Gen I codes
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Collaboration with NASA and
Air Force Core Research
• Boeing* and NASA pre-test and post-test analysis of NASA Glenn
IMX and LIMX experiments
• Boeing* and NASA IMX and LIMX CFD analysis results shared at
JANNAF distortion workshops organized by NASA and AFRL
• NCSU work presented at the 4th Annual Shock Wave/Boundary
Layer Interaction (SWBLI) Flow Control and Modeling Workshop
held at NASA Glenn Research Center, 5-6 April 2011
• Other past and future presentations:
– TBCC Inlet Invited Session* at ASM11
– NCHCCP Invited Session at ASM12
* Supported by Center, NASA NRA, and Boeing funding
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Student / Post-Doc Support
• North Carolina State University – Santanu Ghosh
• Ph.D., 2010, North Carolina State University
• Boeing - University of Southern California – Matt Sexton
• Ph.D. Dissertation Topic: “Design of ablation heat shields for
planetary entry with uncertainty quantification”
• Ph.D. coursework supported by Boeing
• Dissertation research supported by research grant
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Questions ?