TBCC Dual-Inlet Mode Transition - SHANTI Pages · – Previous IMX CFD analysis (2008) and...

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


2

2

Goals and Approach


3

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


4

4

Project Overview


5

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.


6

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


7

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


8

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


10

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


11

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


12

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


13

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


14

14

NASA LIMX Test Status


17

LIMX Installed in the 10x10 SWT

17

Low-Speed

Cowl / Splitter

High-Speed

Cowl


18

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


20

20

Boeing CFD Analysis Status


21

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


22

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


23

LIMX Low-Speed (LS) Duct Features


24

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


25

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


26

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


27

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


28

LIMX LS Duct – Constant BP, Original Grid


of total pressure loss




29

LIMX LS Duct – Blowing, Original Grid

Blowing BP Method is at a

slightly higher backpressure






30

LIMX LS Duct – Blowing, Refined Grid

Shock is aft for refined grid More

blowing required to achieve

equivalent backpressure






31

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


32

LIMX LS Duct – Fixed Bleed, Original Grid






33

LIMX LS Duct – Slater Bleed, Original Grid

Shock forward compared

to fixed bleed case






34

LIMX LS Duct – Slater Bleed, Refined Grid

Refined grid again requires

more blowing to achieve

equivalent backpressure






35

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


36

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)


37

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


38

IMX LS Duct – Constant BP 85, Fixed Bleed

Terminal shock aft






39

IMX LS Duct – Constant BP 90, Fixed Bleed

Terminal shock moves forward






40

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


41

IMX LS Duct – Blowing BP 88, Fixed Bleed

Terminal shock location closely matches

Slater Bleed case on following chart






42

IMX LS Duct – Blowing BP 88, Slater Bleed






43

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


44

LIMX LS Duct – Time-Accurate BCFD Solution

Delayed-DES (DDES) starting from

Unsteady RANS (URANS) solution


45

45

45

NC State Immersed Boundary

Bleed Modeling Status


46

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


47

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


48

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)


49

• 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)


50

50

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)


51

51

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


52

52

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


53

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


54

Summary


55

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


56

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


57

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


59

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