(NASA-KP-132£) NAS T_C HNICAL

SUMMARIES- NUMERICAL AERODYNAMIC

SIMULATION pROGRAM. MARCH 199Z -

FEBRUARY 1993 (NASA) Ibl p

N94-27764

UncI as

H1/99 0000339
The use of color graphics is a powerful

and flexible visualization tool for the

science of computational fluid dynamics.

Without color graphics, the millions of

numbers which form the supercomputer

solutions to the studies contained in this

report and on the cover would be

virtually incomprehensible.

On the cover: Nateri K. Madavan is a

researcher with the MCAT Institute at NASA

Ames Research Center. The cover image is

from Madavan's summary, "Unsteady Multistage

Turbomachinery Applications," on page 77.

h
NAS

Technical

Summaries

Numerical Aerodynamic Simulation Program

National Aeronautics and Space Administration

Ames Research Center
ii
Preface

NASA created the Numerical Aerodynamic Simulation (NAS)

Program in 1987 to focus resources on solving critical problems

in aeroscience and related disciplines by utilizing the power of

the most advanced supercomputers available. The NAS Program

provides scientists with the necessary computing power to solve

today's most demanding computational fluid dynamics problems

and serves as a pathfinder in integrating leading-edge supercomputing technologies, thus benefiting other supercomputer

centers in government and industry.

During the 1992-93 Operational Year, the scientific community

was given access to a Cray-2 and a Cray Y-MP. The Cray-2,

a first generation supercomputer, had four processors,

256 megawords of central memory, and a total sustained

speed of 250 million floating point operations per second.

The Cray Y-MP, a second generation supercomputer, had eight

processors, 256 megawords of solid state device memory, and a

total sustained speed of 1 billion floating point operations per

second. Because of its higher performance, the Cray Y-MP

delivered 67 percent of the total number of supercomputer

hours used by the NAS scientific community during the

1992 93 Operational Year.

The 1992-93 Operational Year concluded with 399 high-speed

processor (HSP) projects and 91 parallel (PAR) projects representing NASA, the Department of Defense, other Government

agencies, private industry, and universities. This document--the

NASTechnicaISummaries provides a glimpse at some of the

significant scientific results for the year.

The content of the NAS Technical Summaries for the 1992-93

Operational Year has been expanded from previous years.

Traditionally, this report contains selected scientific results

computed using HSP resources. This is the first edition to also

contain results computed using PAR resources. An icon above

the text of each technical summary indicates the resource

used-- _C_ represents HSP, and _._ represents PAR.

The organization of the summaries also differs from previous

years. Traditionally this report was organized alphabetically by

principal investigator. This edition is organized by research

disciplines and alphabetically within each discipline to make the

book more functional.

Additional copies of the NAS Technical Summaries for this year

and for previous years are available on request. Contact the

NAS Documentation Center at NASA Ames Research Center,

MS 258-6, Moffett Field, CA 94035-1000, or via email addressed

to doc-center@nas.nasa.gov, or telephone (415) 604-4632.

Number o! Cray Projects by Category

100

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Number of Parallel Projects by Category

35 33

30

25

20

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Number of Cray Projects by Discipline

B Administrative I Hypersonics

[] Aerodynamics • Propulsion

_] Rolorcraft

B Turbulence

[] Other

Astrophysics

Atmospheric Sciences

Chemistry

Computer Science

Electrical Engineering

Electromagnetlcs

Geology

Hydrodynamics

Hydrogeology

Life Sciences

Material Sciences

Mathematics

Physics

Structural Mechanics

Number of Parallel Projects by Discipline

• Administrative i Hypersonics [] Rotorcra.

[] Aerodynamics • Propulsion • Turbulence

[] Other

Astrophysics

Atmospheric Sciences

Chemistry

Computer Science

Electromagnetics

Fluid Mechanics

Hydrogeology

Life Sciences

Material Sciences

Other Fluids

Physics

Structural Mechanics

p_s_'--1OtN¢= PAGE Ett.ANK

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iii

Table of Contents

Principal

Investigator NAS Summary Page

Aeronautics" Basic Research

Ramesh K. Agarwal

Harold L. Atkins

Timothy J. Barth

Alan B. Cain

Richard L. Campbell

Frank Caradonna

Neal M. Chaderiian

Richard D. Crouse

Daniel F. Dominik

Thomas A. Edwards

Datta Gaitonde

Philip B. Gingrich

David Halt

Carolyn R. Kaplan

Steve L. Karman, Jr.

Numerical Solution of Three-Dimensional Maxwell Equations ............................................................ 3

Co-investigators: Mark Shu and Mark R. Axe

McDonnell Douglas Corporation

Simulation of Acoustic Scatter ............................................................................................................. 4

NASA Langley Research Center

A Three-Dimensional Implicit Unstructured Euler Solver .................................................................... 5

Co-investigator: Samuel W. Linton

NASA Ames Research Center/Sterling Software Systems

Simulation of Supersonic Jet Screech .................................................................................................. 6

Co-investigators: William W. Bower and William W. Romer

McDonnell Douglas Aerospace

Efficient Constrained Aerodynamic Design ......................................................................................... 7

Co-investigators: Leigh A. Smith and Raymond E. Mineck

NASA Langley Research Center

Computation of Helicopter Rotor and Wake Flows ............................................................................. 8

Co-investigators: K. Ramachandran, I. Bridgeman, and D. Poling

U.S. Army Aeroflightdynamics Directorate, AVSCOM/Flow Analysis, Inc./

Woodside Summit Group, Inc./The Boeing Company

High-Incidence Static/Dynamic Roll Computations ............................................................................. 9

Co-investigators: John Ekaterinaris, Yuval Levy, and Lewis B. Schiff

NASA Ames Research Center/Naval Post Graduate School/Stanford University

Cavity Aeroacoustic Loads Prediction ............................................................................................... 10

Northrop Corporation

High-Fidelity Space Shuttle Simulation ............................................................................................. 11

Co-investigators: K. Rajagopal, S. Vuong, J. Wisneski, C. Oiling, G. Hock, and I. Sikora

Rockwell International, Space Systems Division

Aerodynamic Optimization of Supersonic Aircraft ............................................................................ 12

Co-investigators: Samson H. Cheung and I. C. Chang

NASA Ames Research Center/MCAT Insitute

Three-Dimensional Turbulent Double-Fin Interactions ...................................................................... 13

Co-investigator: Joseph Shang

WL/FIMM, Wright Patterson AFB

Design-By-Optimization Method ...................................................................................................... 14

Rockwell International, North American Aircraft DMsion

F/A- 18 Store Carriage Analyses ......................................................................................................... 15

Co-investigators: Mori Mani and Todd Michal

McDonnell Aircraft Company

Lifted Ethylene Jet Diffusion Flames .................................................................................................. 16

Co-investigator: Elaine S. Oran

Naval Research Laboratory

Unstructured Grid/Flow Solver Calibration ....................................................................................... 1 7

Co-investigators: Doug Howlett and Elizabeth A. Robertson

Lockheed Engineering and Sciences Company

p_"_(W)_ PA':_E BL&NK NOT F!,L_D

PAGE_ INTENTIONALLY,3LANK
Principal

Investigator NAS Summary Page

Stephen P. Klotz

Rainald Lohner

Fred W. Martin, Jr.

Parviz Moin

Yehia Rizk

David M. Schuster

Michael J. Siclari

Eli Turkel

Veer N. Vatsa

Simulation of Flow About an Airborne Observatory. ......................................................................... 18

Co-investigator: Christopher A. Atwood

NASA Ames Research Center

Finite-Element Euler Solver for a Distributed Memory Computer ....................................................... 19

Co-investigator: Alexander Shostko

George Washington University

Space Shuttle Flow Field ................................................................................................................... 20

Co-investigators: Pieter Buning, William Chan, Ing-Tsau Chiu, Ray Gomez, Jeff Slotnick,

Edward Ma, Dan Pearce, Thomas Wey, and Max Kandula

NASA Johnson Space Center

Direct Computation of Aerodynamic Sound Generation ................................................................... 21

Co-investigators: Sanjiva K. Lele, Tim Colonius, Brian E. Mitchell, and Jonathan B. Freund

NASA Ames Research Center/Stanford University

Computational Fluid Dynamics of F- 18 Flow Field ........................................................................... 22

Co-investigators: Scott Murman and Ken Gee

NASA Ames Research Center

Wings with Control Surfaces ............................................................................................................. 23

Lockheed Engineering and Sciences Company

Sonic Boom Predictions for High-Speed Civil Transport .................................................................... 24

Co-investigator: Kamran Fouladi

Grumman Corporate Research Center/Lockheed Engineering and Sciences Company

Navier-Stokes Equations for Jet Acoustics ......................................................................................... 25

Co-investigator: Ehtesham Hayder

ICOMP/NASA Lewis Research Center

Block-Structured Approach for High-Lift Applications ....................................................................... 26

Co-investigator: Mark D. Sanetrik

NASA Langley Research Center/Analytical Services and Materials, Inc.

Aeronautics: Hypersonics, Subsonics, Supersonics, Transonics

Hypersonics

Ramesh K. Agarwal

Jorge E. Bardina

Charles E. Cockrell, Jr.

Peter A. Gnoffo

Brian L. Haas

Flow Fields Including Chemical Kinetics ........................................................................................... 29

Co-investigators: Jerry E. Deese, Paul Schulte, T. Mark Walter, Thomas P. Gielda,

and Hugh Thornburg

McDonnell Douglas Research Laboratories

Three-Dimensional Compressible Turbulence Modeling ................................................................... 30

Co-investigator: Thomas J. Coakley

MCAT Institute/NASA Ames Research Center

Waverider and Reference Model Computational Study ..................................................................... 31

NASA Langley Research Center

Reentry Heating Effects on the Shuttle Orbiter ................................................................................... 32

NASA Langley Research Center

Particle Simulation of Hypersonic Rarefied Flow ............................................................................... 33

Co-investigators: Michael A. Fallavollita and William J. Feiereisen

Eloret Institute/Stanford University/NASA Ames Research Center

vi
Principal

Investigator

William D. Henline

Lawrence D. Huebner

Ajay Kumar

Scott L. Lawrence

C. C. Lee

K. Kurian Mani

Charles R. McClinton

Grant Palmer

Ramadas K. Prabhu

C. D. Pruett

R. P. Roger

R. Clayton Rogers

Rajiv Thareja

Richard A. Thompson

K. James Weilmuenster

NAS Summary Page

Computation of Space Transport Vehicle Re-Entry Heating ............................................................... 34

Co-investigators: Y.-K. Chen, G. E. Palmer, and S. M. White

NASA Ames Research Center/Eloret Institute

Powered Hypersonic Air-Breathing Configuration Studies ................................................................. 35

-Co-investigator: Kenneth E. Tatum

NASA Langley Research Center/Lockheed Engineering and Sciences Company

Computational Performance Enhancement of a Scram jet Inlet .......................................................... 36

Co-investigator: Dal J. Singh

NASA Langley Research Center

Integrated Hypersonic Vehicle Simulation ........................................................................................ 37

Co-investigators: Gregory A. Molvik, Johnny R. Narayan, and Bradford C. Bennett

NASA Ames Research Center/MCAT Institute

Efficiency Improvements of Boundary Layer Transition Predictions ................................................... 38

Co-investigators: Shawn Hagmeier, Brad Hopping, and Charles Vaporean

McDonnell Douglas Corporation

Flow Past a Dynamic Multiple Body System ..................................................................................... 39

Co-investigator: J. W. Haney

Rockwell International, Space Systems Division

Performance Potential of Full-Scale Injectors .................................................................................... 40

Co-investigator: David W. Riggins

NASA Langley Research Center

Hypersonic Computational Studies ................................................................................................... 41

Co-investigators: Surendra Sharma, Chul Park, Stephen M. Ruffin, Tahir Gokcen,

Dikran Babikian, Susan Tokarcik-Polsky, and Ethiraj Venkatapathy

NASA Ames Research Center/Eloret Institute

Computational Studies of Nonequilibrium Flows .............................................................................. 42

Co-investigator: George C. Olsen

Lockheed Engineering and Sciences Company/NASA Langley Research Center

Validation of the Parabolized Stability Equation Method ................................................................... 43

Co-investigator: C-L. Chang

Analytical Services and Materials, Inc./High Technology Corporation

Jet Interaction Aero-Optic Effects on Hypersonic Interceptors ........................................................... 44

Co-investigator: S. C. Chan

Teledyne Brown Engineering

Fuel Plume Imaging/CFD Comparisons ............................................................................................. 45

Co-investigators: Robert D. Bittner and David W. Riggins

NASA Langley Research Center

Adaptive Unstructured Hypersonic Multigrid Solver ......................................................................... 46

Co-investigators: Kenneth Morgan, Jaime Peraire, and Joaquim Peiro

Lockheed Engineering and Sciences Company/University of Wales/

Massachusetts Institute of Technology/Imperial College

Aerothermodynamic Benchmark for Candidate AMLS Vehicle .......................................................... 47

NASA Langley Research Center

Na vier-Stokes Simulations of Orbiter Aerodynamic Characteristics .................................................. 48

Co-investigators: William L. Kleb and Francis A. Greene

NASA Langley Research Center

vii
Principal

Investigator NAS Summary Page

Thomas A. Zang Modeling for High-Speed Transitional Boundary Layers .................................................................... 49

Co-investigator: Nabil M. Et-Hady

NASA Langley Research Center/Analytical Services and Materials, Inc.

S. H. Konrad Zhu DSMC Simulation of High-Altitude Plume Interaction ....................................................................... 50

Co-investigator: Leonardo Dagum

Rockwell International, Rocketdyne Division/NASA Ames Research Center

Subsonics

Kenneth M. Jones Subsonic High-L ift Analysis ............................................................................................................... 51

Co-investigators: Kevin Kjerstad and Victor Lessard

NASA Langley Research Center/ViGYAN, Inc.

Anutosh Moitra Analysis and Design of High-Lift Systems .......................................................................................... 52

Co-investigator: Wendy B. Lessard

High Technology Corporation/NASA Langley Research Center

Lewis B. Schiff Aircraft Forebody Flow Control Technology ...................................................................................... 53

Co-investigators: Ken Gee, Scott Murman, Gabriel Font, and Roxana Agosta

NASA Ames Research Center

David T. Yeh Steady and Unsteady Flows at High Angles of Attack ........................................................................ 54

Rockwell International, North American Aircraft Division

Supersonics

Dharmanshu L. Antani CFD Applications to HSCT Design and Analysis ............................................................................... 55

Co-investigators: Shreekant Agrawal, David L. Rodriguez, Ramesh K. Agarwal,

and Raymond Hicks

McDonnell Aircraft Company/NASA Ames Research Center

Transonics

Osama A. Kandil Transonic Flow Around a Delta Wing ............................................................................................... 56

Old Dominion University

Aeronautics: Propulsion

John J. Adamczyk Simulation of Turbomachinery Flows ................................................................................................ 59

Co-investigators: Tim Beach, Mark Celestin, Kevin R. Kirtley, Rick Mulac, Wai-Ming To,

Jeff Yokota, and Aamir Shabbir

NASA Lewis Research Center/Sverdrup Technology, Inc./ICOMP

Richard A. Blech Investigation of Scalability of Parallel Turbomachinery Codes ........................................................... 60

Co-investigators: Edward J. Milner, Scott E. Townsend, and Angela Quealy

NASA Lewis Research Center

Pieter G. Buning Subsonic Transport Propulsion/Airframe Integration .......................................................................... 61

Co-investigators: Lie-Mine Gea, Cathy Maksymiuk, and W. R. Van Dalsem

NASA Ames Research Center/McDonnell Douglas Corporation

Jean-Luc Cambier Pulsed Detonation Wave Augmentor Concept .................................................................................. 62

Co-investigators: Henry G. Adelman and Gene P. Menees

EIoret Institute/NASA Ames Research Center

viii
Principal

Investigator

Jeffrey A. Catt

Richard D. Cedar

C. L. Chen

H. C. Chen

Wei J. Chyu

Sanford M. Dash

Frederik I. de long

Richard L. Gaffney, Ir.

Karen L. Gundy-Burlet

Chunill Hah

Doug G. Howlett

J. Mark Janus

Jinho Lee

E. D. Lynch

Nateri K. Madavan

NAS Summary Page

Evaluating Nozzle Drag Using Computational Fluid Dynamics ......................................................... 63

Co-investigators: Tracy J. Welterlen and Brett W. Denner

Lockheed Engineering and Sciences Company

CFD for Engine-Airframe Integration ................................................................................................. 64

Co-investigator: Craig M. Kuhne

General Electric Aircraft Engines

Computation of Separated Nozzle Flows .......................................................................................... 65

Co-investigators: S. R. Chakravarthy and C. M. Hung

Rockwell International Science Center/NASA Ames Research Center

Inverse Design of Installed Nacelle ................................................................................................... 66

Co-investigators: T. Y. Su, T. I. Kao, and D. A. Naik

Boeing Commercial Airplane Group/NASA Langley Research Center/ViGYAN, Inc.

Airframe/Inlet Aerodynamics ............................................................................................................. 67

Co-investigators: Tom I-P. Shih and David A. Caughey

NASA Ames Research Center/Carnegie Mellon University/Cornell University

Three-Dimensional High-Speed Plume-Propulsive Flow Fields ......................................................... 68

Co-investigators: Neeraj Sinha, Brian I. York, Robert A. Lee, Ashvin Hosangadi,

and Donald C. Kenzakowski

Science Applications International Corporation

Turbine-Blade Tip Clearance Flows .................................................................................................. 69

Co-investigator: Tony Chan

Scientific Research Associates, Inc.

Interaction of Turbulence and Chemical Reaction ............................................................................. 70

Co-investigators: leffery A. White, Sharath Girimaji, and J. Phillip Drummond

Analytical Services and Materials, Inc./Institute for Computer Applications

in Science and Engineering/NASA Langley Research Center

Unsteady Turbomachinery Computations ......................................................................................... 71

Co-investigator: Akil Rangwalla

NASA Ames Research Center/MCAT Institute

Three-Dimensional Transonic Compressor Stage Flows .................................................................... 72

Co-investigators: J. Loellbach, S. L. Puterbaugh, and W. W. Copenhaver

NASA Lewis Research Center/ICOMP/WL/FIMM, Wright Patterson AFB ?

Evaluating High-Angle-of-Attack Inlet Characteristics ........................................................................ 73

Co-investigators: Brett W. Denner and Chris L. Reed

Lockheed Engineering and Sciences Company

Computational Analysis of a Tip-Engine Configuration ...................................................................... 74

Co-investigators: Animesh Chatterjee and Chris Cave

Mississippi State University

Studies of Mixing and Combustion in Supersonic Flows .................................................................... 75

Sverdrup Technology, Inc.

Turbulent Base Heating Computational Fluid D/namics ................................................................... 76

Co-investigators: R. Lagnado and Y. Hsu

Rockwell International, Rocketdyne Division

Unsteady Multistage Turbomach inery Applications .......................................................................... 77

Co-investigator: San jay Mathur

MCAT Institute/NASA Ames Research Center/Iowa State University
Principal

Investigator

Charles R. McClinton

Charles R. McClinton

Suresh Menon

Suresh Menon

John C. Otto

Thanh T. Phan

Tom I-P. Shih

Clifford E. Smith

C. S. Tan

Jong H. Wang

Kurt F. Weber

Kurt F. Weber

Perry A. Wooden

David T. Yeh

Shaye Yungster

NAS Summary Page

Base Pressurization Methods for Scram jet Combustors ...................................................................... 78

Co-investigator: Paul Vitt

NASA Langley Research Center/Analytical Services and Materials, Inc.

Simulation of Scram jet Combustor Flow Field ................................................................................... 79

Co-investigator: S. Srinivasan

NASA Langley Research Center

Parallel Simulation of Unsteady Combustion .................................................................................... 80

Co-investigator: Sisira Weeratunga

Georgia Institute of Technology/NASA Ames Research Center

Unsteady Combustion in a Ramjet .................................................................................................... 81

Co-investigator: Thomas M. Smith

Georgia Institute of Technology

Three-Dimensional Parallel Mixing Computations ............................................................................ 82

NASA Langley Research Center

Launch Vehicle Base Flow Simulations ............................................................................................. 83

Co-investigator: Timothy J. Ventimiglia

General Dynamics, Space Systems Division

Flow in Turbine-Blade Coolant Passages ........................................................................................... 84

Co-investigators: Mark A. Stephens and Kestutis C. Civinskas

Carnegie Mellon University/NASA Lewis Research Center

Jets-ln-Crossflow Mixing ................................................................................................................... 85

Co-investigators: Daniel B. Bain and James D. Holdeman

CFD Research Corporation/NASA Lewis Research Center

Flow Phenomena in Turbomachinery ............................................................................................... 86

Co-investigators: S. Arif Khalid and Ted Valkov

Massachusetts Institute of Technology

Hydrocarbon Scramjet Combustor Flows .......................................................................................... 87

Rockwell International, North American Aircraft Division

Chimera Domain Decomposition Applied to Turbomachinery Flow ................................................. 88

Co-investigator: Dale W. Thoe

General Motors Corporation, Allison Gas Turbine Division

Turbofan Compression System Simulation Using Chimera Domain Decomposition .......................... 89

Co-investigator: Dale W. Thoe

General Motors Corporation, Allison Gas Turbine Division

Multifunction Propulsion Systems ..................................................................................................... 90

Co-investigator: Timothy P. Nobel

LTV, Aircraft Division

Numerical Study of Thrust Vectoring Characteristics ......................................................................... 91

Rockwell International, North American Aircraft Division

National Aero-Space Plane External Burning Nozzle Studies ............................................................ 92

ICOMP/NASA Lewis Research Center
Principal

Investigator NAS Summary Page

Aeronautics: Rotorcraft

W. J. McCroskey

W. J. McCroskey

J. C. Narramore

Roger C. Strawn

H. Tadghighi

Aerodynamics and Acoustics of Rotorcraft ........................................................................................ 95

Co-investigators: J. D. Baeder, R. L. Meakin, V. Raghavan, and G. R. Srinivasan

NASA Ames Research Center

Calculations of High-Performance Rotorcraft .................................................................................... 96

Co-investigators: J. D. Baeder, E. P. N. Duque, S. K. Stanaway, and G. R. Srinivasan

NASA Ames Research Center

Rotorcraft Drag Prediction ................................................................................................................ 97

Co-investigators: A. G. Brand, J. J. Shillings, and D. W. Axley

Bell Helicopter Textron

Dynamic Adaption for Three-Dimensional Unstructured Grids ......................................................... 98

Co-investigators: Rupak Biswas and Michael Garceau

RIACS/Stanford University

Simulation of a Complex Three-Dimensional Internal Flow Field ...................................................... 99

Co-investigator: A. A. Hassan

McDonnell Douglas Corporation

Aeronautics: Turbulence

David E. Ashpis

J. Philip Drummond

Kai-Hsiung Kao

J. J. Kim

Linda D. Kral

Linda D. Kral

Parviz Moin

Parviz Moin

Transition in a Highly Disturbed Environment ................................................................................. ] 03

Co-investigator: Philippe R. Spalart

NASA Lewis Research Center/Boeing Commercial Airplane Group

High-Speed Turbulent Reacting Flows ............................................................................................ 104

Co-investigators: Peyman Givi, Cyrus K. Madnia, Steven H. Frankel, and Virgil Adumitroaie

NASA Langley Research Center/State University of New York at Buffalo

Simulation of a Turbine-Blade Cooling Passage .............................................................................. 105

Co-investigator: Meng-Sing Liou

ICOMP/NASA Lewis Research Center

Numerical Simulation of Turbulence .............................................................................................. 106

Co-investigators: Robert D. Moser, R. S. Rogallo, Nagi N. Mansour, and J. R. Chasnov

NASA Ames Research Center/Center for Turbulence Research

Receptivity, Transition, and Turbulence Phenomena ....................................................................... 107

Co-investigators: William W. Bower and John F. Donovan

McDonnell Aircraft Company

Turbulence Modeling for Three-Dimensional Flow Fields ............................................................... 108

Co-investigators: John A. Ladd, Mori Mani, and John F. Donovan

McDonnell Aircraft Company

Drag Reduction Mechanism by Riblets ........................................................................................... 109

Co-investigators: John Kim and Haecheon Choi

Center for Turbulence Research/Stanford University/NASA Ames Research Center

Turbulent Flow Over a Backward Facing Step ................................................................................. 110

Co-investigator: Hung Le

Center for Turbulence Research

xi
Principal

Investigator NAS Summary Page

J. Blair Perot

Man Mohan Rai

Michael M. Rogers

Bart A. Singer

Philippe R. Spalart

Turbulence Simulation on the Connection Machine ....................................................................... 111

Co-investigator: Paul Malan

Stanford University

Direct Simulations of Airfoil Flows .................................................................................................. 112

NASA Langley Research Center

Forced Plane Mixing Layers ............................................................................................................ 113

Co-investigators: Robert D. Moser, S. Scott Collis, and Chris Rutland

NASA Ames Research Center/Stanford University/University of Wisconsin, Madison

Development of a Turbulent Spot ................................................................................................... 114

Co-investigator: Ronald D. Joslin

High Technology Corporation/NASA Langley Research Center

Effect of Suction Holes .................................................................................................................... 115

Boeing Commercial Airplane Group

General" Astronautics, Astronomy, Atmospheric Science, Chemistry, Computer Science,

Electromagnetics, Fluid Mechanics, Hydrodynamics, Life Science, Reactive Flow,

Space Science, Structural Mechanics

Astronautics

Paul McConnaughey

Astronomy

Space Station Flow Analysis ............................................................................................................ 119

Co-investigators: Eric Stewart, Bruce Vu, and Lee Kania

NASA Marshall Space Flight Center/Sverdrup Technology, Inc.

Jeffrey D. Scargle

Bruce F. Smith

Atmospheric Science

Parallel Integration of N-body Gravitational Systems ....................................................................... 120

Co-investigators: Jeffrey N. Cuzzi, Anthony Dobrovolskis, Luke Dones, Robert Hogan,

Creon Levit, Mark Showalter, and Karl Young

NASA Ames Research Center

Numerical Experiments in the Formation and Evolution of Galaxies ............................................... 121

Co-investigators: Richard H. Miller and Thomas Y. Steiman-Cameron

NASA Ames Research Center/University of Chicago/University of California, Santa Cruz

Jeffrey R. Barnes

Christian L. Keppenne

James B. Pollack

Richard E. Young

Numerical Simulations of Baroclinic Instability ............................................................................... 122

Oregon State University

Orographically Forced Oscillations in the Martian Atmosphere ...................................................... 123

Jet Propulsion Laboratory

Dynamics of the Martian Atmosphere ............................................................................................. 124

Co-investigator: Robert Haberle

NASA Ames Research Center

Simulation of Volcanic Aerosol Clouds ........................................................................................... 125

Co-investigators: O. B. Toon and J. B. Pollack

NASA Ames Research Center

xii
Principal

Investigator

Chemistry

Richard L. Jaffe

Stephen Langhoff

Michael A. Wilson

James P. Wolfe

Computer Science

Thomas W. Crockett

Jerry C. Yan

Electromagnetics

Michael J. Schuh

Vijaya Shankar

Fluid Mechanics

Samuel Ohring

Hydrodynamics

Donald W. Davis

Life Science

Jeffrey R. Hammersley

NAS Summary Page

Propane-Air Combustion Mechanism ............................................................................................. 126

Co-investigators: Charles W. Bauschlicher, Harry Partridge, David Schwenke,

Christopher Dateo, and Stephen P. Walch

NASA Ames Research Center/Eloret Institute

Boost Phase Detection Studies ........................................................................................................ 127

Co-investigators: Charles Bauschlicher, Richard Jaffe, Timothy Lee, and Harry Partridge

NASA Ames Research Center

Computer Simulation of Astrophysical Ices ..................................................................................... 128

Co-investigator: Andrew Pohorille

NASA Ames Research Center

Radiation Effects in Intermetallic Compounds ................................................................................. 129

Co-investigators: Robert Averback and Huilong Zhu

University of Illinois, Urbana/Champaign

Performance and Scalability of Parallel Graphics Libraries .............................................................. 130

ICASE

Performance Monitoring of Parallel Programs ................................................................................. 131

Co-investigators: Philip J. Hontalas, Catherine H. Schulbach, Pankaj Mehra,

Sekhar R. Sarukkai, Melisa A. Schmidt, and Tarek S. Elaydi

Recom Technologies/NASA Ames Research Center/Lawrence Berkeley Laboratory

Computational Electromagnetics for Massively Parallel Processors .................................................. 132

Co-investigators: Alex C. Woo and John D'Angelo

NASA Ames Research Center/General Electric Corporation

Time-Domain Computational Electromagnetics .............................................................................. 133

Co-investigators: William F. Hall, Alireza Mohammadian, Chris Rowell, and Michael Schuh

Rockwell International Science Center/NASA Ames Research Center

Vortex-RingFree-Surface Interaction ............................................................................................... 134

Co-investigator: Hans J. Lugt

Naval Surface Warfare Center

Computational Hydrodynamic Performance Evaluation .................................................................. 135

Co-investigator: Keith C. Kaufman

General Dynamics, Electric Boat Division

Small Airway Fluid Dynamics ......................................................................................................... 136

Co-investigators: Rama Reddy, Dan E. Olson, Boyd Gatlin, and Joe F. Thompson

University of Arkansas/Medical College of Ohio/Mississippi State University

xiii
Principal

Investigator

Robert D. MacEIroy

Muriel D. Ross

Robert T. Whalen

Reactive Flow

K. Kailasanath

Space Science

Jeffrey N. Cuzzi

George S. Dulikravich

Structural Mechanics

Susan W. Bostic

Osama A. Kandil

Ahmed K. Noor

James H. Starnes, Jr.

NAS Summary Page

Computer Simulation of the Structure and Function of Membranes ................................................. 137

Co-investigators: Andrew Pohorille and Michael Wilson

NASA Ames Research Center

Visualizing Neurons in Three Dimensions ...................................................................................... 138

Co-investigator: Kevin Montgomery

NASA Ames Research Center/Sterling Software Systems

Large-Scale Numerical Simulations of Human Motion .................................................................... 139

Co-investigators: Marcus G. Pandy and Frank C. Anderson

NASA Ames Research Center/University of Texas at Austin

Multidimensional Burner-Stabilized Flames .................................................................................... 140

Co-investigator: G. Patnaik

Naval Research Laboratory/Berkeley Research Associates

Protoplanetary Particle-Gas Dynamics ............................................................................................ 141

Co-investigators: Anthony R. Dobrovolskis, Joelle M. Champney, and Robert C. Hogan

NASA Ames Research Center/University of California, Santa Cruz/Synernet, Inc.

Magnetoelectrohydrodynamics of Microgravity Crystal Growth ...................................................... 142

Co-investigator: Vineet Ahuja

Pennsylvania State University

Computational Structural Mechanics Applications .......................................................................... 143

NASA Langley Research Center

Simulation of Vertical Tail Buffet ..................................................................................................... 144

Old Dominion University

Buckling and Post-Buckling of Multilayered Composite Panels with Cutouts ................................... 145

Co-investigator: Jeanne M. Peters

NASA Langley Research Center/University of Virginia

Nonlinear Analysis of Damaged Stiffened-Fuselage Structure .......................................................... 146

Co-investigator: Vicki O. Britt

NASA Langley Research Center

xiv

k

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Numerical Solution of Three-Dimensional Maxwell Equations

Ramesh K. Agarwal, Principal Investigator

Co-investigators: Mark Shu and Mark R. Axe

McDonnell Douglas Corporation

Research Objective

To develop computational codes capable of predicting the

electromagnetic (EM) signature of an aerospace vehicle. Such

a code would be of great value in the design and development

of fighter aircraft and missiles because a greater number of

candidate configurations could be considered and evaluated

with a reduction in cost and design time.

Approach

Three-dimensional (3-D) Maxwell equations were solved by

employing mature computational fluid dynamics technology.

Specifically, a compact, higher-order, finite-volume, timedomain/frequency-domain, multizone method was used to

compute the EM field about scattering bodies.

Accomplishment Description

The EM scattering code was extended to calculate the scattering

from geometrically and electrically complex two-dimensional

(2-D) and 3-D objects. Accomplishments included 2-D EM

scattering calculations from Iossless and Iossy dielectric objects

such as dielectric cylinders, coated cylinders, and treated ducts.

Also accomplished were the 3-D EM scattering calculations from

perfectly conducting bodies such as spheres, cone spheres, finite

cylinders, and almonds. Especially noteworthy were the 3-D EM

scattering calculations from an F-15.

Significance

The project represents the development of a leading-edge

technology in computational EM, and it will have substantial

impact on the development of both fighter aircraft and missile

programs.

Future Plans

A CFD-based Maxwell solver will be developed and validated to

calculate the radar cross sections of complete, real-world 3-D

configurations.

Publication

Huh, K. S.; Shu, M.; and Agarwal, R. K.: A Compact High-Order

Finite-Volume Time-Domain/Frequency-Domain Method for

Electromagnetic Scattering. AIAA 30th Aerospace Sciences

Meeting, Reno, Nev., AIAA Paper 92-0453, Jan. 1992.

Amplitude of the magnetic field about a Iossy coated perfectly

conducting cylinder. The circular white line represents the edge

of the dielectric. Red represents maximum amplitude and blue

represents minimum amplitude.

Amplitude of the electric current on the surface of a perfectly

conducting sphere. Red represents maximum amplitude and

blue represents minimum amplitude. The white contour lines

represent electric current contours of equal amplitude in the

near field.

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Aeronautics/Basic Research 3
Simulation of Acoustic Scatter

Harold L. Atkins, Principal Investigator

NASA Langley Research Center

Research Objective

To evaluate the performance of a newly developed nonreflective

inflow/outflow boundary condition for multi-dimensional

unsteady flows.

Approach

A fifth-order shock-capturing finite-difference method was used

to simulate flow over a sphere that either produced sound or

reflected an incident sound wave. The boundary condition

applied at the outer boundary was designed to allow waves to

exit the domain without reflection and, at the same time, allow

desired inbound waves to be prescribed. The boundary condition was evaluated by comparing two solutions in which the

outer boundary was applied at different distances from the

sphere.

Accomplishments Description

Simulations were performed for oscillating spheres and spheres

interacting with plane waves. The figures show an acoustic

plane wave that propagates from right to left and impinges upon

the sphere to produce a scatter wave. The upper half of each

figure corresponds to a case with the outer boundary at

13 sphere radii; in the lower half, the outer boundary is 9 radii.

The boundary condition was designed on the premise that only

isolated acoustic and convective waves of a single, known

orientation exist. This test case consisted of two sets of waves

with different orientations and gave a fair indication of how the

boundary condition performs at off-design conditions. At the

resolution illustrated in the figures, the difference between the

two solutions was roughly three orders of magnitude below the

amplitude of the scatter wave. Grid-refinement studies were

performed to investigate the accuracy properties of the highorder method and the nonreflective boundary conditions. A

typical case required about 40 minutes processing time to allow

the plane wave to cross the domain and establish a periodic

scatter wave and required less than I 0 megawords of memory.

Additional execution time was required when temporal analysis

of the scatter wave was desired.

Significance

The present boundary condition allows the computational

domain to be reduced to the minimum region required to

simulate the nonlinear, near-field solution. This simulation

should reduce the cost of predicting noise produced by aerodynamic flows.

Future Plans

The high-order method will be coupled with a Kirchhoff integral

method for predicting the far-field noise. The initial tests will

determine how much simulation data must be saved for input to

the Kirchhoff method. This testing will be followed by a simulation of a high-speed jet.

(a)

(b)

Perturbations produced by a plane wave impinging on a sphere.

The upper half of each figure shows a case with the outer

boundary at 13 sphere radii; the lower half of each figure shows

the outer boundary at 9 sphere radii. Both the (a) density and

(b) v component of velocity are on the order of +0.001.

4 Aeronautics/Basic Research
A Three-Dimensional Implicit Unstructured

Timothy J. Barth, Principal Investigator

Co-investigator: Samuel W. Linton

NASA Ames Research Center/Sterling Software Systems

:

Research Objective

To develop an efficient numerical scheme for solving the Euler

equations of gas dynamics on unstructured meshes using a

preconditioned minimum residual matrix solver•

Approach

The Euler equations were discretized on arbitrary tetrahedral

meshes using a finite-volume approach with upwind dissipation

added to stabilize the scheme. The algorithm exhibited secondorder spatial accuracy on sufficiently smooth meshes. An

implicit time-stepping strategy, which approached Newton's

method for large time steps, was used. This strategy required the

solution of a sequence of large sparse matrix equations• To solve

the equations, a generalized minimum-residual method with

incomplete lower-upper preconditioning was used.

Accomplishment Description

A computer code was designed, developed, and implemented

on a range of computer architectures including vector and

parallel machines. The computer code was used in the computation of aerodynamic flows about wings and helicopter blades at

subsonic and transonic speeds. Using the implicit solution

strategy, rapid convergence to steady state was demonstratecl.

The first figure shows the temporal convergence history of the

algorithm for subsonic flow over multiple spheres.

Significance

The use of unstructured meshes permitted the calculation of

aerodynamic flow over complex geometries (see the second

figure). Moreover, the improved efficiency of the algorithm

permitted the calculation of these flows using modest resources.

Future Plans

Plans include the modification of the code for computation of

viscous flow including the effects of turbulence• In addition, the

code will be modified to support computations in a nonheterogeneous computing environment.

Euler Solver

10"2

10 -4

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._ 10-6

E

0

,- 10-8

E

.__ 10 -lo

x

_ 10-12

10-14

- ?

I I I

0 10 20 30

Iterations

I I

40 50

Convergence history of the implicit scheme. The dashed and

solid lines depict the convergence history of the scheme with

first- and second-order space differencing, respectively•

Surface mesh on the wing and fuselage _ortion of a Boeing 737

aircraft with high-lift devices deployed•

Aeronautics/Basic Research 5
Simulation of Supersonic Jet Screech

Alan B. Cain, Principal Investigator

Co-investigators: William W. Bower and William W. Romer

McDonnell Douglas Aerospace

Research Objective

To simulate supersonic jet screech and provide detailed data for

modeling receptivity and shock-vortex interactions.

Approach

Several algorithms were examined. A formulation based on

upwind-biased, conservative, finite-volume spatial differentiation

in combination with Runge-Kutta time integration produced the

best results.

Accomplishment Description

Self-sustained screech oscillations were simulated and reasonably matched the frequency measured in NASA Langley

experiments. Both axisymmetric and rectangular nozzle flows

were examined. The fully expanded jet Mach numbers considered were 1.05, 1.20, and 1.60 for nozzles designed for perfect

expansion at Mach numbers of 1.00, 1.30, and 1.35, respectively. Analysis is in progress for the acoustic spectra and

direction characteristics. In the rectangular case, the initialization of the flow field produced an initial varicose mode that

gradually evolved into a dominant sinuous mode consistent

with the predictions of linear theory. Typical runs required

64 megawords of memory and 20-100 Cray Y-MP hours.

Significance

The screech prediction tool is important for the design of tactical

aircraft. Eliminating screech substantially reduces nozzle maintenance costs. A preliminary design tool must generate analysis

solutions quickly on a workstation; thus good models of the

basic phenomena are essential. The present simulations are a

starting point for detailed information to guide analytical

modeling of the shock-vortex interaction and the nozzle lip

receptivity.

Future Plans

Further analysis of the current results is needed to determine the

spectral and directivity characteristics. New simulations for a

round jet to examine mode switching will help define current

simulation capabilities.

Publication

Cain, A. B.: Computation of Supersonic Jet Screech. Proceedings

of the ASME Fluids Engineering Forum on Computational Aeroand Hydro-Acoustics Meeting, Washington, D.C., June 1993.

Pressure and vorticity contours for a rectangular nozzle: design

Mach number = 1.35, fully expanded jet Mach number = 7.60,

red = maximum pressure, blue = minimum pressure.

6 Aeronautics/Basic Resea_h
Efficient Constrained Aerodynamic Design

Richard L. Campbell, Principal Investigator

Co-investigators: Leigh A. Smith and Raymond E. Mineck

NASA Langley Research Center

Research Objective

To evaluate a new approach to constrained aerodynamic design

of complex configurations at multiple design points using

advanced computational fluid dynamics (CFD/codes.

Approach

The objective is approached in two parts that will eventually be

merged: (1) evaluation of the Constrained Direct Iterative Surface

Curvature (CDISC) design method for three-dimensional (3-D)

applications and (2) investigation of automated multipoint design

approaches. The CDISC method generates and modifies target

pressure distributions to meet flow and geometry constraints. It

includes an option for minimizing an objective function by

adjusting the constraint values during the design process. The

basic direct iterative surface curvature (DISC) design method

is used to modify the configuration to match the new target

pressure distributions. Several approaches to employ the CDISC

method at multiple design points and determine what additional

adjustments to the constraints may be required are being investigated. For these studies, the CDISC method is coupled with a

thin-layer Navier-Stokes code (TLNS3D) and an unstructuredmesh Euler code.

Accomplishment Description

The design codes were applied to configurations at subsonic,

transonic, and supersonic speeds. These cases included several

dual-point design studies with the goal of defining an automation strategy for efficient multipoint design. Results for a redesign

of a subsonic transport wing at a Mach 0.82 and a lift coefficient

of 0.55 are shown in the figures. The solutions were obtained on

a 193 x 49 x 49 grid, which required about 24 megawords of

memory and about 1.25 Cray-2 hours to reach the convergence

criteria of four orders of magnitude reduction in the residual.

Mach number contours for the baseline wing are shown in the

first figure. The yellow and red regions indicate Mach numbers

greater than 1.25 and 1.30, respectively. The shock just aft of

midchord extends to the symmetry plane, with the maximum

strength occurring between about 25 and 50 percent of the

semispan. The Mach number contours for the CDISC design

(second figure) indicate that the shock has been significantly

weakened over the inner half of the wing (all Mach numbers

are below 1.30). This weakening and a slight redistribution of

the spanwise loading reduced the inviscid drag of the wing by

about 12 percent. This redesign involved 12 design cycles and

required about 16 percent more time than the baseline analysis.

Significance

Results indicate that the CDISC method could provide an

efficient alternative to numerical optimization for constrained

design of aerodynamic configurations. The CDISC method

eliminates the need to define or modify the target pressures

required by the DISC approach and can reduce the central

processing unit time required.

Future Plans

Multipoint design approaches will be evaluated at cruise and

off-design conditions that include separated flow. The CDISC

method will be installed in the multiblock version of TLNS3D

for more complex geometries.

Mach number contours for baseline wing, Mach = 0.82.

Mach number contours for CDISC redesign, Mach = 0.82.

Aeronautics/Basic Research 7
Computation of Helicopter Rotor and Wake Flows

Frank Caradonna, Principal Investigator

Co-investigators: K. Ramachandran, J. Bridgeman, and D. Poling

U.S. Army Aeroflightdynamics Directorate, AVSCOM/Flow Analysis, Inc./Woodside Summit Group, Inc./The Boeing Company

Research Objective

To use large-scale computational models of the limiting flow

phenomena of helicopters and tilt rotors to determine the various

flow losses and vibratory loading mechanisms with sufficient

accuracy to permit reliable design and to obviate the need for

extensive testing.

Approach

Individual flow models were assembled and validated to build a

total flow solver. Two unique methods were used in this work.

Vortex embedding--an extension of potential methods that

permits the free convection of the wake with no dissipation and

no grid restrictions--was used to predict the hover and forwardflight performance of rotors. Vorticity confinement--an extension of Euler/Navier-Stokes codes that counteracts the effects of

numerical dissipation--was used to model wake formation and

blade-vortex interactions. Combined, the two methods provide

accurate near-blade solutions and a wake-convecting, compressible "far" field that includes interaction effects.

are the most accurate wake solvers of any CFD method and

provide a detailed analysis of rotor drag. This should provide a

new capability to design and evaluate advanced rotor configurations with minimal experimental development.

Future Plans

Computational design studies will be performed. Accuracy and

efficiency improvements will be made in the HELIX-II forwardflight code. Vorticity confinement methods will be applied to

improve blade-vortex interaction computations and viscous flow

effects. Confinement solvers will be combined with vorticity

embedding solvers to provide a comprehensive rotor flow

method.

Accomplishment Description

The most extensive computational fluid dynamics (CFD) hover

validations were performed using the HELIX-I code to compute

the performance of the UH-60A rotor. These methods were

applied to the analysis of the hover performance of the new

RAH-66 helicopter and revealed unsuspected regions of

rotor power sensitivity. These computations required about

5 megawords of memory and 6 Cray-2 hours to compute a

typical performance polar. The first forward-flight loads computations and comparisons with flight data were made using the

HELIX-II code to compute the flow on the AH-1G. A typical

run required about 15 megawords of memory and about

10 Cray-2 hours. The first vortex formation and blade-vortex

interaction computations were performed using the code

TURNSCONFINE. Confinement efficiently maintained the

identity and strength of vortex structures and increased the

higher harmonic loading of the blade-vortex interactions. The

runs required about 10 megawords of memory and 4 Cray-2

hours.

Significance

To improve rotor performance, a more detailed knowledge of the

various drag mechanisms is required. The present computations

(a)

(b)

Vorticity contours of a vortex generated by a wing tip computed

(a) without and (b) with the use of confinement.

Wake structure evolution of a starting rotor solution. Periodicity is achieved in 6 revolutions, by which time

blade-vortex interactions are clearly seen.

8 Aeronautics/Basic Research
High-Incidence Static/Dynamic Roll Computations

Neal M. Chaderjian, Principal Investigator

Co-investigators: John Ekaterinaris, Yuval Levy, and Lewis B. Schiff

NASA Ames Research Center/Naval Post Graduate School/Stanford University

Research Objective

To develop a validated computational methodology to predict

and investigate high-angle-of-attack vortex aerodynamics for

static-roll and dynamic-roll motions of simple wing body

geometries.

Future Plans

A similar study will be carried out where vortex breakdown

occurs (30 degrees angle of attack) and free-to-roll motions,

including wing rock, will be computed.

Approach

Time-dependent, three-dimensional, Reynolds-averaged

Navier-Stokes (RANS) equations were utilized to compute

high-incidence vortical flows about a 65-degree-sweep delta

wing at 15 degrees angle of attack. The Navier-Stokes simulation (NSS) code integrated the RANS equations using either

the implicit Beam-Warming algorithm or a diagonal version

developed by Pulliam and Chaussee. Turbulent flow conditions

were treated with the Baldwin-Lomax turbulence model and the

Degani-Schiff modification, which accounted for the cross-flow

separation.

Accomplishment Description

The NSS code was used to compute the viscous flow about a

delta wing in roll with free-stream Mach number = 0.27, angle

of attack = 15 degrees, and Reynolds number = 3.67 million

(based on the wing root chord). Investigations were carried out

at several static roll angles up to 42 degrees of roll and a largeamplitude (40 degree), high-rate (7 hertz (Hz)) forced roll

motion. The first figure shows the leeward-side vortices with

helicity-density contours for the dynamic case at its maximum

roll angle. Primary, secondary, and tertiary vortices are evident.

The flow topology was verified by experimental oil-flow

patterns. Static and dynamic computational surface pressures

were in excellent agreement with experimental data, as were the

forces and moments. The second figure compares the computational and experimental rolling moment coefficients for the static

and dynamic cases. The negative slope of the static rolling

moment curve indicates that the wing is statically stable in roll

under the flow conditions. The dynamic rolling moment exhibits

a significant hysteresis due to rate effects. The computational

and experimental rolling moments move along their respective

curves in a counterclockwise direction, indicating that the

motion is damped. The areas under these curves represent the

damping energy and agree within 3 percent. No vortex breakdown was observed in the computations or the experiment. Each

static case was solved on an 800,000 point grid with a diagonal

algorithm and required 26 hours (8,000 steps) on a Cray Y-MP.

Significance

Highly accurate computational results were obtained for a delta

wing in static and dynamic roll. This step is the first toward an

experimentally validated computational method for computing

static-roll and dynamic-roll motions of aircraft, which will

eventually be applied to wing rock.

Helicity-density contours at roll angle (_) = -40 degrees. Flow

conditions: Mach = 0.27, Reynolds number = 3.67 million,

dynamic roll amplitude = 40 degrees, and frequency = 7 Hz.

Green contours = clockwise flow rotation and blue contours =

counterclockwise flow rotation.

.050

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-.025

-.050

Static experiment

- _ Dynamic experiment

[] Static, NSS

-- -- -- Dynamic, NSS

I I I I I I I I I I

-50 -40 -30 -20 -10 0 10 20 30 40 50

Deg (cp)

Static and dynamic rolling moment comparisons between

computation and experiment. Flow conditions: Mach = 0.27,

Reynolds number = 3.67 million, dynamic roll amplitude =

40 degrees, and frequency = 7 Hz.

Aeronautics/Basic Research 9
Cavity Aeroacoustic Loads Prediction

Richard D. Crouse, Principal Investigator

Northrop Corporation

Research Objective

Experiments performed on aircraft weapons bay cavities have

demonstrated the presence of powerful acoustic pressure waves.

An effort was established to develop and execute a set of

procedures to calculate the unsteady aeroacoustic flow-field

environment using computational fluid dynamic methods.

Approach

The flow field within and around the cavity and the full aircraft is

calculated using existing three-dimensional (3-D) Navier-Stokes

flow codes. The time-varying pressure field is transformed with a

fast-Fourier transform to a frequency domain. This procedure

will allow the identification of the frequencies and the amplitude

of the varying pressure.

Accomplishment Description

A 3-D grid was generated around the U.S. Air Force Innovative

Weapons Cavity (IWC) model, including the open weapons bay

cavity, consisting of 661,586 grid points. An optimum blocking

structure was defined using six blocks. The viscous wall spacing

was determined and correctly resolved the boundary layer.

Computations on the model were completed and gave a steadystate solution. The time-accurate solution was started from this

steady solution. Early results indicated excessive damping,

which was probably due to a problem with the turbulence

model, and the solution was halted. Only 10 Cray-2 hours were

used and each run required 32 megawords of memory.

Significance

Acoustic loads prediction is needed when considering the

increased performance levels of advanced tactical air vehicles.

The study results provide an extremely useful tool to aid in new

vehicle design. This computational means will enable solutions

to be calculated on a variety of topologies and over a vast flight

regime. It will not make empirical results present-case dependent, nor simply recalculate old data. Rather, it will be applicable in all stages of aircraft design, such as further analysis of

existing geometries, development of present configurations, and

research in future programs.

Future Plans

The turbulence model in the code will be corrected for timeaccurate flow. The time-varying solution of the flow in a cavity

in an entire aircraft will be completed. The results will be

compared to wind tunnel test results of this model in early 1994.

Three-dimensional grid lines around the U.S. Air Force IWC

model.

I 0 Aeronautics/Basic Research
High-Fidelity Space Shuttle Simulation

Daniel F. Dominik, Principal Investigator

Co-investigators: K. Rajagopal, S. Vuong, J. Wisneski, C. Olling, G. Hock, and J. Sikora

Rockwell International, Space Systems Division

Research Objective

To develop a full-scale, high-fidelity asymmetric model for the

Space Shuttle Launch Vehicle (SSLV) and to predict the transonic

aerodynamic loads during ascent by performing Navier-Stokes

computations at full-scale flight Reynolds number.

Future Plans

Computations with variable specific heat gaseous SRB and SSME

plume simulations are in progress now and are expected to

improve further the correlation with test data.

Approach

A model with many fine geometry details of the SSLV was built

using the ICEM/computational fluid dynamics (CFD) computer

assisted design (CAD) and grid software. The multizonal grid was

smoothed using the GRIDGEN3D software. Most, but not all, of

the interzonal boundaries are aligned. The numerical solutions

were obtained using the Reynolds-averaged Navier-Stokes

formulation of the USA-RG3D code with the algebraic

Baldwin-Lomax turbulence model. Initial calculations were

without plume simulation; later calculations will include simulation of the three Space Shuttle main engine (SSME) plumes

and the two solid rocket booster (SRB) plumes. Results will be

correlated with available flight test and wind tunnel data.

Accomplishment Description

The SSLV model simulated the forward and rear attached

hardware connecting the Orbiter, the external tank (ET), and the

two SRBs. The primary protuberances on the body surfaces,

including the liquid oxygen feedline on the right side of the ET,

were also modeled. These geometry details enhanced the

simulation accuracy of the multibody flow field. This complete,

asymmetric grid model contained 237 zones and approximately

5 million points. The USA-RG3D flow solver was used to

simulate the turbulent, flight Reynolds number transonic flow

over the geometry. The model was broken into 6 sections and no

section required more than 64 megawords of main computer

memory. Each section had overlapping zones and the solutions

were updated by cycling through the sections to ensure proper

communication. Numerical solutions were computed for the

redesigned solid rocket motor and advanced solid rocket motor

configurations. The model enabled simulation of sideslip

conditions. For each configuration, solutions were computed at

Mach = 1.25 and Mach = 1.05 with 0 and 4 degrees of sideslip.

Significance

There was good correlation with experimental results. The

detailed definition of the airload distributions and the insight

into the complex three-dimensional flow physics illustrate the

benefits that can only be obtained from this type of detailed

CFD simulation.

Surface pressure coefficient at Mach = 1.25, angle of attack =

-5.1 degrees, sideslip = 0 degrees (without plume simulation)

(gray = _9.75 and white = 0.75).

Pressure coefficient between wing and SRB at y = 250 in. for

Mach = 1.25, angle of attack = -5.1 degrees, sideslip ---

0 degrees (without plume simulation).

Aeronautics/Basic Research 11
Aerodynamic Optimization of Supersonic Aircraft

Thomas A. Edwards, Principal Investigator

Co-investigators: Samson H. Cheung and I. C. Chang

NASA Ames Research Center/MCAT Institute

Research Objective

To apply computational fluid dynamics (CFD) technology to

aerodynamic design problems for the High-Speed Civil Transport

(HSCT). The problems include aerodynamic optimization and

sonic boom minimization.

Approach

For aerodynamic design optimization, the CFD code was used as

an analysis module for aerodynamic performance estimates. A

numerical optimization routine based on sequential quadratic

programming directed design modifications based on the

sensitivity gradient vector and imposed constraints. For sonic

boom minimization, a similar approach was used, but the

objective function was related to the sonic boom signature. This

function required additional analysis modules that would accept

Drag Rise Reduction

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Mach Number

Comparison of a supersonic/transonic wing-body surface

pressure distribution before and after drag minimization redesign. Contours: Mach = 1.0, a = 3.5 degrees.

the CFD solution in the near field and extrapolate the pressure

disturbance to ground level. The design variables were related to

the aircraft equivalent area distribution, which was related in a

non-unique way to the actual aircraft geometry.

Accomplishment Description

Both optimization processes were established and applied to

HSCT design problems. Design optimization required many

function evaluations so it was essential that the computer time

per flow solution was kept to a minimum. This was done by

vectorizing the code to the greatest extent possible, using

efficient solution algorithms and moderately resolved grids. A

typical solution for inviscid flow about an HSCT was obtained

in about 3 minutes of Cray-2 time. However, a comparatively

simple design problem required hundreds of function evaluations, so the computer time required for the optimization process

was 5-1 0 hours or more, depending on the number of design

variables and the complexity of the optimization problem.

Significance

Automated design optimization tools allow the computer to

perform what would otherwise be a time-consuming, laborintensive process, freeing the designer to explore a broader range

of design issues, evaluate new concepts, and achieve performance improvements that would otherwise be unattainable.

Future Plans

The design capability will be applied to advanced HSCT

concepts to explore the tradeoffs between aerodynamic performance and sonic boom loudness. The effects of structural

deflections will be included in aerodynamic shaping, and

multidisciplinary design optimization tools will be developed

and applied to HSCT configurations.

4r Baseline

-- -- -- New Design

Psf "_

0 F _ _'_

0 .1 .2 .3

Time (sec)

Comparison of sonic boom signatures at ground level for

baseline HSCT and the new design optimized for low sonic

boom.

12 Aeronautics/Basic Research
Three-Dimensional Turbulent Double-Fin Interactions

Datta Gaitonde, Principal Investigator

Co-investigator: Joseph Shang

WL/FIMM, Wright Patterson AFB

Research Objective

To examine the flow-field structure arising out of the interaction

of a turbulent boundary layer with intersecting shocks of equal

strength. The ability of the chosen method to provide practical

engineering estimates of mechanical and thermal loading is also

investigated.

Approach

The three-dimensional, mean compressible Navier-Stokes

equations in mass-averaged variables were solved with a highresolution scheme for the inviscid fluxes, central differencing

for the viscous fluxes, and an implicit relaxation-based time

integration method. The examination was focused on a Mach 8

interaction for which high-quality experimental data were

available for validation. Sensitivity of the solution to mesh

resolution was examined.

Future Plans

The present interaction will be verified at different values of

relevant parameters (including shock strength and Mach numbers).

Preliminary indications are that the overall structure remains

similar, but the details exhibit significant variation--particularly

in local topological features. The sensitivity of the flow structure

to variations in turbulence modeling will also be examined.

Publication

Gaitonde, D.; and Shang J.: Calculations on a Double-Fin

Turbulent Interaction at High Speed. AIAA Paper 93-3432,

Applied Aerodynamics Conference, Monterey, Calif., Aug. 1993.

Accomplishment Description

The flow fields past two symmetric configurations with 15- and

10-degree fins were simulated. With the exception of heat

transfer rates in regions of large crossflow, the agreement with

experiment was very good. An analysis of the 15-degree configuration was performed. A complex flow field was revealed with

interesting features pertaining to the kinematic structure including the existence of two relatively small secondary vortices near

the center of the interaction.

Significance

Three-dimensional turbulent interactions are common in

supersonic applications such as inlets and wing-body junctures.

The accompanying effects are often undesirable. A detailed

understanding of the kinematic and dynamic structure of flow

fields will provide the insight needed for control and may thus

suggest improvements in the design and performance of practical

devices.

Ribbons exhibiting flow features: BL = boundary layer, VI =

vortex interaction, SV = secondary vortex, PV = primary vortex.

Aeronautics/Basic Research 13
Design-By-Optimization Method

Philip B. Gingrich, Principal Investigator

Rockwell International, North American Aircraft Division

Research Objective

To develop and validate a three-dimensional (3-D) numerical

optimization design method applicable to complete aircraft

configurations.

Approach

A multiblock, implicit time-marching Euler and Navier-Stokes

solver is coupled with a gradient-based, constrained optimization technique. Perturbation functions, with an associated design

variable, operate on the surface grid. The optimizer determines

the design variable set that minimizes a composite function

formed with the objectives and constraints. To simulate an

inverse procedure, the objective is the root-mean-square

difference between the current and a specified target pressure

distribution.

Accomplishment Description

A 3-D design-by-optimization method was developed for

application to complete aircraft configurations. The design

capability has been applied to the F-16XL-2 with a baseline

laminar flow control (LFC) glove. A coarse (160,000 points)

multiblock Euler grid was developed. The baseline analysis at

Mach 1.9 and 3 degrees angle of attack indicated swept isobars

that would promote boundary layer transition due to crossflow.

To increase laminar flow potential, a flat-top-design pressure

distribution was specified for the forward midspan region of the

wing. Design studies were conducted at several span stations to

determine the appropriate shape functions and constraints. The

required number of global optimizer iterations was typically of

the order of the number of design variables. For a design at one

wing station, 0.5-2.0 Cray Y-MP hours were required, using

20 megawords of memory.

Significance

Nonlinear design methods have not kept pace with computational fluid dynamics analysis capability. Emerging direct inverse

and numerical optimization techniques will reduce design cycle

time for projected subsonic and supersonic aircraft.

Future Plans

The feasibility of using the numerical optimization technique to

minimize drag will be investigated. Appropriate shape function

sets and constraints will be established.

j"

(a)

J

(b)

Upper surface isobars on F-16XL-2 with LFC glove at Mach I. 9,

3 degrees angle of attack for the (a) constant-pressure optimization inverse and (b) baseline. Contour increment = 0.02.

] 4 Aeronautics/Basic Research
F/A-1 8 Store Carriage Analyses

David Halt, Principal Investigator

Co-investigators: Mori Mani and Todd Michal

McDonnell Aircraft Company

Research Objective

To investigate the viability of unstructured grid methods and

Chimera methods in the modeling of a complex, fluid flow field

around an F/A-18 with an array of stores in the stowed position.

Approach

In the unstructured grid approach, the Delaunay grid generation

method was used along with a finite-element flow solver,

MDFENS. In the Chimera grid approach the "line intersection"

method was used to cut the hole in the background grid (the

wing grid zone) and set up the fringe boundaries for the

Navier-Stokes time-dependent (NASTD) multiblock flow solver.

Accomplishment Description

The unstructured grid for the full configuration F/A-18 with pylon

and store used 38,162 grid points and 220,118 tetrahedra in the

half plane. The surface Mach contours are shown in the first

figure for an Euler solution with a free-stream Mach number of

0.84 and 0 degrees angle of attack. The Chimera solution used

28 zones. The pylons and a store were constructed using a

Chimera grid and the remaining zones were constructed from

a nonoverlapping grid. The outboard pylon and store grids

intersected and a hole was generated in each zone. Both the

pylon and the store create a hole with the wing grid. The Euler

solution is shown in the second figure for a 0.7 free-stream Mach

number and 0 degrees angle of attack. The solution is smooth

and continuous across the zone boundaries.

Significance

The unstructured grid Euler solution demonstrates the level

of complexity that can be modeled using unstructured grid

technology with flexible control of local grid spacings. The

Chimera solution demonstrates the three-dimensional capability

using the NASTD flow solver and the flexibility to include or

exclude a number of stores and pylons or part of the geometry

from the solution without generating a new grid each time. This

method will reduce the cost and increase the efficiency in

obtaining a new solution each time.

Future Plans

Each methodology for the unsteady case of store separation for

the F/A-18 aircraft will be demonstrated.

F/A- 18 unstructured grid surface Mach contours at free-stream

Mach number 0.84 and 0 degrees angle of attack; blue =

Mach 0.5 and below, red = Mach 1.0 and above.

NASTD Chimera grid surface pressure contours on a fighter

geometry with pylon, store, and tip missile; black = 0.37,

white = 1.34.

Aeronautics/Basic Research 1_5
Lifted Ethylene Jet Diffusion Flames

Carolyn R. Kaplan, Principal Investigator

Co-investigator: Elaine S. Oran

Naval Research Laboratory

Research Objective

To demonstrate the capability to simulate the phenomenon of

flame lift-off and to study the mechanism by which a lifted flame

is stabilized.

Approach

The model solved the multi-dimensional time-dependent

reactive-flow Navier-Stokes equations coupled with submodels

for soot formation and radiation transfer to simulate unsteady

axisymmetric ethylene air-jet diffusion flames. The convection

terms were solved using the barely implicit correction to the flux

corrected transport algorithm. The effects of thermal conduction,

molecular diffusion, and viscosity are included through explicit

finite-differencing formulations. The radiative heat flux was

found from a solution of the radiation transfer equation using the

discrete ordinates method. The soot formation and growth model

are based on a set of coupled ordinary differential equations

describing soot nucleation, surface growth, and coagulation.

Accomplishment Description

Simulations were conducted for jet inflow velocities ranging

from 5 meters/second to 100 meters/second for three different

fuel-jet dilutions. The computations show that (1) the base of the

flame stabilizes in a stoichiometric region, (2)the flame lift-off

height increases linearly with jet exit velocity, and (3) the height

at which the flame stabilizes increases with fuel-jet dilution.

In all cases, the laminar flame base is anchored in a turbulent

region of the gas jet. The stabilization heights compare favorably

with experimental measurements reported in the literature. The

computations support an extinction theory of flame stabilization,

in which the flame is quenched when the rate of heat conduction to the sides of the flame exceeds the rate of heat production

due to chemical reaction.

Significance

The physical mechanisms responsible for flame stabilization are

still controversial. Past work has involved experimental measurements of flame lift-off height, and stabilization mechanisms have

been deduced from these experiments. This work constitutes the

first numerical simulation of flame lift-off and may provide more

insight into the stabilization mechanisms.

Future Plans

Further analysis of the stabilization mechanism will be deduced

from the simulation data. The model will be used to evaluate the

effect of co-flow velocity on the dynamics of jet diffusion flames.

Publication

Kaplan, C. R.; Baek, S. W.; and Oran, E. S.: Dynamics of

an Unsteady Buoyant Ethylene Jet Diffusion Flame. AIAA

Paper 93-0109, Jan. 1993.

30 m/nc 60 m/xc 100 n_sec

14

cm

0 3

cm

Temperature contours show the effect of jet inflow velocity on

flame lift-off height.

16 Aeronautics/Basic Research
Unstructured Grid/Flow Solver Calibration

Steve L. Karman, Jr., Principal Investigator

Co-investigators: Doug Howlett and Elizabeth A. Robertson

Lockheed Engineering and Sciences Company

Research Objective

To demonstrate the capability to numerically model complex

three-dimensional (3-D) time-dependent, multibody flow fields

using unstructured grid methods.

Approach

An unstructured Cartesian grid computational fluid dynamics

(CED) code was used to analyze complex two-dimensional (2-D)

and 3-D multibody flow fields. The grids were generated by

recursively refining a root Cartesian cell that encompassed the

entire domain. The subdivision was initially based on user

defined boundary spacing. Later, the grid was refined based on

flow-field gradients or movement of a body through the domain.

The flow solver used upwind differencing. The 2-D solver used a

point-implicit time-integration scheme that made use of Jacobi

subiterations for improved stability. The 3-D flow solver used a

multistage Runge-Kutta time-integration scheme. Multigrid can

be used for steady-state problems.

Accomplishment Description

Various 2-D problems were solved to prove the concept.

One of the more difficult 2-D problems attempted was the

AGARD 03 test case, which consisted of a NACA 0012 airfoil

run at a free-stream Mach number of 0.95 and an angle of attack

of 0 degrees. The flow field was symmetric from top to bottom

and contained a normal shock in the wake. Grid convergence

studies performed by other researchers showed the normal shock

to appear 3.35 chord lengths aft of the airfoil trailing edge. This

investigation placed the shock at 3.4 chord lengths back, a

difference of 1.5 percent. In the first figure, the final grid with

33,517 active cells and normalized velocity contours from this

study are given. The primary focus of the 3-D work was the gridgeneration process. Although the grid-generation procedures are

still being developed, some preliminary solutions were computed. In the second figure, the inviscidly computed, steady-state

velocity magnitude contours on the surface and on the symmetry

plane of a 3-D generic wing/store configuration are shown.

(a) (b)

Results of AGARD 03 test case for a 2-D code; (a) adapted grid

and (b) velocity magnitude.

Significance

The development of an unstructured CFD capability will greatly

reduce the analysis time for complex 3-D geometries and flow

fields. Automatic grid generation will permit the timely evaluation of parametric studies of complex geometries and the

analysis of phenomena such as store separation.

Velocity magnitude contours on the surface and the symmetry

plane of a generic wingstore configuration.

Future Plans

The calibration of the 3-D Cartesian-based unstructured-grid

code will continue. Various steady-state solutions for complex

geometries will be analyzed. The time-accurate moving body

capability will be added and calibrated. Long-term plans are to

include a viscous analysis capability.

Aeronautics/Basic Research 1 7
Simulation of Flow About an Airborne Observatory

Stephen P. Klotz, Principal Investigator

Co-investigator: Christopher A. Atwood

NASA Ames Research Center

Research Objective

To assess, using computational fluid dynamics (CFD), effects

of an open telescope bay on the aerodynamics and aero-optics

of a Boeing 747-SP, the aircraft platform for the Stratospheric

Observatory for Infrared Astronomy (SOFIA).

Approach

Computations were performed on a sting-mounted aircraft

without empennage. The telescope was mounted behind the

wings and the engines were powered. Solutions were computed

by OVERFLOW and used the diagonalized Beam-Warming

numerical scheme and Baldwin-Lomax turbulence model in

the code. A shear-layer version of the Baldwin-Lomax model

simulated the effects of turbulence in the cavity-port shear layer.

Accomplishment Description

The original design concept for SOFIA placed the telescope

cavity forward of the wings and immediately behind the cockpit.

Recent work focused on computations with the telescope

mounted behind the wings because locating the cavity there

required fewer structural modifications. Nevertheless, an aftmount cavity configuration has potential problems associated

with (1) radiative scattering from the engine plumes, (2) a

boundary layer that is optically thicker, and (3) strong oscillatory

flow in the vicinity of the empennage. The figure shows the

geometrical configuration used in the computations and the

contours of surface pressure coefficients and engine plume

temperatures at SOFIA cruise-flight conditions. The cavity was

fitted with an internal shear-layer ramp and demonstrated

broadband acoustic behavior similar to the forward-mounted

design. However, a region of separated flow is apparent

immediately downstream from the cavity aperture. Optical

clarity of the infrared telescope was also investigated. There is

some degradation of clarity in the near-infrared regime if the

telescope is mounted behind the wings.

Significance

The use of CFD in the design of SOFIA has provided information

that is difficult to obtain experimentally. The generality of the

methodology has attracted industry interest for applications

ranging from the stealthy internal carriage of stores to the design

of airborne laser platforms.

Future Plans

The Boeing 747-SP horizontal and vertical stabilizer geometry

will be included in the computations during the next year.

Calculations will focus on a detailed investigation of the

separation region behind the cavity and fluctuating loads on

the aircraft empennage.

Contours of SOFIA surface pressure coefficients and normalized plume temperatures: magenta =

high, blue = low. Insets (top to bottom): instantaneous diffraction pattern; instantaneous streamlines

near the cavity; quiet cavity (forward mount); and rectangular cavity (forward mount).

18 Aeronautics/Basic Research
Finite-Element Euler Solver for a Distributed Memory Computer

Rainald Lohner, Principal Investigator

Co-investigator: Alexander Shostko

George Washington University

Research Objective

To develop a three-dimensional (3-D) unstructured-grid generator in the multiple-instruction/multiple-data (MIMD) context.

Approach

The advancing-front grid generation technique was used in

conjunction with unstructured grid load balancing and domain

partitioning methodologies with the host-nodes parallel model.

Moreover, distributed data structures were used for the parallel

3-D grid generation.

Accomplishment Description

The unstructured-grid generator prototype was developed and

ported to the MIMD context. The algorithm used a background

grid as the means to separate spatially different regions, enabling

the concurrent, parallel generation of elements in different

domains. After generating elements in parallel in the different

domains, the empty interdomain regions were gridded in

parallel. Several difficulties were overcome by extending the

two-dimensional grid generator to three dimensions. Load

balancing was improved by optimization of initial subdivision,

similar to heat conduction methodology. In 3-D, only a fraction

of the total mesh could reside in any one processor at a given

time. Thus, the grid is kept in the processors, where it was

generated. A numbering method for the generated nodes and

elements was developed and yielded a unique identification for

each node and element generated on all processors. Optimal

tolerance for the interfaces and predefined maximum effort to be

spent affect the quality of the generated grid. The total execution

time is still under investigation. New visualization tools for

debugging complex geometries and for postprocessing were

developed for the 3-D subdomain/interface grid generation.

Significance

For the first time, 3-D grid generation was ported to the MIMD

context.

Future Plans

Grid generation for complex 3-D geometries will be tested and

benchmarked. Front-end preprocessing of the background grid

will be developed for better load balancing in cases with coarse

background grids and background sources.

(a) Cb)

(c) (d)

Parallel grid generation. (a) Domain definition and background grid. (b) Subdomain grids after parallel generation.

(c) Interface and subdomain grids after parallel generation. (d) Assembled grid.

Aeronautics/Basic Research 19
Space Shuttle Flow Field

Fred W. Martin, Jr., Principal Investigator

Co-investigators: P. Buning, W. Chan, I. Chiu, R. Gomez, J. Slotnick, E. Ma, D. Pearce, T. Wey, and M. Kandula

NASA Johnson Space Center

Research Objective

To improve the numerical modeling of the Space Shuttle Launch

Vehicle (SSLV) ascent aerodynamic environment by improving

the geometric resolution of the vehicle.

Approach

The OVERFLOW-Chimera scheme is used to obtain flow fields

about the SSLV in the Mach number range 0.64.5, with

particular emphasis on the transonic range. ICEM-computational

fluid dynamics (CFD), a computer-aided design (CAD)-based

grid generation package, was used to produce the surface grid.

The resulting grid has, on average, a four inch resolution along

the surface. The hyperbolic volume grid generator HYPGEN was

used to create the volume grids.

Accomplishment Description

A 113-grid system comprised of approximately 16 million points

was generated. Hyperbolic volume grids were used to resolve

the boundary layer features. The hyperbolic volume grids

communicated with Cartesian grids, which captured flow

features between the bodies and communicated with a relatively

coarse farfield grid. A thin-layer Navier-Stokes solution of the

SSLV during ascent was initiated.

Significance

The ICEM-CFD model was the first complete CAD model of the

entire SSLV. The model included all surface protuberences larger

than four inches because of the geometric sensitivity observed in

previous solutions. The CAD model exceeds the geometric

fidelity of the wind tunnel model used to develop the original

aerodynamic data for the SSLV.

Future Plans

Flow-field calculations using the 113-grid system are under way

for a Space Transport System-50 flight case including the Space

Shuttle main engine and solid rocket booster plumes. Comparison with available wind tunnel data and flight measurements

will be used to verify the accuracy of the grid system and flow

solver. The required accuracy for this solution is to be within

5 percent of the wing's ultimate capability.

Publications

1. Pearce, D. G. et al.: Development of a Large-Scale Chimera

Grid System for the Space Shuttle Launch Vehicle. AIAA

Paper 93-0533, 1993.

2. Kandula, M.; and Buning, P.: Evaluation of Baldwin-Barth

Turbulence Model with Thin-Layer Navier-Stokes Computation of an Axisymmetric Afterbody-Exhaust Jet Flowfield.

AIAA Paper 93-0418, 1993.

3. Slotnick, J. P.; Kandula, M.; Buning, P. G.; and Martin,

F. W., Jr.: Numerical Simulation of the Space Shuttle Launch

Vehicle FIowfield with Real Gas Solid Rocket Motor Plume

Effects. AIAA Paper 93-0521, 1993.

Surface grids on the SSLV.

Surface grid topology of aft attach hardware.

20 Aeronautics/Basic Research
Direct Computation of Aerodynamic Sound Generation

Parviz Moin, Principal Investigator

Co-investigators: Sanjiva K. Lele, Tim Colonius, Brian E. Mitchell, and Jonathan B. Freund

NASA Ames Research Center/Stanford University

Research Objective

The prediction of the sound produced by turbulent flow requires

a detailed knowledge of acoustic source terms. Computation of

both the acoustic sources and far-field sound using the unsteady

Navier-Stokes equations allows direct validation of aeroacoustic

theories, and will help provide noise control strategies.

Approach

Accurate computation of far-field sound and near-field hydrodynamics requires the Navier-Stokes equations to be solved

using accurate numerical differentiation and time-marching

schemes, with nonreflecting boundary conditions. Sixth-order

compact finite-difference schemes were used to evaluate

spatial derivatives and time was advanced with a fourth-order

Runge-Kutta scheme. Nonreflecting boundary conditions

developed for the two-dimensional (2-D)linearized Euler

equations were modified for use with nonlinear Navier-Stokes

computations of open flow problems.

Accomplishment Description

The numerical scheme, which was previously validated on

several model problems, was used to compute the sound

generated by a 2-D spatially developing subsonic mixing layer.

The layer was forced at its most unstable frequency and its first

three subharmonics, causing a highly organized structure (nearly

periodic in time). Vortices were formed and underwent pairing

twice at fixed locations downstream from the inflow boundary.

The sound sources associated with the flow were computed.

The figure shows the vorticity contours in the near field, indicating the formation of vortices and the two pairings. Also plotted

in the figure are the contours of the dilatation outside the near

field, showing the acoustic waves that are generated by the flow.

The most intense sound is produced by the second pairing. This

computation required 200 hours on the Cray Y-MP with a core

memory of 10 megawords.

Future Plans

Computations of the sound sources in a three-dimensional round

jet are under way. Techniques for reducing computational cost

are being investigated, including the development of a hybrid

numerical scheme which solves a convected wave equation for

the far field and the Navier-Stokes equation in the near field and

matches the two solutions.

Significance

The computation of the sound generated by the 2-D mixing layer

represents the first time that direct numerical simulation has

been used to compute the sound generated by a turbulent shear

flow. The results will shed light on the nature of the sound

generation by flows with noncompact vorticity fields and will

allow a critical evaluation of the accuracy of aeroacoustic theory

in predicting the sound generated by turbulent shear flows.

Contours of near-field vorticity and far-field dilatation in the 2-D

mixing layer.

Aeronautics/Basic Research 21
Computational Fluid

Yehia Rizk, Principal Investigator

Co-investigators: Scott Murman and Ken Gee

NASA Ames Research Center

Research Objective

To use computational fluid dynamics (CFD) to study the effects

of engine inlet flow on the external flow field around the F-18

aircraft and to gain more understanding of the vortex-induced

tail buffet phenomenon, including the effect of installing a

leading edge extension (LEX) fence.

Dynamics of F-1 8 Flow Field

Approach

Simulation of the complete flow field around the F-18 was

achieved by coupling the external flow and the internal flow

through the engine inlet. A Chimera grid zonal scheme was used

to add necessary grids to the existing system to simulate the flow

through the inlet duct up to the compressor face. Boundary

conditions allowing the specification of the mass flow were

used. A weak coupling between the aerodynamics and structures

was used for the tail buffet study. A time-accurate Navier-Stokes

solver combined with Chimera was used for the aerodynamics.

The resulting airloads were used as a forcing function to

compute the structural response of the flexible vertical tail

using modal analysis. Chimera was also utilized to investigate

the effect of the LEX fence to reduce tail buffet.

because the airload frequency is close to the tail's first natural

frequency of bending. Computations also show that installing

the LEX fence introduces an unsteady vortex interaction which

results in reducing the airload frequency compared to the tail's

natural frequency. Coarse grid computations typically consumed

about 100 Cray Y-MP hours and about 12 rnegawords of central

memory.

Significance

Inclusion of the effects of engine inlet flow improved the

correlation between CFD results and flight data. CFD results

also provided more understanding of the vortex/tail interaction.

This increased understanding will lead to more effective means

of controlling or alleviating the buffet and preventing premature

tail structure fatigue for current and future high-performance

aircraft.

Future Plans

The effect of grid refinement on the flow-field structure will be

examined. Also, the computational method will be generalized

to account for a strong structural/aerodynamic coupling.

Accomplishment Description

The results indicate a strong coupling of the engine external and

inlet flows, especially at the maximum power engine setting.

Increasing the mass flow rate through the inlets caused the LEX

vortex breakdown location to move progressively downstream;

this movement is consistent with trends observed in flight tests.

Time accurate computations show that buffet frequency is

proportional to airspeed. At flight speed, tail buffet is observed

4O

30

=h

=,-

@1

,,== =" 20

_ lO

Icl

-.04

Inlet mass flow rate on LEX vortex; Moo = 0.243, Re = 11.0 x 106,

_z= 30.3 degrees. (a) Faired inlet; (b) flight idle; (c) maximum

power.

Sub-scale WT

• Full-scale WT

O Flight data

[] CFD

- • •

J'', ,

50 100 150

I I

200 250

Speed (m/sec)

- _ LEX fence off

/" \ -- -- -- LEX fence on

- ,,-, ,,\ ,

I I I I I

0 10 20 30 40 50

Nondimensional time

Effect of LEX fence on vortex-induced buffet frequency.

22 Aeronautics/Basic Research
Wings with Control Surfaces

David M. Schuster, Principal Investigator

Lockheed Engineering and Sciences Company

Research Objective

To develop a computational aerodynamics method to analyze

and design aeroelastic wings with multiple deflected control

surfaces.

Approach

Wing geometries with specified leading- and trailing-edge

control surfaces were modeled by allowing the computational

grid to shear between control surfaces and by assuming that

there was no gap along the hinge line. The models were

analyzed using an existing Euler/Navier-Stokes aeroelastic

analysis method (ENS3DAE). This method computed steady or

unsteady aerodynamic loads and the structural deflections that

resulted. These loads and deflections were used to determine

the aerodynamic and structural performance of the wings.

Accomplishment Description

This method was applied to the aeroelastic design of an

advanced fighter wing for efficient operation at steady, transonic

maneuver flight conditions without degrading the wing's

transonic cruise performance. A set of leading- and trailing-edge

control surface deflections in conjunction with an aeroelastic

twist distribution were defined to obtain optimal aerodynamic

performance at a specified maneuver flight condition. The

predicted aerodynamic loads, along with the structural deflections required to obtain the optimal aerodynamic performance

at this flight condition, were used to compute the bending and

torsional stiffness characteristics of the wing using an inverse

structural design procedure. From this computation, an unloaded

jig shape for the wing was extracted, which, under cruise

loading, deflected to the optimal cruise geometry of the wing.

Upon deflection of the control surfaces and adjustment of the

flight conditions to the maneuver design point, the wing

aeroelastically deflected to the appropriate shape for optimal

maneuver performance.

Significance

The interaction of controls, structures, and aerodynamics is an

important design consideration for modern commercial and

military aircraft. The method studied demonstrates an efficient

scheme for implementation of complex-flow aerodynamic

analysis methods into the design environment.

Future Plans

Parallel processing will be employed to couple the aeroelastic

method with a finite-element computational structural mechanics method to perform more detailed fluid-structure interaction

analyses and designs.

Surface grid for a wing with deflected control surfaces and aeroelastic pressure coefficient contours on a fighter wing design at

transonic maneuver flight conditions. Pressure coefficient ranges from -1.0 (blue) to 1.0 (magenta) with an increment of 0.1

between contours.

Aeronautics/Basic Research 23
Sonic Boom Predictions for High-Speed Civil Transport

Michael J. Siclari, Principal Investigator

Co-investigator: Kamran Fouladi

Grumman Corporate Research Center/Lockheed Engineering and Sciences Company

Research Objective

To develop a fast Euler solver capable of accurately predicting

sonic boom pressure signatures generated by complex aircraft

geometries.

Approach

The unsteady three-dimensional Euler equations were solved

explicitly within the context of a finite-volume multistage

Runge-Kutta integration scheme using local time stepping and

residual smoothing to accelerate convergence. An implicit

upwind marching scheme was utilized along with an axial

multiblock approach to facilitate the modeling of complex

geometries including multiple lifting surfaces, inlets, exhaust,

and engine nacelles.

Accomplishment Description

Grid spreading that occurs with conventional grid topologies

results in poor resolution of the flow field downstream of an

aircraft without the use of millions of grid points. The current

approach retains resolution of the flow field so that pressure

signatures can be predicted at one to two body lengths below

the aircraft. These near-field pressure signatures are then

extrapolated to the ground using atmospheric propagation

programs based on linear theory. The magnitude and shape

of sonic boom signatures generated by candidate low-boom

High-Speed Civil Transport designs can be evaluated using this

method.

Pressure propagation behind a supersonic vehicle.

Significance

This innovative grid topology reduced grid spreading and shock

smearing and resulted in accurate field computations several

body lengths downstream of an aircraft. The speed of this

computational fluid dynamics method for supersonic flows is

ideally suited for the design optimization environment.

Future Plans

New research will be carried out to couple this methodology

with a nonlinear design optimization program to design aircraft

shapes with optimum aerodynamic performance and minimum

sonic booms.

Pressure propagation below a supersonic vehicle.

24 Aeronautics/Basic Research
Navier-Stokes Equations for Jet Acoustics

Eli Turkel, Principal Investigator

Co-investigator: Ehtesham Hayder

ICOMP/NASA Lewis Research Center

Research Objective

To investigate the noise generated by and the stability properties

of a jet subject to perturbations. This requires coupling the nearfield solution, which is nonlinear, with a linear far-field solution.

Because of the large range of scales that are present one must

use higher order numerical methods.

Approach

The fully time-dependent Navier-Stokes equations were solved

with a high-order finite-difference method that was fourth-order

accurate in space and second-order accurate in time. The

algorithm was extended from one to several space dimensions

using a splitting method. The fluxes were extrapolated at the

boundaries to keep the formal order of accuracy. Various

boundary conditions were compared at the outflow boundary

to prevent waves from being reflected and contaminating the

interior solution. The goal was to simulate aerodynamic noise

associated with a supersonic jet flowing into a subsonic free

stream. The computations were done in two stages. First, a mean

state of the field with steady inflow boundary conditions was

computed, then the transient behavior of the flow was calculated

using a boundary condition that was a time-dependent perturbation of the initial boundary condition. The perturbation could be

either a sinusoidal perturbation or a perturbation based on the

eigenmodes of the linearized problem.

Accomplishment Description

Planar and axisymmetric jets were considered. For the planar

case, the jet Mach number was 1.5 and the Reynolds number

was 1.27 million. Solutions were obtained for steady-state

calculations and a time-dependent simulation. Results for the

axisymmetric jet simulations were similar to those of the planar

jet case. The original algorithm was extended to be fourth-order

accurate in space for the viscous and inviscid terms and was

compared with the solutions for a flat-plate calculation. Several

radiation boundary conditions were also considered. For a

nondimensional time of 200 and a 150 x 200 x 7 grid for a

domain 25 jet diameters long, 8 megawords of memory and

16 processing hours were required.

Significance

The use of a higher order accurate scheme is essential to obtain

accurate solutions with limited resources.

Future Plans

A major problem with three-dimensional Navier-Stokes equations is the treatment of a singular line. Future focus will be on

the correct numerical treatment of a singular line and its

application to jet acoustics.

x

1

Subsonic jet simulation with high resolution in a large domain. Mach = 0.8, angle of attack = 0.0 degrees, Reynolds number =

1.02 x 104. Contour levels: dark blue = 0.0 and magenta = 1.6.

Aeronautics/Basic Research 25
Block-Structu red Approach for H igh-Lift Appl ications

Veer N. Vatsa, Principal Investigator

Co-investigator: Mark D. Sanetrik

NASA Langley Research Center/Analytical Services and Materials, Inc.

Research Objective

To demonstrate the feasibility of using a block-structured grid

approach for computing viscous flows over high-lift configurations of practical interest.

Approach

A newly developed grid generation procedure with automatic

blocking and compact grid enrichment capability was coupled

to a cell-centered, finite-volume Navier-Stokes multiblock flow

solver.

Accomplishment Description

High-quality grids suitable for high-Reynolds number viscous

flows were generated for a three-element, high-lift airfoil

configuration. The resulting grids varied smoothly across block

boundaries, but maintained point-to-point continuity across

block interfaces. Grid points were clustered in selected regions

to resolve high gradients without propagating the grid points in

the far field. This clustering resulted in compact grid enrichment.

A multistage time-stepping algorithm with a multigrid acceleration technique was used to obtain steady-state solutions to the

Reynolds-averaged Navier-Stokes equations. A field-equation

type turbulence model (the Spalart-AIImaras model) was chosen

to accommodate multiple layers of wall-wake boundary-layer

interactions. The computed solutions compared well with

independent calculations and correlated well with experimentally measured pressure and skin-friction data.

Significance

Block-structured grids with point-to-point continuity across

block interfaces can be used to analyze complex aerodynamic

configurations. This methodology offers the advantages of

unstructured grids in its ability to accommodate complex

configurations and cluster grid points in regions of interest while

still maintaining the efficiency of structured-grid flow solvers.

Future Plans

This method will be applied to more complex configurations.

An effort is under way to compute flow over a transport aircraft

configuration and a High-Speed Civil Transport configuration.

Partial view of a 20-block grid for a 3-element airfoil.

Particle traces for a 3-element airfoil; Mach = 0.2, angle of

attack = 8.1 degrees, Reynolds number = 9.0 million.

26 Aeronautics/Basic Research


Flow Fields Including Chemical Kinetics

Ramesh K. Agarwal, Principal Investigator

Co-investigators: Jerry E. Deese, Paul Schulte, T. Mark Walter, Thomas P. Gielda, and Hugh Thornburg

McDonnell Douglas Research Laboratories

Research Objective

To develop computational methods capable of accurate and

efficient prediction of aerospace vehicle flow fields including

chemical kinetic and propulsion effects.

Approach

The Reynolds-averaged Navier-Stokes equations were solved

on body-conforming curvilinear grids using a finite-volume

Runge-Kutta time-stepping algorithm. Euler, thin-layer, slenderlayer, and full Navier-Stokes options were available. Complex

configurations were modeled with a multiblock zonal implementation. Algebraic and two-equation turbulence models

have been incorporated into the code. Nonequilibrium air,

hydrogen-oxygen, and hydrocarbon chemical models were

included. Ionization models were available for analysis of

electromagnetic wave transmission through the flow field.

Future Plans

Chemistry models for solid-rocket-motor plumes will be

incorporated into the code. Additional chemistry models will

be added and validated.

Publication

Gielda, T. P.; Walter, T. M.; and Agarwal, R. K.: Single Stage

Rocket Performance. AIAA Paper 92-1386, March 1992.

Accomplishment Description

Flow fields about a number of high-speed configurations were

computed. Single-stage-rocket technology configurations were

analyzed at angles of attack from 0 to 180 degrees. Several

missile and launch vehicle configurations with plumes were

analyzed to determine base drag. The figure shows the Mach

number distribution in the symmetry plane of a launch vehicle

with three boosters including the core and booster plumes.

High-Mach number calculations were run for reentry vehicles

requiring the nonequilibrium air chemistry model. Typical

calculations required 20-40 megawords of memory and

10 20 Cray-2 hours.

Mach number distribution in the symmetry plane of a launch

vehicle with three boosters.

Significance

High-speed, high-enthalpy flows are difficult to simulate

experimentally. Accurate simulation of these flows will give

designers an important computational data base on vehicle

performance that cannot be obtained experimentally.

q

Aeronautic_-Iypersonics 29
Three-Dimensional Compressible Turbulence Modeling

Jorge E. Bardina, Principal Investigator

Co-investigator: Thomas J. Coakley

MCAT Institute/NASA Ames Research Center

Research Objective

To verify and develop the baseline two-equation turbulence

models that account for the effects of compressibility in highspeed turbulent flows. The focus is on the identification of

models that are superior in their predictive capabilities of threedimensional (3-D) turbulent flows.

Approach

Turbulence models were developed based on insight gained

from experimental and theoretical research. A data base of highspeed turbulent flows was used in this investigation. Specific

topics for the data base were high-speed attached boundary

layers with pressure gradients, supersonic shear-layer mixing,

and shock wave/boundary layer interactions (SWBLI). A 3-D

compressible Favre-averaged Navier-Stokes code in a generalized coordinate system with two-equation turbulence models

was used to predict the experimental SWBLI data base. The

code was based on an implicit second-order total-variationdiminishing flux-difference splitting algorithm with an alternating

symmetric Gauss-Seidel plane relaxation method and a diagonally dominant approximate factorization scheme in each plane.

The characteristic-based boundary approximations were implicit.

Accomplishment Description

The computational method was validated against experimental

data and two-dimensional simulations of a flat-plate boundary

layer at Mach 5 and a compression corner at Mach 3. A set of

Mach 8 numerical simulations of a 3-D SWBLI experiment

generated with a single fin and an intersecting SWBLI experiment generated with two fins were computed. The turbulence

models investigated were the zero-equation Baldwin-Lomax

algebraic model, several forms of the k-c0 two-equation turbulence model, and two low-Reynolds number k-_ two-equation

turbulence models. The model corrections investigated were a

length-scale correction, a rapid-compression correction, and a

rotation correction.

Significance

Turbulence modeling is a critical technology needed to develop

the National Aero-Space Plane. A critical task is the ability of

the turbulence model to predict hypersonic viscous/inviscid

interactions. These computations provide benchmark results to

develop and implement two-equation turbulence models in the

design.

Future Plans

Two simulation experiments are planned: (1) the interaction of

two oblique shock waves with a thick turbulent boundary layer,

and (2) the interaction of shock waves generated with two

straight fins and a top wall with a thick turbulent boundary layer.

A major consideration is the development of improved compressibility corrections in the turbulence models and the identification

of models that have superior predictive capabilities.

Particle traces of Mach 8 intersecting shock waves and thick turbulent-boundary-layer interaction.

30 Aeronautics/Hypersonics
Waverider and Reference Model

Charles E. Cockrell, Jr., Principal Investigator

NASA Langley Research Center

Computational Study

Research Objective

To conduct a computational study of (1) a Mach 4.0 viscousoptimized waverider model, (2) a configuration designed for

shock attachment along the outer leading edge and high lift-todrag ratios at the design Mach number, and (3) a non-waverider

reference flat-top model. Performance characteristics of the

configurations will be compared to illustrate the benefits of

waverider flow-field properties.

Approach

Three-dimensional thin-layer Navier-Stokes (TLNS) and parabolized Navier-Stokes (PNS) solutions were obtained for the

waverider and reference model configurations using the General

Aerodynamic Simulation Program (GASP). Lift and drag values

were obtained by integrating computational fluid dynamics

(CFD) surface pressure predictions for each configuration.

Mach 4.0 waverider model (left) and reference flat-top model

(right).

Accomplishment Description

Solutions for the waverider and reference model configurations

were obtained at Mach 4.0 at 0-10 degrees angle of attack.

Solutions were obtained at Mach 2.5 and Mach 3.5 at 0 degrees

angle of attack for the waverider model only. The CFD surface

pressure predictions and the integrated lift and drag values for

the reference model configuration were lower than those for the

waverider model. The aerodynamic performance of the reference model configuration was slightly better than the waverider

model because it had a lower drag for the same amount of lift.

Flow-field solutions showed that the reference model exhibited

the same waverider shock attachment properties because the

planform shapes were identical. The planform shape of the

waverider created shock attachment at the leading edge resulting

in the aerodynamic performance advantage. Typical solution

times required 2-3 Cray-2 hours and 10-20 megawords of

memory.

Significance

The planform shape of the configuration is the property that

gives the waverider an aerodynamic performance advantage

over conventional hypersonic vehicles. This suggests that the

bottom surface may be altered without a significant decrease in

aerodynamic performance as long as the planform shape is

maintained. A configuration can provide slightly better aerodynamic performance, but it does not necessarily have better

propulsion airframe integration characteristics.

Static pressure values, nondimensionalized by free-stream

pressure, ranging from 0.75 (blue) to 2.0 (red), for the waverider

(top), reference model (bottom), at the centerline (left), and base

of each model (right). The solutions are at Mach 4.0 and

1 degree angle of attack for the waverider model and 2 degrees

angle of attack for the reference model.

Future Plans

A computational study will be conducted for a Mach 4.0

waverider-derived vehicle with realistic vehicle components

integrated.

Aeronaut ics/Hypersonics 31
Reentry Heating Effects on the Shuttle Orbiter

Peter A. Gnoffo, Principal Investigator

NASA Langley Research Center

Research Objective

To evaluate the capabilities and limitations of the Langley

Aerothermodynamics Upwind Relaxation Algorithm (LAURA)

program for calculating laminar convective heating over the

Space Shuttle Orbiter in a hypersonic flight regime with strong

equilibrium and nonequilibrum gas chemistry effects.

Approach

Flow field simulations at three trajectory points for Space

Transport System (STS)-2 between Mach 24 and Mach 12

(72 to 55 kilometers altitude) were obtained with LAURA and

calculated convective heating levels were compared to flight

thermocouple data. A multiblock solution strategy partitioned

the simulated flow field into manageable blocks that required

only a fraction of the computer resources (time and memory)

of a full-domain approach. Partitions were reassembled and a

full-domain solution was generated from a nearly converged

initial condition.

Accomplishment Description

Calculated laminar heating levels along 15 linear cuts on the

windside and leeside shuttle surfaces were in generally good

agreement with flight thermocouple data for laminar flow at all

three simulation points. Transitional and turbulent flow on the

deflected bodyflap below Mach 18 and on elevons below

Mach 13 were not modeled. A factor of ten change in surface

catalysis rates caused a 15-20 percent change in convective

heating rates across this flight regime; the application of a

radiative equilibrium wall boundary condition accentuated these

differences. Typical resource requirements were 100-250 hours

per case, 128 megawords of core memory, and 128 megawords

of solid-state storage device memory. The code achieved

concurrent central processing unit average values greater than

seven in the multitask queue of the Cray Y-MP.

Significance

Multiblock methodology and simulation capabilities were

established for future vehicle studies across the hypersonic flight

regime.

Future Plans

Additional grid convergence tests and multidimensional grid

adaption on leeside shuttle surfaces are planned.

Publications

1. Gnoffo, P. A.; Weilmuenster, K. J.; and Alter, S. J.:

A Multiblock Analysis for Shuttle Orbiter Re-Entry Heating

from Mach 24 to Mach 12. AIAA Paper 93-2813, July 1993.

2. Weilmuenster, K. J.; Gnoffo, P. A.; and Greene, F. A.:

Navier-Stokes Simulations of the Shuttle Orbiter Aerodynamic Characteristics with Emphasis on Pitch Trim and Body

Flap. AIAA Paper 93-2814, July 1993.

Heating levels over the shuttle surface at Mach 18.07 indicate

cooling effects of the expansion leading up to the deflected

bodyflap. A separation zone is evident in the form of a cool (dark

blue) crescent in front of the bodyflap hinge line. Peak heating

for this case occurs on the wing leading edge (dark red) in the

vicinity of the bow-shockwing-shock interaction region.

32 Aeronautics/Hypersonics
Particle Simulation of Hypersonic

Brian L. Haas, Principal Investigator

Co-investigators: Michael A. Fallavollita and William J. Feiereisen

EIoret Institute/Stanford University/NASA Ames Research Center

Research Objective

To develop and enhance a multiprocessor code for modeling

three-dimensional (3-D) reactive flows about realistic vehicles

during rarefied hypersonic aerobraking using the direct simulation Monte Carlo (DSMC) particle method. This code was

applied to realistic flow scenarios to assess code performance,

scaleup, and limits.

Rarefied Flow

Future Plans

Thermochemistry models for rotational, vibrational, and

chemical relaxation will be enhanced in the code. Loadbalancing algorithms will be improved to reduce interprocessor

communication costs and increase throughput. Variable cell

sizes within each block will be introduced to improve statistical

accuracy when simulating flows with large density variations.

Approach

DSMC methods modeled a rarefied flow as a large collection of

discrete particles traveling through space and interacting via

collisions. New collision-selection rules and other algorithmic

changes permitted efficient implementation on parallel-processing

computers. The flow field was divided into a network of small,

uniform, cubic cells. These cells were grouped into indivisible

blocks of a fixed size. The load-balancing scheme then dynamically assigned various numbers of blocks to each processor to

evenly distribute the computational work. Blocks tended to

align with the free-stream direction to reduce diffusion and

interprocessor communication. Sequential processing of cells

within each block reduced memory requirements.

Accomplishment Description

The code was modified for the Intel Gamma machine to run

with one processor serving both as a coworker and as the host

for the rest of the machine. This permitted ready porting to other

machines. As depicted in the first figure, steady 3-D simulations

were performed of a Mach 24 flow past a blunted cone during

aerobraking at 100 kilometers altitude in five-species reactive

air. As plotted in the second figure, nearly perfect scaleup was

observed up to 128 processors, although the costs of load

balancing operations became excessive as more processors

were used. By restricting the number of cells per block from

512 to 64, calculations for a given block remained within the

memory cache of the processor, thus reducing costs which

previously resulted from frequent accessing of main memory.

Significance

Because of limitations of the Navier-Stokes equations, continuum flow computation methods fail to accurately simulate

rarefied flows associated with atmospheric entry of space

vehicles or impingement of thruster exhaust upon satellite

components. DSMC methods are particularly well suited to such

scenarios, but they require excessive computational resources.

This multiprocessor code enables affordable simulation of flows

of engineering interest and exceeds the performance of an

efficient vectorized DSMC code (for the Cray Y-MP) when

running 64 or more processors. Good scalability permits

implementation on larger machines and thus simulation of

extremely large problems or unsteady flows.

Temperature profile for flow about a blunted cone. Temperatures

range from 1 to 125 times the free-stream value.

8 B

7 --

6--

3 -

2 -

D. 5

m4

Perfect scaleup /I-I

"D -- New balancing routine //

m --B--- Old balancing routine .z,_/,JII

L/11"

'z'//"_IS"

I I I I I i I I

16 32 48 64 80 96 112 128

Number of processors

Performance scaleup for the blunted cone simulations.

Comparisons between old and new load-balancing schemes.

Aeronautics/Hypersonics 33
Computation of Space Transport Vehicle Re-Entry Heating

William D. Henline, Principal Investigator

Co-investigators: Y.-K. Chen, G. E. Palmer, and S. M. White

NASA Ames Research Center/Eloret Institute

Research Objective

To apply three-dimensional (3-D) hypersonic flow computational

fluid dynamics (CFD) tools to the computation of re-entry surface

heating for hypersonic and space transportation vehicles.

Material surface boundary conditions were applied to provide

realistic estimates of aerothermal heat loads.

Approach

Hypersonic 3-D Navier-Stokes numerical solvers were used to

test vehicle aerothermodynamic surface heat transfer computational modeling. Material surface and thermal response boundarycondition subroutines were developed to couple with the CFD

flow code to represent the flow field plus heat shield aerothermodynamic response. The flow solver GASP2.0 was used to

calculate surface heating, temperatures, and ablation rates for

three calculations: shuttle tile gap heating, a coupled flow field

with heat transfer and surface catalysis on a high-angle blunt

cone, and a coupled flow field and ablation response for a .....<iiii

hypersonic nosetip. (a)

Accomplishment Description

For the shuttle tile gap heating calculation, the code was used to

model the overlying hypersonic boundary layer at Mach 5 as air

flowed over a tile gap. The downstream tile face was slightly

elevated to create a step-type obstruction for the impinging

flow. This calculation required a very fine numerical grid and

tight tolerences on local time steps. Runs took approximately - _

1 Cray Y-MP hour for simple isothermal wall boundary conditions. The first figure shows the pressure field and associated

velocity fields for the tile gap heating problem. GASP2.0 was

applied to the hypersonic heating to an ablating graphite nosetip

problem. This problem was difficult, and it required nearly

---- Z

50 Cray Y-MP hours. GASP2.0 was applied to the computation

of surface heating with catalysis on very-high-angle blunt cones•

Surface heating distributions compared well with experimental

data.

Significance

Valuable information on the sensitivity CFD algorithms to a wide

range of different complex vehicle surface boundary conditions

was gained. A specific coupling strategy was needed to efficiently obtain a solution for each problem. Only proper explicit

updating of the CFD flow-field variables using relaxation

techniques allowed any solution for coupled ablation problems.

(b)

-- Z

(a) Pressure and (b) velocity fields for Mach 5 air flow over a

transverse tile gap with a step.

Future Plans

The realism of material response boundary condition modeling

will be extended. More efficient ways to couple the flow codes

will be explored. Coupled problems will be extended to

compute flows over more realistic 3-D flight body geometries

and will be compared to experimental data and flight data.

34 Aeronautics/Hypersonics
Powered Hypersonic Air-Breathing Configuration Studies

Lawrence D. Huebner, Principal Investigator

Co-investigator: Kenneth E. Tatum

NASA Langley Research Center/Lockheed Engineering and Sciences Company

Research Objective

To perform computational analyses of hypersonic air-breathing

vehicles under simulated powered conditions. Three-dimensional

(3-D) effects of inlet fairings on the powered aftbody and wings

will be determined, as well as differences in plume characteristics at various points along a typical flight path.

Approach

The inlet-fairing effects and plume characterization studies

were performed on Langley's Test Technique Demonstrator

(TTD) using three different forebody/inlet representations and

a common aftbody with wings at four representative Mach

number/nozzle pressure ratio (NPR) combinations. Exhaust

flow simulation was initiated with internal nozzles designed

for each prescribed free-stream Mach number.

Accomplishment Description

Salient flow-field features about the TTD with a flow-ingested

inlet are shown in the first figure. No strong shocks were created

at the cowl leading edge, and pressurization of the aftbody lower

surface due to the powered exhaust simulation is evident. The

second figure compares the effects of inlet representation and

powered flow conditions by using the exhaust mass-fraction

contours at the body trailing edge. The faired-over inlet cases

show slightly greater plume expansion in the vertical direction,

but very comparable lateral expansion. There is a significant

effect of plume expansion for increasing Mach number/NPR.

At Mach 3.40, the plume remains attached only to the aftbody.

At higher Mach numbers, the plume expands vertically and

laterally, and at Mach 13.66 the entire lower surface of the

aftbody and wings is engulfed in the exhaust flow. Solutions

averaged about 1.5 Cray Y-MP hours and required about

7 megawords of memory.

Significance

Fairing over the inlet of a hypersonic air-breathing wind tunnel

model is a viable approach to obtaining powered aftbody effects

at supersonic and hypersonic speeds. Lower surface forebody

flow-field differences that propagate downstream are isolated

from the aftbody surface by the exhaust plume, resulting in

minimal aftbody force and moment differences. The expansion

of the exhaust plume over the vehicle lower surface and wings at

increased Mach numbers may alter the aerodynamic characteristics of this type of vehicle.

Flow-field Mach number contours (0 = blue to 5.9 = light pink),

and surface pressures (0 = blue to 3,000 = light pink) for the TTD

with a flow-ingested inlet, M_ = 5.92, NPR = 200.

[a) {b)

(c) [

TTD body trailing edge exhaust mass-fraction contours; 0.01 =

blue to 0.99 = light pink. (a) Flow-ingested inlet (left), faired-over

inlet (right), M_ = 3.40, NPR 30. (b) Flow-ingested inlet (left),

faired-over inlet (right), M_ = 5.92, NPR = 200. (c) Flow-ingested

inlet (left), faired-over inlet (right), M_ = 10. 14, NPR = 3,000.

(d) Flow-ingested inlet (left), faired-over inlet (right), M_ = 13.66,

NPR = 40,000.

Future Plans

Scramjet exhaust simulation techniques will be compared to

assess differences in cold-gas and hot-gas exhaust simulation,

including the modeling of the flow with scramjet combustion

products.

Aeronautics/Hypersonics 35
Computational Performance Enhancement of a Scramjet Inlet

Ajay Kumar, Principal Investigator

Co-investigator: Dal J. Singh

NASA Langley Research Center

Research Objective

To conduct a numerical performance enhancement study of a

sidewall compression scramjet inlet and to develop a design/

optimization procedure.

Publication

Korte, J. J.; Singh, D. J.; Kumar, A.; and Auslender, A. H.:

Numerical Study of the Performance of Swept Curved Compression Surface Scramjet Inlets. AIAA Paper 93-1837, June 1993.

Approach

A three-dimensional (3-D) Navier-Stokes code and a marching

Euler code were used to assess numerically the impact of curved

sidewalls on the performance of a baseline swept-sidewall

compression scramjet inlet at Mach 4 conditions. In addition, a

systematic 3-D design procedure was developed by optimizing

performance for a two-parameter inlet design utilizing a flowfield analysis based upon the Euler equations. A viscous solution

of the inviscidly optimized inlet was obtained and compared

with the baseline case. Effects of boundary-layer ingestion on the

overall flow features and performance were also investigated.

Accomplishment Description

Turbulent flow-field solutions were generated by solving

Reynolds-averaged Navier-Stokes equations to obtain inlet

performance parameters such as total-pressure recovery, mass

capture, and flow-field distortion. These parameters were used

to assess the performance enhancement by modifying the planarsidewall compression angle as a function of height, utilizing

sidewall curvature, and employing both the forward-swept and

backward-swept compression surfaces. This study indicated that

both the viscous and inviscid solutions yielded similar performance trends. The resulting optimized inlet geometry yielded

inlet performance improvements; however, inlet top-wall

separation still persisted. The code for an average run used

20 megawords of memory and took about 15 Cray Y-MP hours.

Significance

This numerical study showed that enhanced performance is

obtainable using curved-wall modifications to the standard

planar-sidewall inlet design.

Future Plans

A constrained inlet design/optimization procedure consistent

with multi-dimensional adverse pressure gradient separation

criterion will be developed.

Scramjet inlet flow-field Mach contours.

36 Aeronautics/Hypersonics
Integrated Hypersonic Vehicle Simulation

Scott L. Lawrence, Principal Investigator

Co-investigators: Gregory A. Molvik, Johnny R. Narayan, and Bradford C. Bennett

NASA Ames Research Center/MCAT Institute

Research Objective

To develop reliable tools for numerically predicting hypersonic

vehicle flow fields, including the effects of interactions between

the airframe and propulsion flow fields. Where possible, the

tools developed were used in-house to study the unique physics

associated with hypersonic flight.

Approach

In addition to extending current capabilities by implementing

models for physical effects not previously addressable (for

example, hydrocarbon-air combustion chemistry), existing

physical models were upgraded with more robust or more

accurate models where deficiencies were identified. The basic

suite of codes involved in this effort included the time-dependent

flow solvers, TUFF and SPARK, and the space-marching flow

solvers, STUFF and UPS.

Significance

The coupling of internal and external flow solvers provides the

capability to capture interactions between the airframe and

propulsion system that would be neglected if simulations were

carried out separately. The HWRV results play a role in the

design process through their use in calibrating the aerodynamics

and propulsion modules of a systems analysis code.

Future Plans

Hydrogen-burning scram jet vehicle simulation will be completed.

In conjunction with this work, other turbulence-chemistry

interaction models will be investigated and some design

problems associated with the scramjet combustor will be

addressed.

Accomplishment Description

A finite rate hydrogen-air chemistry model was adapted for the

purpose of analyzing the effect of turbulence-chemistry interactions in high-speed combustion. This model was initially

implemented in the SPARK code and applied to two-dimensional

and axisymmetric configurations involving hydrogen-air

combustion reactions. The model was also added to the TUFF

and STUFF codes and validated. Demonstration nose-to-tail

calculations were performed for a hydrocarbon-burning

waverider and were initiated for a hydrogen-burning National

Aero-Space Plane-type vehicle. The Hypersonic Waverider

Research Vehicle (HWRV) results (see figure) were obtained at

power-on design conditions/Mach 8.0, altitude = 92,500 ft). The

external flow, computed with a space-marching method,

required 5 megawords of main memory and approximately

6 hours of Cray Y-MP time. The time-marching scheme required

for the internal hydrocarbon scramjet flow employed a twoequation turbulence model, consumed approximately 103 hours

of Cray Y-MP time, and used roughly 20 megawords of main

memory. Tip-to-tail analysis of a generic, hydrogen-burning,

hypersonic vehicle (power-on) configuration is under way.

The forebody flow field was solved using TUFF and UPS.

Nonequilibrium air with finite-rate chemistry is considered at

a free-stream Mach number of 16. A scramjet engine consisting

of three intake modules with swept sidewall compression is

being considered for the power-on computations. The multizone

TUFF code is being used for the inlet computations. The engine

computation typically requires 12 Cray Y-MP minutes per

time step.

Nose-to-ta il ana lysis of a hypersonic wa verider research vehicle;

pressure contours, red = high, blue = low.

Aeronautics/Hypersonics 37
Efficiency Improvements of Boundary Layer Transition Predictions

C. C. Lee, Principal Investigator

Co-investigators: Shawn Hagmeier, Brad Hopping, and Charles Vaporean

McDonnell Douglas Corporation

Research Objective

To develop efficient predictions of three-dimensional (3-D)

boundary layer transition using linear stability analysis (LSA).

Approach

Three-dimensional LSA (eMalik) requires a high quality computational fluid dynamics (CFD) solution as input as well as

estimates for the most unstable frequency of the disturbance.

The LSA procedure is more efficient when an improved CFD grid

stretching function, frequency and disturbance growth onset

estimates, and automated frequency search procedures are used.

for LSA. There was up to a 40 percent decrease in the number of

radial grid points required for proper LSA results. A disturbance

frequency and onset of disturbance growth estimation were

incorporated and have automated the frequency search for the

most unstable disturbance.

Significance

Accurate prediction of transition is essential for hypersonic

vehicle design. Improvements are necessary to make the 3-D

LSA design process efficient. Frequency estimation is also useful

to determine instrumentation requirements for flight test vehicles.

Accomplishment Description

Various CFD stretching functions were studied to optimize the

CFD grid for LSA. The stretching function maintained adequate

resolution of the near-wall region for accurate heat transfer and

skin friction calculations and better resolved the boundary layer

Future Plans

Other candidate stretching functions will be tested for greater

efficiency. The process of 3-D LSA will be refined to allow its

use in the design of hypersonic vehicles.

(a) (b)

(a) Mach number distribution from 6 to 14 for a hypersonic flight test vehicle at a free-stream Mach number of 14. (b) Disturbance

frequency distribution from 80,000 to 350,000 hertz. The disturbance frequency at hypersonic speeds is estimated using the velocity

at the boundary layer edge and the boundary layer thickness.

38 Aeronautics/Hyperson ics
Flow Past a Dynamic Multiple Body System

K. Kurian Mani, Principal Investigator

Co-investigator: J. W. Haney

Rockwell International, Space Systems Division

Research Objective

To study multiple body dynamics in the hypersonic flow regime

with zonal decomposition.

Approach

In multibody dynamics analysis, the domain around bodies is

divided into multiple computational zones. The zone boundaries

and the grid nodes belonging to the adjacent zones may or may

not align. The grid was generated using GENGRID. Once the

basic grid structure was generated, nodal velocities were

assigned to the grid points. At first only the edge points were

calculated for the grid. From the edge points, the effect of all the

component bodies will be calculated. Next, each of the bounding surfaces was treated individually. For a selected surface, the

velocities of the points on the bounding curves were known

from the previous calculation. The interior points on the surfaces

were influenced only by the boundary points to maintain

continuity of induced velocities. A smooth blending of the grid

velocities was assured even if certain surfaces contained patches

with prescribed velocities. The solution provided an instantaneous picture of the flow field with the effect of body dynamics

on the flow field. This solution was achieved by accounting for

the nodal velocities of the surface grid points using the unified

solution algorithm code.

Accomplishment Description

The current study analyzed the flow field around a geometry

system consisting of the SR-71 aircraft with a "piggy backed"

waverider derivative at the aft end of the SR-71. The grid

generation was very complex. Once a grid was obtained, the

grid system was made flexible to accommodate the relative

motion. An algorithm was developed to compute grid velocities

across zonal boundaries. This algorithm was followed by a

solution process accounting for flux crossings across cell

surfaces in motion.

Significance

The study resulted in a method to analyze multibody dynamics

using zonal decomposition and grid stretching without a great

computational burden on the solvers.

Future Plans

The method will be extended to large relative motions.

Mach contours on the SR-71 aircraft/waverider geometry at

relative motion.

Aeronaut ics/Hypersonics 39
Performance

Charles R. McClinton, Principal Investigator

Co-investigator: David W. Riggins

NASA Langley Research Center

Potential of Full-Scale Injectors

Research Objective

To evaluate the performance potential of realistic scram jet

combustor injection strategies for flush-wall jets and intrusive

(ramp) injectors using state-of-the-art simulation capabilities. An

important goal is to gain understanding of the generation and

evolution of vorticity in supersonic combustors and its role in

the mixing process.

Approach

The three-dimensional Navier-Stokes family of SPARK reacting

codes was used to model the complex flow physics in scramjet

combustors for representative flight conditions and full-scale

geometries corresponding to a high flight Mach number

(Mach 14). Parametric studies of flush-wall and intrusive injector

geometries were undertaken using full finite-rate H2-O 2 chemistry. Performance potentials of the various strategies were

obtained using an energy-based method. Detailed studies of

vorticity generation and evolution were performed using

subscale injector ramps.

Publications

1. Vitt, P. H.; Riggins, D. W.; McClinton, C. R.; and Rogers,

C. R.: Validation and Application of Numerical Modeling to

Supersonic Mixing and Reacting Flows. AIAA Paper 93-0606,

Jan. 1993.

2. Riggins, D. W.; and Vitt, P. H.: Vortex Generation and

Mixing in Three-Dimensional Supersonic Combustors. AIAA

Paper 93-2144, June 1993.

Accomplishment Description

Flow fields for three full-scale candidate scram jet combustors

were obtained and analyzed for mixing, combustion, and thrust

potential. Details of the flow physics have been obtained and

allowed in-depth interpretation of vortex mixing and combustion

and comprehensive examination of flow losses. Stokes circulation law and moment-of-momentum checks provided confidence in modeling consistency for vorticity and vortex structures

for a series of subscale injector ramps. A correlation between

mixing and axial large-scale vorticity was developed.

Pressure contours over full-scale ramps in a scram jet combustor.

Significance

This study of full-scale flight-scaled injector ramps and jets in

supersonic reacting flows allowed the detailed comparative

study of flow phenomena that contribute to mixing, burning,

and thrust production for hypersonic air-breathing vehicles. In

addition, the understanding of injector-generated large-scale

vorticity and its contribution to fuel-air mixing was increased by

the parametric investigation of subscale ramps in high- and Iowenthalpy flows.

Reaction-generated water mass fraction contours 18 cm downstream of the ramp base (ramp is approximately 36 cm long).

Future Plans

Computational studies of scramjet engines with various injector

strategies will continue, including the specific modeling of a fullscale combustor/nozzle experiment for flight Mach 8 conditions.

40 Aeronaut ics/Hypersonics
Hypersonic Computational Studies

Grant Palmer, Principal Investigator

Co-investigators: S. Sharma, C. Park, S. M. Ruffin, T. Gokcen, D. Babikian, S. Tokarcik-Polsky, and E. Venkatapathy

NASA Ames Research Center/Eloret Institute

Research Objective

To investigate a wide spectrum of phenomena relating to the

computation of hypersonic, chemically reacting flows, including

coupled radiation, vibrational relaxation, and the application of

reacting gas flow solvers to space vehicle configurations.

Approach

Radiative transfer was coupled to an axisymmetric nonequilibrium flow solver for a nongray emitting and absorbing gas. The

resulting set of integro-differential equations were solved using

a modified Gauss-Seidel relaxation technique. In another study,

a new vibrational relaxation model that accurately described

relaxation in expanding base flows was developed using

modified Schwartz, Slawsky, and Herzfeld vibration-translation

transition rates and by considering both long-range and shortrange intermolecular forces when computing vibration-vibration

exchange rates. The performance of a drag device attached to a

lunar return aerobrake was assessed using an effective gamma to

simulate real-gas effects.

Accomplishment Description

The calculated spectra from the coupled radiation code qualitatively agree with those from both shock tube and Fire II flight

experiment data. The new vibrational relaxation model showed

substantial improvement over the standard Landau-Teller

model when applied to expanding flows. The new model

closely reproduced experimental data that showed an order

of magnitude greater energy exchange rate than predicted by

Landau-Teller. It was shown that by varying the drag-device

panel length and deflection angle the drag coefficient of the

lunar return aerobrake could be increased by a factor of 8.4.

Resource requirements for a drag-device computation were

1 5 megawords of memory and 10 Cray Y-MP hours.

1000 --

"0

2

E

0

f-

._o

E

UJ

10

.1

.01

.001

.2

-- Ames experiment

• AVCO experiment

Computation

I I I

.4 .5 .6 .7 .8

Wavelength (t_m)

• =t ,J

•,.

I I I

.3

Emission at the peak radiation point.

Significance

Coupled radiation and vibrational relaxation in regions of

expansion are two areas that have not been successfully integrated into nonequilibrium flow solvers. This work represents a

substantial advancement to the state of the art in these disciplines. The drag-device calculations provide information for a

realistic vehicle configuration that may aid future design of these

spacecraft.

Future Plans

The results of this work will be incorporated into future theoretical and applied hypersonic reacting-gas studies.

Calculated surface oil flow and surface pressure.

Aeronautics/Hypersonics 41
Computational Studies of Nonequilibrium Flows

Ramadas K. Prabhu, Principal Investigator

Co-investigator: George C. Olsen

Lockheed Engineering and Sciences Company/NASA Langley Research Center

Research Objective

To develop computer codes for the solution of two-dimensional

and three-dimensional Navier-Stokes equations assuming the

flow to be in chemical and thermal nonequilibrium, and to use

these codes to determine aerodynamic and thermal loads on

some critical components of hypersonic flight vehicles.

transport terms was completed, and several example cases were

solved. The transport terms have been included in the code. This

code was also vectorized for computational speed, and it is

being used to determine type IV shock interference heating on

the cowl leading edge of a hypersonic vehicle. The remeshing

codes are being modified.

Approach

The Langley adaptive remeshing code and Navier-Stokes solver

employed a cell-centered upwind scheme using Roe's flux

difference splitting technique. This solver, together with the

adaptive unstructured mesh generator, mesh enrichment, and

mesh movement techniques, was very effective in capturing

detailed flow features in complex flow situations.

Accomplishment Description

The perfect gas code was modified to include a five-chemical

species and two-temperature air model. The code with no

Significance

Knowledge of the magnitude of extreme aerothermal loads is

a key element in the design of hypersonic flight vehicles. This

project determined the aerothermal loads on some critical

components and demonstrated the power of unstructured

meshes in capturing intricate flow features.

Future Plans

Several examples of type IV shock interference will be solved

to determine the heat flux amplification. These codes will be

extended to three-dimensional flows.

(a) (b)

(a) Translational and (b) vibrational temperature contours for Mach 15 flow (with shock interference) past a 0. l-in.-radius body.

42 Aeronautics/Hyperson ics
Validation of the Parabolized Stability Equation Method

C. D. Pruett, Principal Investigator

Co-investigator: C-L. Chang

Analytical Services and Materials, Inc./High Technology Corporation

Research Objective

To validate the nonlinear parabolized stability equation (PSE)

method for transitional high-speed boundary layer flows.

100 -- _

Approach

Forced "second-mode" instabilities in a Mach 6.8 boundary 10-1

layer flow along a sharp cone were investigated by two fundamentally different approaches: PSE and spatial direct numerical

simulation (DNS). In PSE, the governing equations were approxi- _ 10-2

mate, but the solution was computationally efficient because of _"

the use of marching methods. In spatial DNS, the compressible E 1°-3

Navier-Stokes equations were solved in three dimensions and

time by highly accurate compact-difference and spectral .E

10-4

numerical methods. Results were compared with particular

attention paid to the evolution of nonlinearly generated harmonics of the forced disturbance. 10-5

Accomplishment Description

Spatial DNS was a yardstick by which other tools were

measured. Spatial DNS was used to assess the fidelity of the

nonlinear PSE method. PSE treats both mean-flow nonparallelism

and nonlinear wave interactions. Previously, DNS and PSE were

studied for incompressible flows, but this study is believed to be

the first comparison for high-speed boundary layer flows. The

first figure compares PSE and DNS computations for forced

transitional hypersonic flow along a sharp cone. The PSE

computation is forced with a single symmetric pair of finiteamplitude oblique second-mode disturbances (identified as

"T11"). All other harmonics arise spontaneously from nonlinear

interactions. The PSE result (lines) is then compared to the

DNS result (symbols). Agreement between the methods is superb

for harmonic amplitudes and harmonic structures (second

figure). The DNS required nearly 200 Cray-2 hours and

25 megawords of storage; the PSE computation required less

than 2 Cray-2 hours.

Significance

The PSE method is a reliable and computationally efficient tool

for investigating convectively dominated instabilities in highspeed boundary layer flows. The method remains valid well into

the moderately nonlinear stage of laminar-turbulent transition.

Moreover, results confirm that "oblique-mode breakdown" is a

possible path to transition in high-speed compressible flows.

Future Plans

The complete laminar-turbulent transition process in high-speed

boundary layer flows will be investigated combining PSEwith

spatial DNS. PSE will be used to compute the nonlinear initial

stages of the transition process to derive a harmonically "rich"

inflow condition for the spatial DNS. DNS will be used to

compute the strongly nonlinear and breakdown stages of

laminar-turbulent transition.

10-6

.9

--T33 ........... T13

_-- T31 _-mT11

------T22 ....... T04

T20 _T02

ii. 

1.1 1.3 1.5 1.7 1.9 2.1 2.3

_'-_exxlO 3

Maximum amplitudes of fundamental disturbance and selected

harmonics as functions of root Reynolds number; lines = PSE,

symbols = DNS. Legend: integers identify harmonics with

respect to temporal frequency and spanwise wave number,

respectively.

0

10-0

10-1

10-2

10-3

10-4

10-5

10-6

10-7

10-8

10-9

0

O T20

-- @ T13

Z_ T31

/ I i I I I I I I

.5 1.0 1.5 2.0

Wall-normal coordinate

Amplitude envelope versus wall-normal coordinate of temperature components of fundamental disturbance and selected

harmonics at %,l'Rex = 1,742.

Aeronautics/Hypersonics 43
Jet Interaction Aero-Optic Effects on

R. P. Roger, Principal Investigator

Co-investigator: S. C. Chan

Teledyne Brown Engineering

Research Objective

To obtain flow-field predictions for a supersonic jet exiting into a

hypersonic stream from a conical surface. Geometric configurations correspond to advanced concepts for defense interceptors

on which jet thrusters are used for maneuver control. The second

objective is to isolate effects that could adversely affect the

performance of an on-board infrared seeker. Effects due to large

angle of attack are of particular interest.

Approach

A three-dimensional full Navier-Stokes flow solver employing

an upwind algorithm was used to predict the complex jet

interaction flow field for the configuration. Computational fluid

dynamics results were generated for wind tunnel test conditions

to be compared to test data for the large recirculation region on

the missile surface upstream of the jet. Subsequently projected

tactical flight conditions were simulated.

Accomplishment Description

Computed results for the slot-cooled infrared seeker window in

a triconic forebody compared well with surface pressure and

temperature distribution data. Computed results for the large

center-of-gravity thruster firing at 0.0 and 16.0 degrees angle of

attack also compared well to force and moment data and to

Schlieren photographs. The 16-degree angle of attack case

required over 3 million grid points and was performed in

3 computational blocks. It required 200 Cray Y-MP hours and

52 megawords of memory for the largest block.

Significance

All proposed designs for advanced endoatmospheric interceptor

concept developments employ infrared seekers and center-ofgravity thrusters. Use of center-of-gravity lateral control during

"end game" is preferred because of the reduced sensitivity to

target miss distance. For maneuvering target intercepts large leeside angles of attack are necessary. In this configuration, the jet

upstream recirculation region could extend onto or over the

seeker window. This project investigated the fundamental flowfield structure for this phenomenon.

Hyperson ic Interceptors

Future Plans

This project has been completed. Tactical flight environment

predictions will be performed at a secure government computer

facility.

Publication

Roger, R. P.; and Chan, S. C.: CFD Study of the FIowfield Due

to a Supersonic Jet Exiting into a Hypersonic Stream From a

Conical Surface. AIAA Paper 93-2926, 1993.

Complex flow features in several cross planes for the 16-degree

angle of attack case.

44 Aeronautics/Hypersonics
Fuel Plume Imaging/CFD Comparisons

R. Clayton Rogers, Principal Investigator

Co-investigators: Robert D. Bittner and David W. Riggins

NASA Langley Research Center

Research Objective

To assess the fuel (hydrogen) distribution in a high-enthalpy

(Mach 13.5) pulse facility using computational fluid dynamics

(CFD) and Mie scattering of silicon dioxide particles. Both

methodologies were analyzed on a quantitative basis to compare

integrated fuel mixing. Because the amount of thrust provided by

an engine is directly related to fuel mixing, the ability to quantify

the degree of mixing with minimum uncertainty is imperative.

Approach

The SPARK codes were used to model this high-enthalpy flow

field. The codes were used in a blocking algorithm, where the

Navier-Stokes code was used only where the flow physics

dictated. The calculations were for mixing only. A laser with a

long pulse length illuminated silicon dioxide particles in the

injected fuel through Mie scattering. The hydrogen jet was

seeded with the silicon dioxide particles by preburning a

silane-hydrogen mixture and adding enough oxygen to consume

all the silane. The pre-burner outflow was modeled as a mixture

of hydrogen and water. The CFD hydrogen-number density flux

was then compared with the experimental Mie image intensity.

Accomplishment Description

A comparison of computational and experimental results three

inches downstream from a pair of 10 degree swept ramp

injectors is shown in the figure. The left side of the figure shows

the CFD planar hydrogen-number density flux normalized by the

total flux in that plane. The right side of the figure shows the

experimental planar laser intensity normalized by the intensity

integral in that plane. The CFD has slightly greater penetration,

less spreading, and a larger jet core. The CFD jet lifted off the

wall slightly more than in the experiment, but overall the

agreement is very good. A procedure developed to deduce the

integral mixing efficiency from an experimental image was

applied to the CFD-generated image. A value of 0.46 was

obtained, which agreed well with the CFD result of 0.44.

Applying the procedure to the data image resulted in efficiencies

between 0.48 and 0.53. This close agreement (_+10 percent) with

the CFD results was considered very good.

Significance

This was the first application of Mie imaging of a fuel plume

with silicon dioxide particles. Eventually, CFD will be used

to (1) predict how well fuel mixes in high-enthalpy flows,

(2) dissect the details of the flow physics, and (3) determine what

factors influence mixing. Long-pulse-length laser image correlation with Reynolds-averaged CFD is encouraging.

Future Plans

Recent tests of different geometries and crossflow plane locations to increase the high-enthalpy mixing will be analyzed. A

study of the effects of two-phase flow on the results has also

been proposed.

Comparison of CFD solution (left) and Mie imaging of fuel plume (right); red = high levels, blue = low levels.

Aeronautics/Hypersonics 45
Adaptive Unstructured Hypersonic Multigrid Solver

Rajiv Thareja, Principal Investigator

Co-investigators: Kenneth Morgan, Jaime Peraire, and Joaquim Peiro

Lockheed Engineering and Sciences Company/University of Wales/Massachusetts Institute of Technology/Imperial College

Research Objective

To further develop the FELISA system for the modeling of

compressible flows using unstructured tetrahedral meshes.

Approach

The algorithm used a centered scheme along with an explicitly

added artificial viscosity of matrix form. A multistage timestepping scheme was used in combination with an unstructured

multigrid algorithm in order to advance the solution to steady

state. The system can be used to model flows from subsonic to

hypersonic regimes. The figures show the computation of a

Mach 0.5 flow over a model of the Grumman X-29 at an angle

of attack of 5 degrees. A sequence of three meshes with

744,551,245,422 and 103,379 tetrahedral elements was

used in the computation. The fine mesh, two auxiliary meshes

used for the multigrid scheme, and the surface density contours

are shown.

Accomplishment Description

An unstructured mesh system was demonstrated as a powerful

method for the computation of three-dimensional Euler aerodynamic flows. The scheme used a side-based data structure with

minimal storage requirements. The practical usefulness of the

method was dramatically improved by incorporating a multigrid

acceleration procedure for the analysis of steady-state problems.

A typical multigrid run required about 100-150 cycles to

achieve convergence in about 3-4 Cray Y-MP hours using

about 8 megawords of memory.

Significance

The algorithm can be used to model Euler flows over complex

three-dimensional bodies using a series of unrelated unstructured

meshes and multigrids to achieve rapidly converged solutions for

flows across a wide range of flow speeds. The system (FELISA 2.0)

is available for testing.

Future Plans

The scheme will be extended for highly stretched meshes to

solve Navier-Stokes equations from subsonic through hypersomc

flows.

(a) (b)

(c) (d)

Grumman X-29 (a) surface mesh, (b) auxiliary mesh I, (c) auxiliary mesh 2, (d) density contours.

46 Aerona utics/Hypersonics
Aerothermodynamic Benchmark for Candidate AMLS Vehicle

Richard A. Thompson, Principal Investigator

NASA Langley Research Center

Research Objective

To provide high-fidelity aerothermodynamic solutions over the

forebody of a candidate Advanced Manned Launch System

(AMLS) vehicle during reentry for validation of preliminary

design efforts and to determine a starting point for follow-on

analyses of the vehicle aft-end.

Approach

Solution of the time-dependent, thin-layer Navier-Stokes

equations was achieved using the Langley Aerothermodynamic

Upwind Relaxation Algorithm (LAURA). Models for chemically

reacting flow were included to assess both equilibrium and

nonequilibrium air effects. Wall temperatures were computed

simultaneously with the flow-field solution using a radiationequilibrium wall boundary condition. Calculations were

performed for the vehicle near peak-heating flight conditions

where design of the thermal protection system (TPS) is critical.

Accomplishment Description

Flow-field and surface heating predictions, including the effects

of real-gas chemistry, were obtained for the vehicle forebody

at a moderately high altitude (78 km) and angle of attack

(32 degrees). Radiation-equilibrium wall temperatures were

computed at these conditions assuming both equilibrium-air

and finite-rate chemistry (seven species) in the shock layer. In

the latter case, the vehicle surface was considered to be fully

noncatalytic. The point-implicit solution strategy in LAURA

complemented the macrotasking capabilities of the Cray Y-MP

and Cray C-90. Sustained concurrent central processing unit

usage was measured up to 15.07 out of 16 processors on the

Cray C-90 using this multitasking approach. Typical solutions

required up to 95 megawords of core memory with supplemental solid-state storage of 50 megawords. Complete solutions

required on the order of 25 processing hours.

101

"G

_== 100

--z

×

0 10 -1

10-2

0

O Engineering Code_ Equilibrium air

LAURA J

m m m LAURA } Nonequilibrium

Windward

__-- Leeward

- _" _ .,,... .,..,......,."'" " _

I I I I I

20 40 60 80 100

Axial distance (ft)

Comparison of computed heat-transfer distributions along the

windward and leeward centerlines of the candidate AMLS

forebody.

Significance

Surface pressures and heat-transfer rates were computed over

the vehicle forebody for validating approximate engineering

techniques used in preliminary design efforts. Wall temperatures

computed at the peak-heating condition showed the current

design to be within bounds of the thermal protection system

limits. No regions of over-critical heating occurred on the

forebody.

Future Plans

Aerothermodynamic analysis of the vehicle aft-end, including

wings and tip-fin controllers, will be done. Investigation and

validation of the approximate engineering design techniques

will continue.

Contours of computed radiation-equilibrium wall temperatures

for equilibrium air chemistry on the forebody. Highest wall

temperature (1,650 K) is shown in red and the lowest (_700 K) is

in cyan.

Aeronautics/Hypersonics 47
Navier-Stokes Simulations of Orbiter Aerodynamic Characteristics

K. James Weilmuenster, Principal Investigator

Co-investigators: William L. Kleb and Francis A. Greene

NASA Langley Research Center

Research Objective

To develop and validate surface modeling, grid generating, and

flow solving software for the aerodynamic analysis of winged

entry vehicles.

Approach

Vehicle surface definition and control surface orientation were

implemented through ICEM software. GRIDGEN and 3DMAGGS

(three-dimensional multiblock advanced grid generation system)

codes were used for grid generation. Fluid dynamic simulations

were implemented through the Langley Aerothermodynamic

Upwind Relaxation Algorithm (LAURA) code, which is appropriate for solving viscous, hypersonic flows in thermochemical

equilibrium and nonequilibrium. Shuttle Orbiter wind tunnel

and flight data were used as a validation data base.

Accomplishment Description

Flow-field solutions were obtained for the complete Orbiter

configuration at flight conditions and for a simplified Orbiter

shape at tunnel and flight conditions. Surface properties were

integrated to determine the longitudinal aerodynamic characteristics of these configurations. A parametric study using the

simplified Orbiter configuration showed increases in flight

bodyflap trim angle over those required at tunnel conditions

when compared with the wind tunnel based, pre-STS-1 aerodynamic data book. This study also showed that the increase in

bodyflap trim angle was caused by an increased nose-up basic

body pitching moment caused by real-gas effects. Computed

aerodynamics for the complete Orbiter configuration were

within 5 percent of the flight values and the bodyflap trim angle

was predicted to within 10 percent of the flight value.

Significance

Ground-based facilities cannot simulate the aerothermal

environment encountered by entry vehicles at hypersonic

speeds. This study shows that the computational tools are in

place to define accurately the longitudinal aerodynamic

characteristics of such vehicles in this flight regime.

Future Plans

The technology developed in this study will be directly applied

to the analysis of the Advanced Manned Launch System, HL-20,

and other candidate winged entry vehicles.

2O

15

E 10

5

-5

35

m

0 Computed flight

[] Computed tunnel

40 45

O_(deg)

Comparison of computed Orbiter bodyflap trim requirements for

tunnel and flight conditions as a function of angle of attack.

48 Aeronautics/Hypersonics
Modeling for High-Speed Transitional

Thomas A. Zang, Principal Investigator

Co-investigator: Nabil M. EI-Hady

NASA Langley Research Center/Analytical Services and Materials, Inc.

Research Objective

To apply the state-of-the-art dynamic modeling of small-grid

scales to a transitional boundary layer at high speed. A central

issue in large-eddy simulation (LES) is model development for

the small subgrid scales. The model allows the transfer of the

right amount of energy from the large to the subgrid scales, or

vice versa, near the wall. In previous studies, the ad hoc manner

in which the model constants or the model modifications have

been treated was not satisfactory.

Approach

The subgrid scales were modeled dynamically in a temporal LES

of transitional boundary-layer flow along a cylinder at Mach 4.5.

The coefficients of the dynamic eddy-viscosity model were

determined from the local information of the smallest resolved

field. This was accomplished by filtering the equations of motion

with a test filter having a width double that of the grid filter. The

difference between the two resulting small-grid scale fields,

which are related by the Germano identity, is the contribution

by the scales of motion between the test and the grid scales.

Both the formulations of Germano et al. and Lilly (least-square

technique) were used for the dynamic modeling to compare and

assess their results. The results of the LES were compared with

accurate results from direct numerical simulations (DNS).

Boundary Layers

Significance

Simulation of temporal forced transition through laminar

breakdown and beyond can be accomplished accurately and

cheaply at high speeds. Location of transition onset (rise of skin

friction), the length of the transition zone, and peak skin friction

can be predicted accurately using dynamic LES modeling.

Future Plans

The LES method will be thoroughly tested with other state-of-theart small-grid scale models and will then be extended to spatial

transition.

(a)

Accomplishment Description

The dynamic eddy-viscosity subgrid-scale model has been

successfully applied to a high-speed transitional boundary layer.

The coefficients of the dynamic model were qualitatively

evaluated at various stages of the laminar turbulent transition

process. The model gave the proper asymptotic behavior of

modeled quantities near the wall and in the free stream, with no

dissipative character in the linear and early nonlinear regions. A

remarkable agreement exists between LES calculations and the

accurate DNS results concerning the resolved Reynolds stresses,

heat flux, and the time evolution of the skin friction. LES of the

transitional flow along a cylinder at Mach 4.5 is achieved with

one-sixth of the grid resolution used for DNS. A typical LES case

for laminar breakdown takes about 20-25 CPU hours with

central memory requirements from 2 to 8 megawords compared

to 125 CPU hours and 30 megawords for DNS.

(b)

Spanwise vorticity flow structure in the middle of the skin

friction rise from (a) DNS (96 x 48 × 144) and (b) LES

(36 × 24 × 96). Mach = 4.5, t = 54.

Aeronautics/Hypersonics 49
DSMC Simulation of High-Altitude Plume Interaction

S. H. Konrad Zhu, Principal Investigator

Co-investigator: Leonardo Dagum

Rockwell International, Rocketdyne Division/NASA Ames Research Center

Research Objective

To characterize the flow field from two simultaneously firing

reaction control system (RCS) engines on the Space Shuttle

Orbiter during proximity operation about Space Station Freedom.

Approach

The direct simulation Monte Carlo (DSMC) method was applied

to analyze the problem. PSiCM, a three-dimensional (3-D)

particle simulation for rarefied flows, was used to obtain the

flow-field solution.

Accomplishment Description

Solutions for the multiple RCS plume interactions were computed and demonstrated a complex, 3-D structure for the flow

field from two RCS plumes separated by distances of 7.8 meters

and 18.5 meters. The computational domain covered a downstream distance about the same as the separation distance. The

solutions provide essential information about the near-field flow

structure and physics for the concerned plume interaction.

PSiCM was compared to experimental results for the interaction

between two rarefied free jets. The computed flow fields show

excellent agreement.

Significance

This work was prompted by the Space Station design and the

possibility of excessive loads on the solar array caused by RCS

Density field from two RCS plume interactions separated by a

distance of 7.8 meters. Half the flow field is presented. There is a

20 degree computation cutoff angle measured from the plume

centerline. A shock is formed adjacent to the symmetry plane

between the two plumes.

plume impingement. These solutions indicate that there is still a

significant interaction that suggests the practice of predicting the

multiple-plume flow field as a linear superposition of each single

RCS plume is inaccurate in the present Space Station design.

Future Plans

A parametric study of the multiple RCS firings will be done to

provide a data base to establish an accurate engineering model

for the Space Station design. A multizonal capability and an

impingement capability will be added to PSiCM to extend

computational domain further downstream.

1.0

0

.9

.8

._ .7.6

.4

.3 O

z

.2

.1

I I I I I I

0 1 2 3 4 5 6 7 8 9 10

Y/D

Density profile along the symmetry plane centerline between

two interacting jets; circles = experimental data.

220

200

180

160

140

o

I,-

120

100

8O

60

0

I ] I I I I I I _'1"_

1 2 3 4 5 6 7 8 9 10

Y/D

Rotational temperature distribution along the symmetry plane

centerline between two interacting jets; circles = experimental

data.

50 Aeronautics/Hypersonics
Subsonic High-Lift Analysis

Kenneth M. Jones, Principal Investigator

Co-investigators: Kevin Kjerstad and Victor Lessard

NASA Langley Research Center/ViGYAN, Inc.

Research Objective

To predict the flow characteristics of several high-speed research

(HSR) configurations with attached-flow leading-edge flaps and

vortical-flow leading-edge flaps at subsonic takeoff and landing

conditions, and to compare the numerical and experimental

data. Numerical methods were assessed as analysis and design

tools for high-lift systems on HSR configurations.

Approach

Two approaches were used for numerical analysis. A threedimensional (3-D), thin-layer, Reynolds-time-averaged

Navier-Stokes code (CFL3D) was used to simulate an attachedflap HSR configuration on a two-block structured-volume grid.

The 3-D Euler code USM3D was used to simulate a vortex-flap

HSR configuration (including trailing-edge flaps) on an unstructured volume grid.

Accomplishment Description

The CFL3D code (with the Baldwin-Lomax turbulence model)

was used to predict the forces and moments of the HSR configuration with an attached-flow flap over an angle-of-attack range

from 0 to 20 degrees at Mach 0.22 and a Reynolds number of

1.4 million per foot. Surface pressures on the attached-flow flap

compared well with experimental data except at higher angles of

attack. The discrepancies at the higher angles of attack were due

to the inability numerically to predict transition and capture the

correct separation location for turbulence flow. The USM3D

code was used to model leading-edge vortex flaps coupled with

deflected trailing-edge flaps on a generic HSR configuration. At

the design condition for the vortex flaps, USM3D accurately

predicted the reattachment line of the primary vortex when

compared to experimental flow visualization. Forces, moments,

and surface pressures compared welt with the experimental data.

Significance

Accurately modeling and predicting the effects of leading- and

trailing-edge flaps of a High-Speed Civil Transport design are

critical in meeting high-lift requirements at takeoff and landing.

Future Plans

Viscous effects for split leading- and trailing-edge flaps will be

included using specialized domain decomposition techniques.

Pressure coefficient (Cp) contours and surface traces on an HSR

configuration with attached-flow flaps; Mach = 0.22, angle of

attack = 8 degrees, Reynolds number = 1.4 million per foot.

Pressure coefficient (Cp) contours on an HSR configuration with

vortex flaps; Mach = 0.22, angle of attack = 12 degrees.

Aeronautics/Su bsonics 51
Analysis and Design of High-Lift Systems

Anutosh Moitra, Principal Investigator

Co-investigator: Wendy B. Lessard

High Technology Corporation/NASA Langley Research Center

Research Objective

To simulate the viscous subsonic flow about a representative

High-Speed Civil Transport (HSCT) configuration with and

without deployment of control surfaces. The concept of optimally deflecting the leading- and trailing-edge flaps to achieve

attached flow was studied to improve the high-lift system

performance of an HSCT aircraft.

Approach

A vortex-lattice technique (WINGDES and AERO2S) based on

linear, attached-flow theory was used to determine the optimum

flap deflection to provide maximum performance for specific

flight conditions. The flow was simulated using CFL3D, which

solved the three-dimensional, compressible, Reynolds-averaged

Navier-Stokes equations.

Accomplishment Description

Numerical simulations for flow about a representative HSCT

configuration with and without leading-edge flaps deflected

were performed at Mach 0.08 and a Reynolds number of

640,000 at various angles of attack. As shown in the figure (left

side), a reduced vortical flow field (approaching attached flow)

results when the leading-edge flaps are optimally deflected. The

computations correlated very well with the experimental

coefficients at all angles of attack, and the predicted streamline

patterns were similar to experimental oil flow results.

Significance

Accurate simulation and prediction of the vortex-dominated flow

about an HSCT aircraft is essential to the design and analysis of

high-lift systems.

Future Plans

An unstructured gridding technique will be applied to an HSCT

configuration with the leading- and trailing-edge flaps deflected.

Total pressure contours for an HSCT configuration without

leading-edge flaps deflected, angle of attack = 8.0 degrees (left),

and with leading-edge flaps deflected, angle of attack =

9.7 degrees (right).

52 Aeronaut ics/Subsonics
Aircraft Forebody Flow Control Technology

Lewis B. Schiff, Principal Investigator

Co-investigators: Ken Gee, Scott Murman, Gabriel Font, and Roxana Agosta

NASA Ames Research Center

Research Objective

To determine the effect of pneumatic forebody flow control on

the flow field about a full aircraft geometry and simple configurations flying at high angle of attack using computational fluid

dynamics (CFD).

Approach

The flow about an F-18 geometry with tangential slot blowing

was computed using a multizone thin-layer Navier-Stokes code.

The slot geometry was modeled in the computations. The jet

conditions were introduced as boundary conditions in the slot

grid. A computation using the isolated forebody was obtained for

wind tunnel test conditions, and computations using the full

aircraft geometry were obtained for flight conditions. Also,

computations were carried out for simple bodies incorporating

tangential slot blowing at high incidence.

Accomplishment Description

Computations were accomplished for the F-18 isolated forebody

and complete F-18 configuration at wind tunnel and flight-test

conditions. In the figure part (a), the computed surface flow

pattern with blowing active on the left side of the isolated F-18

fuselage is shown. The flow is highly asymmetric on the nose of

the aircraft. An analogous computation for the full F-18 configuration is shown in the figure part (b). Here, the asymmetric

breakdown of the wing leading-edge-extension (LEX) vortices is

clearly visible. The left LEX vortex breaks down in advance of

the right vortex. Asymmetric flows generate large yaw moments

on the aircraft, which have the potential of being exploited for

enhanced control of the aircraft in high-angle-of-attack maneuvers. Computations were used to optimize the configuration of

the slot location, as seen in the figure part (c). A comparison of

the computed surface pressure distributions with data measured

in the full-scale test is in the figure part (d). It shows excellent

agreement with the experimental ' measurements.

Significance

The computations show the ability of CFD methods to model the

complex high-angle-of-attack flow field. The computations,

carried out in advance of the wind tunnel test, guided the design

of the wind tunnel experiment. This illustrates how the coordinated use of computational and experimental methods can

provide increased understanding of complex high-incidence

flows.

Future Plans

Time-accurate, full aircraft computations at wind tunnel

conditions are under way to more accurately resolve the

interaction of the forebody vortex and the vertical tails.

Computations for simple bodies are being done to investigate

the basic physics underlying the development of the flow

asymmetry at high angles of attack.

(a)

(b)

(c)

-2

-1

Cp

0

1

(d)

Computation

9O 180 270 36O

Slot blowing for enhanced aircraft high-angle-of-attack agility. (a) CFD prediction; (b) computed

asymmetric flow with blowing; (c) wind tunnel test configuration," (d) CFD/wind tunnel comparison.

Aeronaut cs/Subsonics 53
Steady and Unsteady Flows at High Angles of Attack

David T. Yeh, Principal Investigator

Rockwell International, North American Aircraft Division

Research Objective

To extend and validate a computational capability to assist in

the design of advanced technology aircraft that are required to

operate in a high-angle-of-attack flow environment.

Approach

The unified solution algorithms (USA) code and the ENSAERO

code were used for high-angle-of-attack flow simulations. The

USA code solved the three-dimensional Reynolds-averaged

Navier-Stokes equations through an upwind, flux-difference,

finite-volume approach coupled with a high-resolution totalvariation-diminishing scheme. The ENSAERO code coupled the

Navier-Stokes solutions with modal structural equations of

motion for computing aeroelastic responses of flexible aerodynamic surfaces. The X-31 enhanced fighter maneuverability

demonstration aircraft was the configuration studied. This

configuration was chosen because of its inherently designed

high-angle-of-attack characteristics and capabilities, the existence of a wind tunnel data base, and the ongoing flight-test

program.

solution. The figure shows the rigid solution of a simplified

X-31 under the free-stream conditions of Mach = 0.3 and

angle of attack = 30 degrees. A typical computation required

25 megawords of memory and 30 Cray Y-MP hours.

Significance

This study paralleled the X-31 flight-test program and established

and extended computational fluid dynamics capability for

analysis of high-angle-of-attack flow phenomena. This study wilt

have significant impact on current and future advanced aircraft

design.

Future Plans

The study will be extended to include structural response in an

unsteady vortex dominated flow environment based on the

current rigid solution.

Accomplishment Description

The X-31 configuration was gridded and numerical solutions

using the USA code were analyzed for prestall and poststall flight

regimes and compared with wind tunnel and flight-test data.

A nonaligned block interface technique was utilized to model

multiple canard deflection angles where a cylindrical grid

around the canard can be rotated about the hinge line. The

effects of the movable canard on the aerodynamic control in the

post-stall regime were demonstrated and validated with wind

tunnel test data. Aeroelastic calculation of the structural response

was initiated for applications such as the vertical tail buffet

phenomena observed in high-angle-of-attack environments.

A simplified one-block grid was generated to focus on the

demonstration of the aeroelastic simulation of the X-31. The

aeroelastic computation was started from the rigid steady-state

Pressure contours for a simplified X-31 rigid model.

54 Aeronautics/Su bsonics
CFD Applications to HSCT Design and Analysis

Dharmanshu L. Antani, Principal Investigator

Co-investigators: Shreekant Agrawal, David L. Rodriguez, Ramesh K. Agarwal, and Raymond Hicks

McDonnell Aircraft Company/NASA Ames Research Center

Research Objective

Computational fluid dynamics (CFD) codes are calibrated and

validated for supersonic wing designs to improve the cruise

aerodynamic performance of a H igh-Speed Civil Transport

(HSCT) configuration. A three-dimensional (3-D) Navier-Stokes

code for the analysis of current HSCT configurations was applied

for a highly loaded condition. A 3-D Navier-Stokes code was

calibrated and validated to analyze high-lift concepts for HSCT

configurations and to provide design guidelines for improving

aerodynamic performance of the high-lift system.

Approach

A 3-D wing optimizer was used in conjunction with an advanced

CFD code to improve the aerodynamic performance of an HSCT

wing originally designed using linear theory. The new wing

design was performed with an Euler code in the presence of the

fuselage. The improved wing was tested in the NASA Ames 9- by

7-Foot Supersonic Wind Tunnel to validate the design optimization methodology. A Navier-Stokes code was used to analyze a

current HSCT configuration in a 2.5 G flight condition to predict

aerodynamic loads for future application to structural design. A

3-D Euler/Navier-Stokes code with available data was used to

analyze an HSCT high-lift configuration.

Accomplishment Description

The 3-D Euler-based wing-body optimizer was used to design

a new HSCT supersonic wing with a predicted improvement in

the supersonic cruise lift-to-drag ratio of about 5 percent

(compared to a linear-theory designed wing) based on flight

Reynolds number conditions. An analysis of the wing using a

Navier-Stokes solver predicted an improvement in performance.

Navier-Stokes solutions were obtained for a current HSCT

wing-body configuration in 2.5 G subsonic, transonic, and

supersonic flight conditions for prediction of aero loads. Results

compared well with wind tunnel data. Euler solutions were

obtained for a high-lift HSCT configuration and were compared

to experimental data. Several cases were analyzed in each area

of study. An average job run for the wing optimization and

aerodynamic loads prediction required 8 Cray Y-MP hours and

64 megawords of central memory.

Future Plans

Wing and fuselage design optimization methods will be improved.

Wing design methods in the presence of nacelles and diverters

will be developed and applied. Wing-body design optimization

methodologies will be validated against 1993 supersonic wind

tunnel tests.

Pressure contours on upper surface of baseline (left) and

modified (right) wings.

Significance

This method shows large potential for improving aerodynamic

performance of HSCT configurations originally designed using a

linear method. Accurately modeling nonlinear aerodynamics is

the key to this success. CFD-based design optimization methods

will minimize expensive wind tunnel tests. Once verified,

spanwise loading from CFD predictions will have a strong

impact on the low-weight structural design of the wing.

Aeronautics/Supersonics 55
Transonic Flow Around

Osama A. Kandil, Principal Investigator

Old Dominion University

a Delta Wing

Research Objective

To predict steady and unsteady, transonic, vortical flows around

delta wings in the moderate- to high-angle-of-attack range. A

complex shock-vortex system develops in the transonic-flow

regime, which is important for the performance of the new

generation of super-maneuver aircraft. Experimental measurements of transonic flows around a 65-degree cropped delta wing

show that the shock system appearing over the wing consists of a

ray shock wave beneath the leading-edge primary vortex and a

transverse, time-dependent shock wave (known as a terminating

shock).

Approach

The unsteady, compressible, fulI-Navier-Stokes equations

were solved with an implicit, upwind, flux-difference, finitevolume scheme (FTNS3D). The scheme was a modified fullNavier-Stokes solver of the CFL3D program.

Accomplishment Description

A 65-degree swept-back, sharp-edged, cropped delta wing of

zero thickness was considered for the computational solutions.

The cropping ratio was 0.15. An O-H grid of 1 x 258 x 584 in

the wraparound, normal, and axial directions was used. The

minimum grid size normal to the wing surface is 5 x 104 . The

figure shows the results for a wing at 20 degrees angle of attack,

Mach 0.85, and Reynolds number 3.23 x 106. A plane-ofgeometric-symmetry view of the computed surface pressure and

Mach contours along with particle traces of the leading-edge

vortex core is shown. The terminating shock and the supersonic

pocket ahead of the wing are clearly captured along with the

vortex breakdown of the leading-edge core. The computed

surface pressure is in good agreement with the existing experimental data, and a complete reconstruction of the flow field

behind the terminating shock was presented for the first time.

The effects of the angle of attack and Mach number were also

studied.

Significance

Prediction and study of transonic flows around delta wings

are important to increase the performance quality of the new

generation of super-maneuver aircraft. Unlike the low-speed

regime, vortex breakdown in the transonic flow regime occurs

at moderate angles of attack and is caused by shock waves.

Future Plans

Work is under way to study the transonic flow around delta

wings undergoing pitching, rolling, and combined oscillations

with high amplitudes and during the vortex-breaking mode.

Coupling the computational fluid dynamics with the wing

dynamics will also be considered.

Surface pressure, Mach contours, and particle traces on wing

and symmetry planes; Moo = 0.85, Reynolds number =

3,230,000, a = 20 degrees.

56 Aeronautics/Transonics

t
Simulation of Turbomachinery Flows

John J. Adamczyk, Principal Investigator

Co-investigators: Tim Beach, Mark Celestin, Kevin R. Kirtley, Rick Mulac, Wai-Ming To, Jeff Yokota, and Aamir Shabbir

NASA Lewis Research Center/Sverdrup Technology, Inc./ICOMP

Research Objective

To develop a flow model to analyze the flow field within

multistage/single-stage turbomachinery.

Approach

In a parallel effort, an unsteady code based on the computational tool V-STAGE was developed to investigate the effect of

unsteadiness on turbomachines. V-STAGE is a finite-volume

cell-centered scheme that solves the average-passage equation

system. These simulations will be used with experimental data

to develop better average-passage equation system models.

Accomplishment Description

The effect of tip clearance on stall in a high-speed fan and the

blade redesign on a one-and-a-half-stage turbine were studied.

Simulations of two types of bypass fans, a compact two-stage

high-speed compressor, and an unsteady simulation of a highspeed single-stage fan were completed. A four-stage high-speed

compressor was analyzed. Typical run times were in excess of

20 Cray Y-MP hours per completed case and required up to

20 megawords of memory.

Significance

The design of advanced turbomachinery components requires

the development of a three-dimensional model that can predict

the complex flows associated with these devices. A model was

developed to answer many of the issues associated with the flow

field within single-stage and multistage turbomachinery.

Density distribution of an advanced high-speed compressor

rotor.

Future Plans

Refinements to V-STAGE will include higher order turbulence

models and closure models for the deterministic stresses.

V-STAGE-unsteady is being used to simulate unsteady flow

through a blade row.

PAGE_ -_ -INTEi,Ii-LOi'iALL'_' _LANK

Aeronautics/Propulsion 59
Investigation of Scalability of Parallel Turbomachinery Codes

Richard A. Blech, Principal Investigator

Co-investigators: Edward J. Milner, Scott E. Townsend, and Angela Quealy

NASA Lewis Research Center

Research Objectives

To develop a parallel, portable version of the average-passage

turbomachinery code and investigate the scalability of this code

on several parallel architectures.

Approach

A distributed-memory, message-passing programming model

was used. Portability was achieved through use of the Application Portable Parallel Library (APPL). Initial demonstrations

focused on the inviscid, single-blade-row code (ISTAGE) on

several parallel computers. The experience gained from this

exercise will be used to develop a parallel, viscous, multistage

(MSTAGE) version of the average-passage code.

Accomplishment Description

A parallel version of ISTAGE was demonstrated on the NASA

Lewis Hypercluster, the NASA Lewis Intel iPSC/860, and the

Intel Delta machines. A fine-grain partitioning of the grid in the

axial direction was used. The grid size used was 120 x 11 x 11.

A speedup of about 12 was achieved on all three machines using

20 processors. The code was run using up to 60 processors on

the NAS iPSC/860, which increased the speedup to about 15.

The results are summarized in the figure. A larger problem size

would likely result in better performance. The larger cases were

not run, however, so that the focus could be shifted to the

viscous code, MSTAGE. A coarse-grain parallel version of

MSTAGE was developed by running individual blade rows on

separate processors. A four-blade-row compressor simulation

was demonstrated on the Hypercluster and a network of IBM

RS6000 workstations. Four processors were used in each case.

Speedups of 3.91 and 3.78 were achieved on the Hypercluster

and the RS6000s, respectively. More significantly, an overall

performance of 44 MFLOPS, 30 percent of the Cray Y-MP single

processor performance, was achieved on the four RS6000s. This

translates to a performance/cost ratio of 220 MFLOPS per million

dollars for the RS6000s versus 50 MFLOPS per million dollars for

an eight-processor Cray Y-MP. APPL was used for all these cases,

allowing the same parallel code to run on four different parallel

computers without additional programming.

Significance

The initial demonstration of the ISTAGE code shows that

significant performance improvement is possible using parallel

processing. The results for the viscous code, MSTAGE, on the

RS6000 workstations offer hope for a cost effective alternative to

supercomputers. A parallel version of the average-passage code

will allow more cost effective execution by industry, as well as

within NASA Lewis for applications such as integrated computational fluid dynamics and experiments and numerical propulsion

system simulators. Parallel processing is also an important first

step toward developing codes which can be run interactively.

Future Plans

The parallel blade-row version of MSTAGE is operational, and

will be used in the ICE project for computational fluid dynamics

calculations of the Low Speed Compressor Facility. A fine-grain

parallel version of MSTAGE is currently being tested. This

version will be better suited to massively parallel processors,

and will be demonstrated on the Hypercluster, Intel iPSC/860,

and Delta machines. An interesting comparison will be the

performance of the coarse-grain version of MSTAGE on a few

powerful processors versus the fine-grain version using many

lower-performance processors.

1000 --

¢D

I--

100

I0

• iPSC/860

O iPSC/860*

0 Hypercluster

Cray Y-MP (1P)

• Delta

O

.o,o_o,oo..o_o

0 0 0 0 0

I I I I I I

2 5 10 20 50 100

Number of processors

*Compiler and algorithm optlmizations

Timings for the/STAGE code on several parallel computers;

20 iterations on a 120 x 11 x 11 grid.

60 Aeronautics/Propulsion
Subsonic Transport Propulsion/Airframe Integration

Pieter G. Buning, Principal Investigator

Co-investigators: Lie-Mine Gea, Cathy Maksymiuk, and W. R. Van Dalsem

NASA Ames Research Center/McDonnell Douglas Corporation

Research Objective

To test and refine the application of overset grid methodology

to Navier-Stokes flow simulation for subsonic transport

wing-body-pylon-nacelle geometries.

Future Plans

This effort is being extended to develop gridding strategies

for automatic generation of overset structured grids for

propulsion-airframe integration applications.

Approach

Pylon and nacelle grids were added to a suitably refined

wing-body grid generated using McDonnell Douglas grid

generation capabilities. The Air Force PFGSUS grid-joining code

was used to cut holes in the wing-body grid to accommodate

the pylon and nacelle. The Ames Research Center OVERFLOW

code was then employed to compute the flow on the overset

grids.

Accomplishment Description

The OVERFLOW flow solver was first validated in comparison

to the Langley Research Center thin-layer Navier-Stokes code

for wing-body configurations. Methodology was established for

adding pylon-nacelle grids, and comparisons were made with

wind tunnel and flight-test data. Areas in need of additional

investigation were identified. In addition, a new time-step

scaling was implemented that accelerates convergence to

steady state (in some cases by a factor of four).

Surface pressure on a subsonic transport; Mach = 0.85, Reynolds

number = 49 million (based on chord).

Significance

The capability for investigating propulsion-airframe integration

issues through Navier-Stokes flow simulation was established.

Overset grids on nacelle symmetry plane.

Aeronautics/Propulsion 61
Pulsed Detonation Wave Augmentor Concept

Jean-Luc Cambier, Principal Investigator

Co-investigators: Henry G. Adelman and Gene P. Menees

Eloret Institute/NASA Ames Research Center

Research Objective

To investigate the use of unsteady waves in a supersonic

combustor (scramjet), in order to enhance mixing and combustion rates. In this concept, blast waves are produced at the end

of small detonation tubes embedded in the combustor walls and

propagated into the main combustor stream. The interaction

between the strong pressure waves and the mixing layer

stimulates the formation of mixing vortices, while shock-heating

of the mixture allows for very rapid combustion. The detonation

tubes are rapidly cycled and triggered in sequence.

Approach

Preliminary computational fluid dynamics studies of various

generic engine configurations were performed to estimate

the effectiveness of the concept and to identify the key

elements for future parametric studies. The numerical simulations were performed using a two-dimensional, time-accurate

Navier-Stokes code, coupled to the full chemical kinetics of

hydrogen-air combustion.

Future Plans

We will investigate these unsteady interactions in more detail,

using sophisticated nonequilibrium, compressible turbulence

models. Other configurations will also be studied, notably

external combustion and the stimulation of hydrocarbon

combustion.

Publication

Cambier, J. -L. et al.: Numerical Simulations of a Pulsed Detonation Wave Augmentation Device. 29th AIAA Joint Prop. Conference, Monterey, CA, June 1993, AIAA Paper 93-1985.

Accomplishment Description

The studies demonstrated the validity of the concept for a

specific engine configuration, shown in the figures, where a

short detonation tube was located beneath the fuel layer of a

scramjet engine. The wave diffracted at the tube exit and

decayed into a blast wave. The exiting shock was transmitted

through the fuel layer, which it compressed and heated to

ignition temperature. A rapid burning occurred at that location,

and it is clearly seen in the figure (a). The fuel layer was then

pushed and rolled-up by the transmitted shock, creating a large

vortex. After only 275 microseconds, the fuel layer had significant amounts of vorticity and stretching present--both contribute

to mixing enhancement (figure (b)). The shock heating was also

very efficient at stimulating the combustion, which was seen to

occur at several locations. Another important feature was the

transmittal of a shock within the fuel layer. This particular feature

may be used in other engine configurations, such as external

combustion.

Significance

Strong, unsteady shocks are a powerful mechanism for stimulating the mixing and combustion in a scramjet combustor. This

may lead to the design of shorter, lighter engines, with higher

performance at high flight speeds. The detonation tubes also

significantly contribute to the engine thrust, and they have

minimal operating penalties.

(a)

(b)

Temperature contours, at (a) 75 and (b) 275 microseconds.

62 Aeronautics/Propulsion
Evaluating Nozzle Drag Using Computational

Jeffrey A. Catt, Principal Investigator

Co-investigators: Tracy J. Welterlen and Brett W. Denner

Lockheed Engineering and Sciences Company

Research Objective

To define a methodology that uses computational fluid dynamics

(CFD) to predict jet-induced nozzle drag. The methodology

should accurately predict the difference between two operational exhaust system conditions as compared to test data.

Approach

Using the HAWK3D flow solver, three-dimensional full

Navier-Stokes solutions of several F-16 jet effect test points

were obtained. Two configurations simulating a reference

"zero-nozzle drag" configuration and the actual nozzle installation were evaluated at two Mach numbers. The drag of each

configuration was calculated by integrating pressure over

surfaces that represented the metric portion of the model. Drag

increments between the two configurations were compared to

tested drag increments.

Accomplishment Description

Comparisons of predicted afferbody pressure distributions with

pressure data gathered during jet effects testing showed that

HAWK3D accurately modeled jet entrainment. Balance readings

for incremental drag differences also matched test data very well.

These comparisons showed the validity of using CFD to predict

jet-induced drag increments. However, highly accurate computational procedures must be applied in a rigorous manner to

achieve valid results. The computational model must accurately

reflect the test environment, especially in the definition of the

tested model. It is important that drag is evaluated on only

those portions of the model that are force sensitive. Tracking

quantities of merit (drag, for this study) as a solution progresses

provides needed insight on convergence. This tracking is

critical for jet-induced drag evaluation, because numerical

quantities can indicate convergence long before fluctuations

in afterbody pressure distribution damp out. Convergence of a

typical 2.5 million point problem required approximately

20,000 iterations, 92 megawords of memory, and

300 Cray Y-MP hours.

Fluid Dynamics

Surface pressure contours and sonic regions produced on the

F-16 at Mach 0.9; low pressure regions = blue, high pressure

regions = magenta.

Significance

Propulsion integration is a driving technology in the development of advanced fighter aircraft. This is evident in the design of

the F-16 afterbody, which requires that the nozzle be cleanly

integrated into the airframe. Predicting drag characteristics with

CFD provides cost-effective and timely input into the design

process and provides risk reduction before manufacturing is

initiated. CFD may prove to be a viable alternative to wind

tunnel testing.

CFD predicted nozzle pressure contours (left) compared with test

data (right). Test contours interpolated from 40 pressure taps.

Future Plans

Preliminary CFD results show that refinements to external nozzle

shaping can result in significant station improvement. Work is

continuing through the prediction of jet-induced drag of

advanced nozzle concepts installed on the F-16.

Aeronautics/Propulsion 63
CFD for Engine-Airframe Integration

Richard D. Cedar, Principal Investigator

Co-investigator: Craig M. Kuhne

General Electric Aircraft Engines

Research Objective

To develop an analytical tool to determine the mutual aerodynamic interaction between an engine and an airframe. Emphasis

is on the development of an analysis system that can be integrated into the engine-airframe design process.

Approach

The current computational fluid dynamics (CFD) based design

and analysis systems at General Electric Aircraft Engines were

used to assess the mutual aerodynamic interaction of isolated

components (engine inlets, exhaust systems, clean wings). This

assessment was pursued using the Chimera domain decomposition technique to build the complete configuration from the

isolated component grids. The PEGSUS code was used to

determine the interpolation coefficients required by the Chimera

method, and the CFL3D code was used as the flow solver.

Accomplishment Description

An advanced high-bypass-ratio engine was tested in the

NASA Langley National Transonic Facility wind tunnel. Threedimensional Euler and Navier-Stokes calculations were performed for the isolated engine, the clean-wing airframe (with

no engine installation), and the complete configuration (airframe

including the engine and pylon). Comparisons with test data

showed that the major aerodynamic interactions were predicted.

These included an increase in the wing upper-surface Mach

numbers inboard of the installation and a reduction outboard

(see figure) due to the engine installation. A typical Navier-Stokes

solution takes about 20 Cray-2 hours and 64 megawords of

memory.

Significance

Assessment of engine and airframe aerodynamic interaction is

critical for optimum aircraft design.

Future Plans

Work is in progress to make more efficient use of the computational grid while concentrating the grid in the areas where it is

deficient for high-quality Navier-Stokes solutions. Further

evaluation of the results is planned when test data become

available from a highly instrumented model of the MDC

cooperative study aircraft that is to be tested in the NASA

Langley 16-Foot Wind Tunnel.

Publications

1. Ostrander, M. J.; and Cedar, R. D.: Analysis of a High Bypass

Ratio Installation Using the Chimera Domain Decomposition

Technique. AIAA Paper 93-1808, June 1993.

2. Cedar, R. D.; Dietrich, D. A.; and Ostrander, M. J.: Engine/

Airframe Installation CFD for Commercial Transports. An

Engine Manufacturers Perspective. SAE Aerotech '93,

Sept. 1993.

Decreased Mach

Number

Increased Mach

Number

Surface pressure distribution on the (top) clean wing and

(bottom) installed engine configurations.

64 Aeronautics/Propulsion
Computation of Separated Nozzle Flows

C. L. Chen, Principal Investigator

Co-investigators: S. R. Chakravarthy and C. M. Hung

Rockwell International Science Center/NASA Ames Research Center

Research Objective

To understand separated overexpanded nozzle flows using a

time-accurate Reynolds-averaged Navier-Stokes solver.

Approach

The equations solved were the unsteady, axisymmetric

Reynolds-averaged Navier-Stokes equations. The time-accurate

unified solution algorithm (USA) Navier-Stokes code was

employed to calculate the separated nozzle flow. Time accuracy

was required to study the transient flow and the unsteady nature

of the separated flow. The USA code, which is a finite-volume

multizonal Navier-Stokes solver, utilizes a high-resolution total

variation diminishing scheme. The code was validated from low

subsonic to hypersonic flows.

Accomplishment Description

The axisymmetric overexpanded nozzle flows were studied, and

the flow structures of the startup and throttle-down processes

were examined. During the impulsive startup process, observed

flow features included the Mach disc, separation shock, Mach

stem, vortex core, contact surface slip stream, initial shock

front, and shocklet. The movement of the Mach disc was not

monotonical in the downstream direction. For a range of

pressure ratios, hysteresis phenomenon occurred; different

solutions were obtained depending upon different processes

(see accompanying figure).

Significance

The separated flow structures associated with the hysteresis

phenomenon were revealed. A typical time-accurate calculation

to reach final-state solution required 40 Cray Y-MP hours and

10 megawords of memory.

Future Plans

The bifurcation diagram will be further investigated and examination of three-dimensional flow structure will be continued.

Publication

Chen, C. L.; Chakravarthy, S. R.; and Hung, C. M.: Numerical

Investigation of Separated Nozzle Flows. AIAA Paper 93-3016,

July 1993.

Schlieren-type snapshots and streamlines at pressure ratio = 45. (top) Startup. (bottom) Throttle down.

Aeronautics/Propulsion 65
Inverse Design of Installed Nacelle

H. C. Chen, Principal Investigator

Co-investigators: T. Y. Su, T. J. Kao, and D. A. Naik

Boeing Commercial Airplane Group/NASA Langley Research Center/ViGYAN, Inc.

Research Objective

To reshape an installed nacelle to achieve a specified nacelle

pressure distribution.

Approach

The flow solver was a general, multiblock, multigrid Euler code

(GMBE) developed for the analysis of complete airplane

configurations with underwing turbofans. GMBE was coupled

with the direct iteration surface curvature algorithm, and the

resulting new code (GMBEDS) was used for the inverse design

of a turbofan nacelle installed on a transport airplane.

Accomplishment Description

GMBEDS was executed successfully for a generic low-wing

transport with two different turbofan nacelles (bypass ratio 6).

The analysis run from the first installed nacelle was used to

obtain the initial pressure distribution. This nacelle provided the

initial geometry from which to start the design iteration. The

pressure distribution from an analysis of the second installed

nacelle was used as the target. The goal was to see how close

the GMBEDS-designed nacelle would be to the target nacelle.

The volume grid contained approximately 1.2 million grid

points and was composed of 32 blocks. A typical analysis

of 500 multigrid steps for third-order convergence takes

220 minutes and 20 megawords of memory on the Cray Y-MP.

The analysis after each design iteration requires only first-order

convergence. Typically, 5-10 iterations are required. The figure

shows how the initial and target pressures compare with those of

the designed nacelle after six design iterations.

Significance

The computational fluid dynamics code developed in this

project is an effective tool for the inverse design of installed

turbofan nacelles. The computational cost of the design is of

the same order as that of the analysis.

Future Plans

The same methodology will be applied to a chosen multiblock

Navier-Stokes code.

Cp

Cp

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--.2

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-.6

--,2

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.4

.6

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.4

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Pressure comparison for installed nacelle design; free-stream Mach number = 0.85, angle of attack = 0.5 degrees. Dotted line = initial,

solid line = target, dashed line = design.

66 Aeronautics/Propulsion
Ai rframe/I n let Aerodynam ics

Wei J. Chyu, Principal Investigator

Co-investigators: Tom I-P. Shih and David A. Caughey

NASA Ames Research Center/Carnegie Mellon University/Cornell University

Research Objective

To develop the analytical capability to predict integrated

performance of forebody-inlet systems for highly maneuverable

aircraft.

Approach

Navier-Stokes codes were adapted and developed using two

numerical approaches, one with the F3D Navier-Stokes code

combined with the Chimera grid-embedding technique, and the

other with a finite-volume method with multiblock and multigrid

diagonal implicit schemes.

Accomplishment Description

Studies were made on (1) an inlet boundary layer bleeding with

one or multiple flat plates with circular holes that vent into a

plenum, (2) an inlet flow through a highly contoured and offset

super-elliptic diffuser with a newly developed three-dimensional

finite-volume Navier-Stokes code, and (3) a transonic flow

around the forebody-inlet of the AV-8B II Harrier aircraft.

Significance

The study permitted analysis of complex inlets, making possible

the investigation of a broader range of design variables associated with integration of the forebody and inlet systems and inlet

performance with boundary layer control.

Future Plans

Efforts will focus on code validation cases with various turbulence models to determine the effects on the flow physics and

inlet performance.

Publications

1. Chyu, W. J.; Howe, G. W.; and Shih, T. I-P.: Bleed Boundary

Conditions for Numerically Simulated Mixed-Compression

Inlet Flow. J. of Propulsion and Power, vol. 8, 1992.

2. Wang, L.; and Caughey, D. A.: Multiblock/Multigrid Euler

Method to Simulate 2D and 3D Compressible Flow. AIAA

Paper 93-0332, Jan. 1993.

3. Mysko, J. M.; Chyu, W. J.; and Chow, C. Y.: Navier-StOkes

Simulation of External/Internal Transonic Flow on the

Forebody/Inlet of the AV-8B Harrier II. AIAA Paper 93-3057,

July 1993.

4. Chyu, W. ].; Rimlinger, M. J.; and Shih, T. I-P.: Effects of

Bleed-Hole Geometry and Plenum Pressure on ThreeDimensional Shock-Wave/Boundary-Layer/Bleed Interactions. AIAA Paper 93-3259, July 1993.

_ 0,0 2.0 3.9

Boundary layer bleed through multiple slanted circular holes.

P/PTO

1,00 |

A V-8B II Harrier forebody-inlet surface pressure.

Aeronautics/Propulsion 67
Three-Dimensional High-Speed Plume-Propulsive Flow Fields

Sanford M. Dash, Principal Investigator

Co-investigators: Neeraj Sinha, Brian J. York, Robert A. Lee, Ashvin Hosangadi, and Donald C. Kenzakowski

Science Applications International Corporation

Research Objective

To establish and advance computational methodology for

the simulation of steady and transient chemically reacting,

multiphase, high-speed plume-propulsive flow fields. Advanced

turbulence models and new techniques for simulating nonequilibrium thermochemical and multiphase flow processes are

being developed.

Approach

The CRAFT implicit/upwind code with strongly coupled, largematrix inclusion of multi-equation turbulence models and

chemical species was extended to incorporate particulate

nonequilibrium capabilities using Eulerian and Lagrangian

formulations. In addition, a gas-liquid formulation was incorporated with rate-dependent vaporization and combustion. For

steady problems, Eulerian particle formulation and Reynoldsstress-based turbulence modeling was utilized. For nonsteady

problems, the Lagrangian particle formulation with large-eddysimulation based turbulence modeling was utilized.

Accomplishment Description

Grid methodology was upgraded to provide zonal patching in a

dynamic framework. Particulate methodology was upgraded to

account for volumetric contributions, and a new Lagrangian

solver was incorporated and integrated with a solid-propellant

ball-grain formulation. Large-eddy-simulation subgrid stress

models were incorporated for transient studies. Problems

analyzed included rocket plume-propulsive flows, ducted

rockets, and plume-launcher interactions. The upgrades

produced unique computational capabilities for steady and

transient combusting/multiphase flows. Each run required

30-120 megawords of memory and 10-50 Cray Y-MP hours,

depending on the grid dimensions and the number of chemical

species included.

Significance

CRAFT plays a major role in missile design, in vertical launcher

design, and in propulsive system design. It also supports major

hypervelocity gun programs.

Future Plans

Multiphase flow modeling will be significantly advanced.

Volumetric contributions, particle-particle interactions, and

combustion modeling will be emphasized, as well as strong

coupling to ablative boundaries and a detailed treatment of

particle-surface interactions. Advanced work on droplet

formation at transient interfaces will also be performed.

Interaction flow field for a rocket plume in a vertical launcher. Velocity vectors colored by Mach number.

68 Aeronautics/Propulsion
Turbine-Blade Tip Clearance Flows

Frederik I. de Jong, Principal Investigator

Co-investigator: Tony Chan

Scientific Research Associates, Inc.

Research Objective

To analyze the tip clearance region flow for the gas generator

oxidizer turbine (GGOT) blade and to study advanced concepts

to reduce tip clearance losses.

Approach

The finite-difference form of the compressible, three-dimensional

Navier-Stokes equations was solved for a single blade row in a

rotating frame of reference using a linearized block-implicit

alternating-direction-implicit procedure. The grid was generated

by stacking a series of two-dimensional grids from hub to casing;

grid points were clustered near solid boundaries and in the tip

clearance region. At the inflow boundary, located approximately

one axial chord from the blade leading edge, flow profiles were

specified, while at the outflow boundary the static pressure was

prescribed.

Accomplishment Description

Results were obtained for the baseline GGOT blade at the fullscale Reynolds number on a grid containing approximately

216,000 points: 90 points in the streamwise direction (36 points

on the blade), 60 points in the circumferential direction

(30 points on the blade), and 40 points from hub to casing

(12 points in the clearance region). Particle traces in the near-tip

region showed vortical flow behavior of the fluid which passed

through the clearance region and exited at the downstream edge

of the gap. In an effort to reduce clearance flow losses, the minishroud concept, which added a fence to the pressure side of the

blade at the blade tip, was proposed. Calculations performed on

the GGOT geometry with the minishroud, however, indicated

that the minishroud had only a marginal effect on the tip

clearance losses at the design tip clearance. Calculations at

larger tip clearances are currently in progress. Each calculation

required approximately 10 Cray Y-MP hours and 11 megawords

of memory.

Significance

The ability to predict accurately the tip clearance flow and to

perform parametric studies on changes in the geometry (both

blade shape and tip clearance) is essential to the design of highperformance turbines.

Future Plans

Further three-dimensional simulations will be performed to

assess the effect of tip clearance and to study additional

advanced concepts.

Publication

Chan, Y.-T.; and de Jong, F. J.: Navier-Stokes Analysis of an

Oxidizer Turbine Blade with Tip Clearance with and without a

Mini-Shroud. Presented at the 1 lth Workshop for Computational

Fluid Dynamics Applications in Rocket Propulsion, Huntsville,

Ala., April 1993.

Pressure contours and particle traces on the baseline GGOT blade.

Aeronautics/Propulsion 69
Interaction of Turbulence and Chemical Reaction

Richard L. Gaffney, Jr., Principal Investigator

Co-investigators: Jeffery A. White, Sharath Girimaji, and J. Phillip Drummond

Analytical Services and Materials, Inc./Institute for Computer Applications in Science and Engineering/NASA Langley Research Center

Research Objective

To investigate the interaction of turbulence and chemical

reaction in high-speed turbulent chemically reacting flows.

Approach

The Favre-averaged Navier-Stokes equations and 8-species

continuity equations with an 18-step reaction model were solved

to provide various mean quantities in chemically reacting flows.

A k-_ turbulence model for the Reynolds stresses provided part

of the interaction of turbulence and chemical reaction. The

effects of temperature fluctuations on the mean species reaction

rate were modeled by an assumed marginal probability density

function (PDF) for temperature. The temperature variance that

affected the shape of the PDF was determined from the solution

of an equation for the enthalpy variance.

Accomplishment Description

Assumed Gaussian and beta PDFs were used to model the

effects of temperature fluctuations on chemical reactions in

turbulent flows. Results indicated that the effects of temperature

fluctuations either increase reaction rates or decrease them

depending on the mean temperature, the fluctuation intensity,

and the reaction mechanism under consideration. Numerical

simulations indicated that in hydrogen-air reactions, temperature fluctuations can have a significant effect on the ignition

delay time. These fluctuations directly affect scram jet combustor

length and efficiency. Simulations of a hydrogen-air coaxial

jet configuration were completed. Typical calculations took

60 megawords of memory and 40 Cray Y-MP hours of computer

time.

Significance

Temperature and species fluctuations can influence the performance of a scramjet by affecting chemical reaction rates.

The effects of temperature and species fluctuations must be

understood and modeled appropriately to optimize scram jet

performance.

Future Plans

This work will be extended to include the effects of species

fluctuations and determine the effect that they have on scramjet

performance.

Publication

Gaffney, R. L, Jr.; White, J. A.; Girimaji, S. S.; and Drummond,

J. P.: Modeling Turbulent/Chemistry Interactions Using Assumed

PDF Methods. Joint Propulsion Conference, Nashville, Tenn.,

AIAA Paper 92-3638, 1992.

Mach contours inside the outer (vitiated air) nozzle and downstream of the exits of the hydrogen and air nozzles (solution inside the

hydrogen nozzle not shown). Dark blue = low Mach number, red = high Mach number.

70 Aeronautics/Propulsion
Unsteady Turbomachinery Computations

Karen L. Gundy-Burlet, Principal Investigator

Co-investigator: Akil Rangwalla

NASA Ames Research Center/MCAT Institute

Research Objective

To investigate unsteady rotor-stator interaction in multistage

turbomachines.

Approach

The three-dimensional (3-D) Navier-Stokes codes, STAGE-3 and

ROTOR-4, incorporate the most modern, high-order (third-order

spatial and second-order time-accurate) upwind-biased schemes

for the solution of the thin-layer Navier-Stokes equations. The

schemes were set in an iterative implicit framework. The codes

used a multizone grid with some zones moving relative to

others. Information was transferred between the various zones

using temporally- and spatially-accurate zonal boundary

conditions.

Accomplishment Description

A fine-grid 3-D calculation of a turbine stage was completed;

however, the numerical results indicated that twice as many

points would be needed in the radial direction to obtain

sufficient accuracy for a turbine-loss calculation. Because of

insufficient computational resources, this grid refinement was

not carried out. A fine-grid 3-D calculation of a multistage

compressor was initiated. Initial results from the code compared

well with experimental data for time-averaged surface-pressure

distributions, but the unsteady flow field was still developing.

The code required 16 megawords of memory and 70 megawords

of solid-state device memory for the 1.8-million point grid. A

typical converged solution was estimated to require approximately 500 Cray Y-MP hours for a converged time-averaged

solution. The figure shows preliminary instantaneous surface

pressures within the compressor.

Significance

Unsteady turbomachinery flow fields are extremely complex,

especially in the latter stages of multistage turbomachines.

Wake-wake and wake-airfoil interactions cause complex timevarying forces on the downstream airfoils. It is important to

understand these interactions to design turbomachines that are

light, compact, reliable, and efficient.

Future Plans

The 3-D computation of the flow within the multistage compressor is continuing. On completion of the 2.5 stage compressor

computation, the results will be compared with experimental

data from a 2.5 stage compressor investigated. Future computations will include a 4.5 stage compressor geometry and a

multistage/multi-airfoil turbine computation.

Preliminary instantaneous surface pressures in a 2.5 stage compressor.

Aeronautics/Propulsion 71
Three-Dimensional Transonic Compressor Stage Flows

Chunill Hah, Principal Investigator

Co-investigators: J. Loellbach, S. L. Puterbaugh, and W. W. Copenhaver

NASA Lewis Research Center/ICOMP/WL/FIMM, Wright Patterson AFB

Research Objective

To simulate numerically the unsteady flow field inside a

transonic, high-through-flow axial compressor stage, and to

compare computed results with available experimental data.

Approach

The three-dimensional, unsteady Navier-Stokes equations are

solved on structured grids using second-order accurate implicit

time integration, third-order accurate upwind differencing for the

inviscid fluxes, and second-order accurate centered differencing

for the viscous fluxes. A two-equation turbulence model with a

low-Reynolds-number modification is used for turbulence

closure.

Publications

1. Hah, C.: Unsteady Aerodynamic Flow Phenomena in a

Transonic Compressor Stage. AIAA Paper 93-1868,

June ] 993.

2. Hah, C.; and Puterbaugh, S. L.: A Critical Evaluation of a

Three-Dimensional Navier-Stokes Method as a Tool to

Calculate Transonic Flows Inside a Low-Aspect-Ratio

Compressor. AGARD Proceedings, 1990.

Copenhaver, W. W.; Hah, C.; and Puterbaugh, S. L.:

Three-Dimensional Flow Phenomena in a Transonic HighThrough-Flow Compressor Stage. ASME J. Turbomachinery,

vol. 115, no. 2, 1992.

3.

Accomplishment Description

The numerical method simulated the unsteady flow field inside

a transonic, high-through-flow compressor stage consisting of

20 low-aspect-ratio rotor blades and 31 stator blades. For the

simulation, a set of two rotor blades and three stator blades were

used, along with a time-lagged periodic boundary condition. The

flow fields at three typical operating conditions (near choke, near

peak efficiency, and near stall) were computed. Approximately

630,000 grid points were used for each calculation. A typical

run required approximately 300 hours of single CPU time and

32 megawords of memory on a Cray Y-MP. Time-averaged

quantities from the computed results compare well with those

of the experimental data. The figure shows the instantaneous

static pressure distribution on the blade and hub surfaces for

two different time steps at the near-peak-efficiency operating

condition. Small differences can be seen between different

passages at the same time step and between corresponding

passages at different time steps, illustrating both the nonperiodicity and the unsteadiness of the computed flow field.

Significance

This work is a direct extension to unsteady flow analysis of an

existing steady Navier-Stokes solver that has been tested and

applied for a wide range of turbomachinery flows. Unsteady

analysis is necessary for accurate flow-field prediction for closely

coupled blade rows within a stage or between successive stages.

Instantaneous static pressure distributions at two different time

steps.

Future Plans

Efforts are under way to analyze the unsteady effects of various

inlet flow distortions and to perform design optimizations.

72 Aeronautics/Propulsion
Eval uati ng H igh-Angle-of-Attack In let Characteristics

Doug G. Howlett, Principal Investigator

Co-investigators: Brett W. Denner and Chris L. Reed

Lockheed Engineering and Sciences Company

Research Objective

To establish a rigorous methodology for evaluating

shock-boundary layer interaction using the three-dimensional

full Navier-Stokes flow solvers HAWK3D and CDFalcon. This

project served as a validation exercise for the newly developed

CDFalcon flow solver.

Approach

Flow-field predictions were obtained using CDFalcon and

HAWK3D. CDFalcon is a finite-volume flow solver similar to

HAWK3D. Validation of these codes focused on their ability to

model complicated flow phenomena. A test condition at

Mach 0.3 and 30 degrees angle of attack was chosen because it

produced a strong normal shock that interacted with a separation bubble downstream of the inlet cowl lip. Since inlet

performance and basic airflow are highly dependent upon the

ability to model accurately shock-boundary layer interaction

and separation, accurate analysis of this condition is a demanding test for any flow solver. Analysis of these solutions consisted

of obtaining a converged solution on a grid that was not

optimized for separated flow, adapting the grid to local flow

gradients, interpolating the existing solution onto the adapted

grid, and resolving the flow field.

Surface pressure contours and strake vortex particle traces. Blue =

low pressure regions and magenta = high pressure regions.

Accomplishment Description

Predicted flow characteristics showed good agreement between

HAWK3D, CDFalcon, and test data. Flow-field phenomena in

and around the inlet developed as anticipated. CDFalcon

predicted duct surface pressure distributions that showed good

agreement with test data. Although HAWK3D analyses are not

complete, they show that all primary flow field features have

been captured. Duct pressure distributions, pressure recovery,

and compressor face distortion were tracked as solutions

converged and were very near or approaching tested values.

Significance

High-angle-of-attack maneuver flight conditions produce

complex flow fields that are difficult to predict numerically.

Recursive grid adaptation can overcome this limitation, but it

usually imposes unwieldy increases in computational requirements. Single-step grid adaptation provides a practical way to

achieve accurate results for these solvers. The most desirable

solution would be an efficient unstructured-grid flow solver with

automated grid refinement capability.

Centerline Mach contours shovel significant lip separation and a

subsequent normal shock. Blue = static regions and magenta =

the maximum Mach (approximately 1.5).

Future Plans

The HAWK3D solution of this flow field and the use of similar

procedures to analyze advanced boundary layer management

concepts for fighter aircraft application will be pursued.

Aeronautics/Propulsion 73
Computational Analysis of a Tip-Engine Configuration

J. Mark Janus, Principal Investigator

Co-investigators: Animesh Chatterjee and Chris Cave

Mississippi State University

Research Objective

To perform a computational flow analysis of a design concept

centered around induced drag reduction and tip-vortex energy

recovery.

Approach

The tested geometry was a semispan model having an unswept,

untapered wing with a tip-mounted SR-2 blade design highspeed propeller in a pusher configuration. The flow conditions

were set to Mach 0.70 and angles of attack ranging from

approximately -2 to 4 degrees. The flow model solved the

unsteady three-dimensional Euler equations, discretized as a

finite-volume method, utilizing a high resolution approximate

Riemann solver for cell interface flux definitions. The numerical

scheme was an approximately factored block, lower-upper

implicit Newton iterative-refinement method. Multiblock domain

decomposition was used to partition the field into an ordered

arrangement of blocks. Block-block relative motion was

achieved using local grid distortion.

Future Plans

Thus far, all solutions obtained have been inviscid; therefore no

physical flow separation at the wing tip has been modeled. This

separation may have a pronounced effect on the lift-off and

trajectory of the tip vortex. Efforts are presently under way to

obtain a viscous solution in an attempt to better determine the

precise physics of the tip-vortex and its lift-off from the solid

surface. Also under way are efforts to quantify better induced

drag using a wake integral technique in lieu of surface pressure

integration.

Accomplishment Description

Three configurations were analyzed: (1) a baseline fuselage-wing,

(2) a fuselage-wing-nacelle, and (3) a fuselage-wing-nacelle

propfan. In order to correlate adequately the relationship

between angle of attack and interference effects, several

numerical simulations at differing inflow angles were necessary.

Data collected from the numerical solutions included aerodynamic force coefficients, propeller performance coefficients,

and flow-field maps. Integration of surface pressure to obtain

the coefficients was satisfactory for all but the drag calculations.

Very good agreement has been obtained with configuration lift

comparisons. An excellent correlation with the expected wind

tunnel interference effects was found by comparing a free-flight

simulation and a simulation including tunnel walls. The processing rate on a Cray Y-MP for each iteration of each time step was

approximately 22,000 computational cells per second. A typical

steady run took 15 Cray Y-MP hours and required 10 megawords

of internal memory and 40 megawords of solid-state storage

device memory. Unsteady runs took about 25 Cray Y-MP hours.

Significance

NASA wind tunnel tests have indicated large reductions in

induced drag for wings with tip-mounted engines. This study

used numerical simulation to investigate innovative methods for

significantly reducing the induced drag of wings. Numerical

simulations similar to those of the NASA investigation were

performed to corroborate their data and to obtain insight into

the performance enhancing mechanisms involved. Calculations

were performed to complement NASA's results and to help guide

future NASA research and system studies that address induced

drag reduction technologies and vortex hazard alleviation

studies.

Fuselage-wing-nacelle-propfan shaded surface pressure

contours (Mach = 0.7, angle of attack = 4 degrees).

74 Aeronautics/Propulsion
Studies of Mixing and Combustion

Jinho Lee, Principal Investigator

Sverdrup Technology, Inc.

in Supersonic Flows

Research Objective

To simulate numerically various three-dimensional (3-D)

turbulent/chemically-reacting and nonreacting (mixing) flow

fields inside supersonic combustors to investigate possible

fuel-air mixing/combustion enhancement techniques to increase

thrust. Different aspects of the turbulent and finite-rate chemistry

models and their influence on combustor predictions are

studied.

Approach

The analysis was performed using a reactive propulsion code

based on the lower-upper scheme RPLUS. The effects of

hydrogen-air chemical reactions were modeled using a

9-species/14-reaction finite-rate chemistry model. A modified

zero-equation Baldwin-Lomax model was used to model the

effects of turbulence. The inverse square length scale relationship was used to extend the turbulence model into a 3-D regime,

and a simple gradient-diffusion relationship was used to model

turbulent mass diffusion. A multiple-block structured-grid

approach was used to model complex flow geometries.

Accomplishment Description

Steady mixing enhancement mechanisms used in existing wall

mounted fuel injector designs were analyzed. The mixing

characteristics of the transverse injector model, the multiple

rectangular swept-ramp injector model, and the swept-ramp

injector model were studied. A typical swept-injector model

geometry, the flow condition, and the three-block grid system

are shown in the first figure. This injector/combustor configuration was simulated using a three-block two-species (air-air)

turbulent nonreacting model using 220,000 cells, approximately

13 megawords of memory, and 10 Cray Y-MP hours. A typical

streamwise cross-section contour of the predicted injected gas

mass fraction is compared with experimental data in the second

figure•

Significance

The Baldwin-Lomax algebraic turbulence model predicted the

near-wall turbulence in the near-equilibrium region of the

combustor, but the model formulation was weak in predicting

the mixing-layer behavior and the non-equilibrium behavior

around the injectors. There was reasonable agreement between

the predictions and the experimental data. The swept-injector

design had superior mixing behavior over a typical transverse

injector design, and enhanced mixing was developed through

the streamwise vorticity generated by the swept ramp.

Future Plans

Effects of parameters in the injector exit conditions, injector

model, and injector grid distribution on the streamwise vorticity

generation will be studied. Mixing behavior modifications

caused by physical models and their effectiveness in modeling

combustor environments will also be studied.

_h 2o

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Mach 2.0, 9.5-deg_ree swept-ramp streamwise injector model

and a typical three-block grid system.

(a)

(b) llllllllllJl

Streamwise injected mass-fraction contour comparisons;

experimental data (left) and computed solution (right).

Nondimensional streamwise locations: (a) 1.06, (b) 2.13, and

(c) 3.19. Mass: black = 0.0 to white = 1.0.

Aeronautics/Propulsion 75
Turbulent Base Heating Computational

E. D. Lynch, Principal Investigator

Co-investigators: R. Lagnado and Y. Hsu

Rockwell International, Rocketdyne Division

Research Objectives

To predict the level of base fuel entrainment and heating to be

anticipated under conditions characteristic of National Launch

System (NLS)-Iike vehicles and to examine the physical behavior

of turbulence models appropriate to those systems.

Approach

Traditional approaches for predicting base heating have used

empirically based correlations to scale existing flight data.

However, the configuration and flow conditions for modern

launch systems are significantly different than previous configurations in how densely the engines are packed, the type of

exhaust gases, and the degree of overexpansion of the nozzle

plumes. The unified solution algorithm computational fluid

dynamics (CFD) code was used to model the NLS base region

flow at 10,000 and 50,000 feet altitude to predict the level of

base heating. Full Navier-Stokes, three-dimensional (3-D),

200,000-grid-point analyses were made of the geometry,

including all 6 nozzles and their external contours. The effects

of the physical models were found through parametric axisymmetric analyses with various turbulence and kinetics models.

Accomplishment Description

The results of the computations showed little fuel entrainment

into the base region and inconsequential base heating at either

Fluid Dynamics

altitude. At 10,000 feet the nozzle plume was overexpanded and

an oblique shock formed downstream in the nozzle plume. The

vehicle airstream was responsible for driving a 3-D recirculation

region in the base; however, no fuel was driven into the base. At

50,000 feet the plumes were underexpanded and interacted in

the line between the plumes, but the level of base entrainment

was insufficient to raise the base temperature above acceptable

limits. These 3-D base-heating analyses required an average of

50 Cray Y-MP hours and 20 megawords of memory.

Significance

Use of fuel-rich turbine exhaust gases and film cooling to cool

the nozzle walls in NLS-like launch systems could result in hot

gases recirculating and burning in the vehicle base region and

could lead to an unacceptable level of base heating. These CFD

analyses indicated that base heating was not a significant issue

and prevented a costly redesign of the NLS.

Future Plans

Future analyses will employ turbulence models more appropriate

to defining the flow field in combusting flows such as vehicle

base regions.

2.50

1.88

1.25

0.63

0.00

Mach number contours at 50,000 feet altitude in planes proceeding outward from vehicle base.

76 Aeronautics/Propulsion
U nsteady MuLtistage Turbomachinery Applications

Nateri K. Madavan, Principal Investigator

Co-investigator: San jay Mathur

MCAT Institute/NASA Ames Research Center/Iowa State University

Research Objective

To evaluate the capabilities of the CM-2/CM-5 in predicting

unsteady flows in multistage turbomachines.

Approach

In recent years, the ROTOR and STAGE families of codes have

been developed at NASA Ames Research Center for the prediction of unsteady flows in turbomachines. These codes solve

the thin-layer Navier-Stokes equations in a time-accurate

manner using state-of-the-art upwind algorithms and zonal

methodologies. Implementation of these codes has been limited

to vector machines. The thrust of the present research is toward

implementation of these codes on the CM-2/CM-5 and toward

algorithm modifications to exploit the CM-2/CM-5 architecture

to improve run-time performance and decrease overall computing times. Although three-dimensional capability remains the

eventual goal, the focus here is on the two-dimensional

STAGE2 code.

Accomplishment Description

Migration of the STAGE2 code to the CM-2/CM-5 was completed and the flow in a 3.5 stage experimental compressor

was computed. Code migration included selecting optimal data

structures and bookkeeping strategies, revisions to the algorithm

that were necessitated by the CM-2/CM-5 hardware, and

thorough checking of the interim results with the Cray version.

The current implementation on the CM-2 incorporates the

communication compiler for improved zonal boundary

information transfer and the new slicewise compiler. The

communication compiler has not been incorporated in the

CM-5 implementation. Several routines in the STAGE2 codes

result in extremely poor performance on the CM-2 when coded

entirely in CM FORTRAN. Significant performance improvements were achieved by replacing the FORTRAN code with

calls to the CM FORTRAN utility library.

Significance

A time-accurate, implicit algorithm for the Navier-Stokes

equations in two dimensions was successfully implemented on

the CM-2/CM-5. The algorithm used multiple overlapped and

patched zones and high-order upwind techniques. Modifications

to the original algorithm resulted in more optimal performance

on the CM-2/CM-5.

Future Plans

Future plans include algorithm developments for improved

performance and extensions to three dimensions.

Publication

Madavan, N. K.: Unsteady Turbomachinery Flow Simulations

on Massively Parallel Architectures. Computing Systems in

Engineering, vol. 3, nos. 1-4, 1992, pp. 241-249.

Entropy contours in a 3.5 stage axial compressor. Results from a two-dimensional multistage calculation.

Aeronautics/Propulsion 77
Base Pressurization

Charles R. McClinton, Principal Investigator

Co-investigator: Paul Vitt

NASA Langley Research Center/Analytical Services and Materials, Inc.

Methods for Scram jet Combustors

Research Objective

To use computational fluid dynamics (CFD) modeling to

evaluate parametrically seven different methods of decreasing

pressure drag on step expansions in supersonic combustor flows

at Mach 16. Main fuel injection combustion effects and step

height variation were also evaluated.

Approach

The General Aerodynamic Simulation Program was used in the

steady-state elliptic, fulI-Navier-Stokes mode for all the computations in this study. Most of the cases were modeled as twodimensional flow fields, but calculations were also made for two

three-dimensional flow fields, to evaluate the effect of relieving

around a single jet. All the cases involved turbulent inflow and

finite-rate hydrogen-air chemistry. Comparisons were made with

empirical and analytical closed-form models to anchor the

results with an established experimental data base.

Accomplishment Description

The numerical simulations of the base pressurization designs

focused on providing flow visualization, which was used to

enhance the understanding of the flow physics that occurred

in the near-base region. The designs used supersonic slot or

circular orifice fuel injection, axial transpiration (bleed) from the

face of the step, vertical transpiration along the wall downstream

of the step, and a combination of supersonic slot from the step

lip and axial bleed beneath. The effects of fuel flow rate,

mainstream combustion, and variation in step height to combustor height ratio were also evaluated. The numerical results

predicted the expected flow features and gave insight into the

physics that determine the base pressure. The figure compares

flow fields that result for no base injection, showing the large

low-pressure recirculation region, to the flow fields for axial

base bleed with two mass flow rates (right), and to the mixed

supersonic/subsonic base injection case (lower left). Each

solution required 10 megawords of memory and between

6 (supersonic injection) and 20 (subsonic injection) Cray Y-MP

hours.

Significance

This study provides guidance for future design of base pressurization systems and indicates how the injection will affect the

local flow fields. Also demonstrated are some of the possible

limitations of pressurizing the base and the value of CFD

solutions as a tool for understanding the interaction of complex

flow physics.

Future Plans

The study will be enlarged to include evaluation of scramjet

combustor injection methods at high flight Mach numbers.

(a) (b)

(c) (d)

(a) Streamlines over pressure contours for flow over a two-dimensional backward-facing step in a scramjet combustor with a flight

Mach number of 16; blue = low pressure and red = high pressure. (b) Effect of bleeding hydrogen fuel from the base at a fuel equivalence ratio of 0.036. (c) Base bleed rate of 0.068. (d) Pressure and streamlines for a case with a Mach 2 supersonic fuel slot at the step

lip and base bleed below it at a fuel equivalence ratio of 0.05.

78 Aeronautics/Propulsion
Simulation of Scramjet Combustor Flow Field

Charles R. McClinton, Principal Investigator

Co-investigator: S. Srinivasan

NASA Langley Research Center

Research Objective

To simulate scramjet combustor flow fields using the General

Aerodynamic Simulation Program (GASP) flow solver and to

assess the combustor performance.

Approach

GASP, a state-of-the-art computational fluid dynamics (CFD)

code, was used to simulate high-speed scramjet combustor

flow fields for comparison with the experimental Planar LaserInduced Iodine Fluorescence Images (PLIIF) and with the results

obtained using the SPARK flow solver. The evaluation effort

was methodically completed by choosing several test cases to

encompass both the high- and Iow-enthalpy flow regimes. The

result for the Iow-enthalpy case is shown here.

Accomplishment Description

A simulation of normal injection of air at sonic conditions

(staged at the University of Virginia) into a Mach 2 low-enthalpy

airstream was performed. The mixing simulation took approximately 10 Cray Y-MP hours and about 6 million words of central

memory. Good agreement with the experimental PLIIF images

was obtained for both the low- and high-enthalpy cases. The

figure compares the CFD and PLIIF of the injected air mole

fraction contours along the plane of symmetry.

Significance

The GASP flow solver was successfully applied to simulate

high-speed combustor flow fields. The results obtained from the

simulation were compared with the experimental data, and they

yield confidence in the use of GASP for routine scramjet flowfield simulations.

Future Plans

The GASP flow solver will be used to simulate several highenthalpy scram jet combustor flow fields. Detailed comparisons

of the CFD results with Mie scattering images obtained experimentally will be performed.

Publication

Srinivasan, S.; Bittner, R. D.; and Bobskill, G. J.: Summary of

the GASP Code Application and Evaluation Effort for Scramjet

Combustor Flowfields. AIAA Paper 93-1973, June 1993.

Comparison of CFD (top) and PLIIF (bottom) injectant mole fraction contours along the centerline in the streamwise direction.

Aeronautics/Propulsion 79
Parallel Simulation of Unsteady Combustion

Suresh Menon, Principal Investigator

Co-investigator: Sisira Weeratunga

Georgia Institute of Technology/NASA Ames Research Center

Research Objective

To utilize the parallel code to simulate combustion instability in

ramjet engines and then to implement a dynamically adjusting

active secondary fuel-injection controller. The benefits and

problems associated with the use of highly parallel distributed

processing machines for unsteady simulations of combustion

will be determined.

Approach

A large-eddy simulation (LES) model for unsteady compressible

reacting flows was implemented on the Intel iPSC/860. The

explicit finite-volume numerical scheme solved the axisymmetric

form of the Navier-Stokes equations and a thin-flame model for

premixed combustion. A one-equation model for the subgrid

turbulent kinetic energy was used to model the unresolved

subgrid momentum and heat fluxes.

Accomplishment Description

It is possible to realize a highly scalable, distributed memory,

data parallel implementation for a class of explicit time integration schemes used in unsteady combustion simulations. A

geometry-tailored, two-level data partitioning strategy was

adopted to deal with the L-shaped computational space and has

resulted in a very efficient implementation, in spite of the added

complexity. The implementation is general enough to be

portable to any multiple-instruction/multiple-data parallel

computer that supports processor-to-processor communication

primitives. Parallel computers such as the Intel iPSC/860 appear

to provide a scalable, yet reasonably low-cost, alternative to

traditional vector supercomputers for simulations of unsteady

flow phenomena, except in the critical area of external input/

output performance. For combustion simulation with a oneequation subgrid model, 64 nodes of the iPSC/860 achieved

performance much better than a single-processor Cray Y-MP,

and, with 128 nodes, the performance exceeded a single

processor Cray C-90. A simulation with a 640 x 120 grid using

64 nodes required about 40 central processing unit hours to

obtain enough data for statistical analysis.

Significance

The parallel simulation model allows the exploration of unsteady

combustion in a parameter space that was uneconomical and

difficult to accomplish with the available resources on the Cray

computers. A new approach to simulate combustion in the

subgrid was developed and is uniquely suited for parallel

computations. Coupling the subgrid combustion model with the

parallel LES model will result in a new simulation capability.

Future Plans

Simulations of unsteady combustion with the subgrid combustion model will be carried out. A secondary fuel-injection

controller based on a theoretical model will be coupled in

parallel to the simulation model so that the controller will learn

while the simulation is going on and inject an optimized

secondary fuel stream to suppress the combustion instability.

Flame propagation in an axisymmetric ramjet: yellow = product, red = flame, cyan = fuel.

Vortex motion in an axisymmetric ramjet.

80 Aeronautics/Propulsion
Unsteady Combustion in a Ramjet

Suresh Menon, Principal Investigator

Co-investigator: Thomas M. Smith

Georgia Institute of Technology

Research Objective

To simulate and investigate unsteady combustion and combustion instability in a three-dimensional (3-D) ramjet engine and to

characterize the effects of the 3-D flow field on the premixed

flame structure.

Approach

An explicit, finite-volume scheme was used to solve the 3-D

unsteady, compressible Xavier-Stokes equations using a largeeddy simulation (LES) technique. A compressible eddy viscosity

was used to model the unresolved momentum and heat subgrid

fluxes. Premixed combustion was simulated using a thin-flame

model in which the turbulent flame speed explicitly appears as a

function of the laminar flame speed and the subgrid turbulent

kinetic energy.

Accomplishment Description

Unsteady combustion in a 3-D ramjet engine was simulated for

conditions that were studied earlier using axisymmetric and twodimenional (2-D) ramjet configurations. For stable, unsteady

combustion (characterized by low-amplitude, low-frequency

pressure oscillations), the vortex motion in the shear layer was

quite different in 3-D flows when compared to 2-D flows. In

3-D flows, conversion of the fuel into product was more efficient

due to 3-D mixing, and the vortices in the shear layer were

smaller and more concentrated due to vortex stretching effects.

In spite of these flow-dominated differences, the analysis of the

flame structure in 3-D flows showed that the flame sheet tended

to be predominately 2-D (cylindrically shaped). This result

agreed with data obtained by direct numerical simulation of

flame propagation in temporal mixing layers and homogeneous

isotropic turbulence, which showed that the flame sheet was

highly 2-D with the vorticity vector aligning with the intermediate strain direction while the flame sheet was normally aligned

with the most compressive strain direction. A simulation with a

grid resolution of 192 x 64 x 16 cells needed approximately

22 megawords of memory and required about 75 hours on the

Cray C-90 to obtain enough data for statistical analysis.

Significance

Statistical results suggest that the flame structure in complex

3-D flows has a preferred tendency to become a 2-D cylindrical

shape. This tendency suggests that a 2-D simulation model,

corrected for the out-of-plane vortex stretching effect, may be

able to predict the statistical properties of premixed flame as

in a 3-D flow. This ability would result in a major reduction in

computational resource requirements and perhaps make LES of

premixed flames a practical engineering tool.

Future Plans

The 3-D LES data base will be used to develop a new approach

to carry out 2-D LES of premixed flames that include the effects

of out-of-plane vortex stretching. Active control of combustion

instability in the 3-D ramjet will also be investigated using a

secondary fuel injection controller that was investigated earlier

in axisymmetric flows.

burnt flame fuel

Flame propagation in the combustion chamber of a ramjet at

various streamwise locations; yellow = product, red = flame, and

cyan = premixed fuel.

burnt flame fuel

Middle spanwise plane showing the flame structure and the

vortex motion (shown as contour lines) in the combustor.

Aeronautics/Propulsion 81
Three-Dimensional Parallel

John C. Otto, Principal Investigator

NASA Langley Research Center

Mixing Computations

Research Objective

To develop a parallel version of the three-dimensional chemically reacting, computational fluid dynamics, Navier-Stokes

code, SPARK, and to use the parallel code to examine a

nonreacting, fuel-mixing problem.

Approach

SPARK was moved from the sequential supercomputing

environment of the Cray, where it runs on a single processor,

to the distributed memory environment of the Intel iPSC/860

hypercube. The parallel efficiency of the code was measured

when the problem size was scaled with the number of processors and when the problem size was constant. Performance

was also compared to the sequential performance on a single

processor of the Cray Y-MP. SPARK was used to examine a

mixing enhancement strategy for scram jet engine applications.

This mixing involved adding a swirling component to the

supersonic fuel jet prior to injecting it at an angle of 30 degrees

into a Mach 2.0 flow in a rectangular duct. An algebraic eddyviscosity turbulence model was added to the code to model

the turbulence in the region where the fuel-air mixing occurs.

Mixing efficiency was computed as a function of axial distance

to determine if the swirl enhanced the fuel-air mixing. The

hydrogen fuel envisioned for use in scramjet engines was

modeled by helium in the experiment.

Future Plans

Work is under way to port the SPARK code to other parallel

computers, including the CM-5 and the Kendall square research

KSR1. The code will also run on the Intel Paragon machines to

solve very large problems that cannot be readily solved on

conventional supercomputers.

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Accomplishment Description

The parallel efficiency of SPARK asymptotes to a value of about

67 percent when a power-law viscosity model is used and the

global problem size scales with the number of processors. When

Wilke's mixture viscosity is used, the computational requirement

is significantly higher and the parallel efficiency is asymptotic

to about 76 percent. The performance of the code has been

compared to that of a single processor of the Cray Y-MP.

When the power-law viscosity model is used, the parallel code

operates at 95.3 MFLOPS on 32 processors and at 161.5 MFLOPS

on 64 processors. On a single processor of the Cray Y-MP, the

problem that was run on 32 processors of the hypercube ran at

a rate of 146.5 MFLOPS. The parallel code was used to study a

fuel mixing problem. Two swirl levels and a nonswirling case

have been examined. The mixing efficiency has been observed

to increase as the swirl level increases. For the swirling cases,

the mixing efficiency was significantly higher than for the

nonswirling case. Initial comparisons between the numerical

solutions and flow visualization data taken in the experiment

show the development of similar flow structures. All comparisons to this point have been qualitative.

Significance

The code efficiency has been shown to be fairly good and the

parallel code offers another resource for solving very large

problems. Swirl enhances the mixing levels of the fuel. Qualitative comparisons to experimental data show the development of

similar structures, but are insufficient to validate the numerical

solutions.

Parallel efficiency versus the hypercube order when the problem

size scales with the number of processors; squares = mixture

viscosity, circles = power-law viscosity.

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Mixing efficiency _'ersus axial distance from the duct inlet;

solid = no swirl, dotted = low swirl, dashed = high swirl.

82 Aeronautics/Propulsion
Launch Vehicle Base Flow Simulations

Thanh T. Phan, Principal Investigator

Co-investigator: Timothy J. Ventimiglia

General Dynamics, Space Systems Division

Research Objective

To develop an accurate and economical technology capable

of numerically simulating the complete flow field around a

multibody launch vehicle interacting with its multi-engine

plumes.

Approach

The three-dimensional Reynolds-averaged Navier-Stokes flow

solver, FALCON, was used to simulate the flow field over an

entire launch vehicle. The code employed a finite-volume,

cell-centered, upwind-differencing scheme that includes a

k-kl turbulence model. The GRIDGEN codes were used to

create the initial multiblock structure grid. As the solution

evolved, the adaptive version of the EAGLE code was used to

adapt the grid to the flow gradients.

Accomplishment Description

Perfect gas simulations were extended to include the effects of

the forebody for the Atlas II launch vehicle at several altitudes.

The results showed that the forebody significantly alters the flow

in the base region and therefore must be included to obtain a

realistic base flow solution. Flow gradient grid adaptation was

proved to greatly enhance the quality of the solutions, especially

at higher altitudes where the plumes expand to greater size

and create stronger interactions. A typical simulation required

approximately 46 megawords of memory and 90 Cray Y-MP

hours.

Significance

Currently, the environment in the base region of a launch

vehicle is predicted with very large uncertainties. These uncertainties lead to designs of thermal protection systems that include

many overconservatisms. This research will provide valuable

understanding and information which can help in the design

of optimal protection systems that can be used with higher

confidence.

Future Plans

Perfect-gas simulations for the Atlas II configuration, which

include four solid-rocket boosters, are planned. Real-gas effects

will also be modeled to capture plume interactions more

accurately.

Mach number contours in the base region at 54,000 feet;

free-stream Mach number = 2.1.

Temperature contours in the base region at 54,000 feet;

free-stream Mach number = 2. I.

Aeronautics/Propulsion 83
Flow in Turbine-Blade Coolant Passages

Tom I-P. Shih, Principal Investigator

Co-investigators: Mark A. Stephens and Kestutis C. Civinskas

Carnegie Mellon University/NASA Lewis Research Center

Research Objective

To develop and evaluate advanced computational tools to study

three-dimensional (3-D) flow and heat transfer inside coolant

passages of turbine blades.

Approach

When 3-D flow fields inside geometrically complex turbineblade coolant passages are computed, the hours involved in the

grid generation process account for most of the time required to

obtain a solution. To reduce grid generation time, overlapping

structured (Chimera) grids were employed with solutions to

the compressible Navier-Stokes equations obtained by the

OVERFLOW and PEGSUS codes.

Accomplishment Description

The OVERFLOW code was adapted to be used to compute flow

fields in coolant passages of turbine blades that have steady-state

solutions with respect to a rotating frame of reference. The

Chimera grid approach was ideally suited for the design and

analysis of coolant passages because the grid generation process

was reduced from about one month for a continuous multiblock

grid to a few days for a Chimera grid. Also, parametric studies to

optimize locations for pin fin and divider wall were performed

without generating new grids because grids about pin fins and

divider walls can be easily moved. Finally, because each overlapping grid was structured, very efficient relaxation schemes

were available to generate solutions. A typical 3-D computation

for the coolant passage required about 30 megawords of

memory and 3-5 Cray Y-MP hours.

Significance

Currently, coolant passages are designed by using either quasione-dimensional analyses or multidimensional analyses with

simplified geometries. This study produced a code that allows

the correct geometry to be analyzed in a timely and costeffective manner.

Future Plans

Accuracy of predictions will be improved by using turbulence

models that account for rotation and streamline curvature. The

physics of flow and heat transfer inside coolant passages will be

studied.

Publications

1. Steinthorsson, E.; Shih, T. I-P.; and Roelke, R. J.: Enhancing

Control of Grid Distribution in Algebraic Grid Generation.

Int. J. Numer. Methods Fluids, vol. 15, 1992, pp. 862-868.

2. Stephens, M. A.; Rimlinger, M. J.; Shih, T. I-P.; and Civinskas,

K. C.: Chimera Grids in the Simulation of Three-Dimensional

Flowfields in Turbine-Blade-Coolant Passages. AIAA

Paper 93-2559, June 1993.

Composite Plot of All Sixteen Meshes:_ q Plane

Two-dimensional view of a Chimera grid along the midplane of

a coolant passage.

Mach number contours along the midplane of a coolant passage.

84 Aeronautics/Propulsion
Jets-ln-Crossf[ow Mixing

Clifford E. Smith, Principal Investigator

Co-investigators: Daniel B. Bain and James D. Holdeman

CFD Research Corporation/NASA Lewis Research Center

Research Objective

The objectives of this work were to identify improved mixing

schemes for rich-burn/quick-mix/lean-burn combustors applicable to advanced aircraft engines and to assess the effects of

design parameters on mixing effectiveness. Efforts were focused

on jet mixing in rectangular cross-sectional geometries.

Approach

Numerical parametric studies were performed on different orifice

configurations. The computational analyses were performed

using a three-dimensional Navier-Stokes flow solver (CFD-ACF)

capable of analyzing turbulent reacting flows in complex

geometries. The numerical results were examined using an

interactive visualization package (CFD-VIEW).

Accomplishment Description

Over 60 orifice configurations were analyzed and assessed. The

modeled geometry consisted of a rectangular duct with a row of

jets on the top and bottom walls. Isothermal flow conditions

were assumed. Systematic, parametric analyses were performed,

consisting of variations in (1) orifice lateral alignment (in-line

and staggered), (2) jet-to-mainstream (J) momentum flux ratio

(16, 36, 64), (3) orifice aspect ratio (4:1,2:1,1:1, circle), (4) jetto-mainstream mass flow (MR) ratio (2.0, 0.50, 0.25), and

(5) orifice spacing-to-duct height (S/H) ratio (0.125 _<S/H <_ 1.5).

The numerical calculations were performed with grids varying in

size from 50,000-150,000 cells that required approximately

2-4 Cray Y-MP hours and 10-15 megawords of memory.

Significance

At optimum S/H, in-line lateral arrangements produced faster

initial mixing than non-impinging, staggered arrangements due

to their smaller geometric orifice size. Previous publications

show that S/H(_) is a practical design parameter for mixing

optimization even at high MR, and increasing J improves initial

mixing at optimum S/H.

Future Plans

The effect of mass flow ratio and aspect ratio on mixing effectiveness will be assessed, and emission characteristics will be

inferred from existing cold-flow data.

Publication

Bain, D. B.; Smith, C. E.; and Holdeman, J. D.: CFD Mixing

Analysis of Axially Opposed Rows of Jets Injected Into Confined

Crossflow. AIAA Paper 93-2044, June 1993.

Typical jet mixing numerical results. In-line circular orifices: S/H = 0.375, J = 36, MR = 2.0. Jet mass fraction: red = 0.0 to blue = 1.0.

Aeronautics/Propulsion 85
Flow Phenomena in Turbomachinery

C. S. Tan, Principal Investigator

Co-investigators: S. Arif Khalid and Ted Valkov

Massachusetts Institute of Technology

Research Objective

To investigate the role of compressor endwall flow in setting the

pressure rise capability in axial compressors, and to characterize

the unsteady flow phenomena involved in blade row interaction

in turbomachinery.

Approach

For compressor endwall flow, three-dimensional Reynoldsaveraged Navier-Stokes simulations of isolated compressor

blade passages were done using the VSTAGE finite-volume

compressible flow solver. Techniques were developed to

quantify the effect of clearance and loading on the endwall flow

field. For unsteady flow, direct Navier-Stokes simulations were

done for passage flows with and without upstream wakes at

Reynolds numbers 5,000-50,000. An algebraic turbulence

model was incorporated in the solver and the simulations were

repeated for Reynolds numbers of about 1,000,000. Results were

compared to those from the direct calculations to take into

account the uncertainty inherent to turbulence models.

Accomplishment Description

The compressor endwall flow computations were used to

develop an extension of the boundary layer displacement

thickness concept to quantify blockage in axial compressors.

The loss in the shear between the leakage and passage flow is

the source of blockage associated with the leakage vortex core.

Increasing the clearance substantially affects the sensitivity of the

endwall flow field to increases in passage pressure rise. Typical

computations used 8 megawords of memory and 6 Cray Y-MP

hours. The unsteady flow over the suction surface of the stator

blade was dominated by a moving row of vortical disturbances

produced at the leading edge upon interception of the wake.

The strength of the disturbances over the suction surface was

reduced by tailoring the pressure gradient over the foremost part

of the blade, and could be virtually eliminated by removing

boundary-layer fluid from the foremost part of the suction

surface. Typical direct simulation computations required

64 megawords of memory and 4 Cray Y-MP hours.

Significance

Flow in the endwall regions is known to have a critical influence

on axial compressor performance, and this study established

causal links between clearance and stall onset. Turbomachinery

design may be improved if the effects of unsteadiness on the

performance metrics are taken into account. Control strategies

identified in our study may be used to alleviate pressure

fluctuations.

Future Plans

We will assess the importance of different mechanisms of

blockage present in the endwall region. Casing treatment will

also be studied. Three-dimensional simulations of the unsteady

flow in wake-stator interaction will incorporate a full model of

the upstream wakes, including interaction with streamwise

vortices.

Particle traces in clearance vortex and contours of rotary total

pressure coefficient in the trailing-edge plane. Red = 0.0,

blue = -0.5.

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Disturbance vorticity contours showing the principal features of

the unsteady flow during wake-stator interaction.

86 Aeronautics/Propulsion ___.__--_
Hydrocarbon Scram jet Combustor Flows

Jong H. Wang, Principal Investigator

Rockwell International, North American Aircraft Division

Research Objective

To validate Navier-Stokes methodology for predicting hydrocarbon scram jet combustor flows.

Approach

The unified solution algorithm (USA) code is applied to existing

combustor-flow test cases to validate the code and provide

direction for future development of numerical and chemical

models.

Accomplishment Description

A hydrocarbon scram jet combustor was numerically investigated. The combustor was first modeled as a two-dimensional

(2-D) flow with the fuel-injector opening adjusted for a correct

fuel-air ratio. Several chemical kinetics models were investigated for the oxidation of ethylene fuel. For a fuel-air ratio

below 0.76, the predicted wall surface pressure was in good

agreement with experimental data. However, for a higher

fuel-air ratio, the predictions began to deviate from the data, and

in the main combustor the wall pressure was underpredicted by

50 percent. This deviation indicated that the mixing for the 2-D

numerical model was much less than that for the physical flow

model. To investigate the effects of three-dimensional (3-D) flow,

a 3-D numerical study was performed. The detailed configuration of the combustor, including all the fuel injectors, was

considered in the numerical model. The grid system consisted

of 347,030 points. The model was divided into two parts that

required .S and 14 megawords of memory, respectively. The

solutions took 125 Cray Y-MP seconds per iteration and the

entire process took about 130 hours to converge. The predicted

wall surface pressure for the 3-D model compared better with

data than the predicted pressure for the 2-D model. The maximum difference between the predictions and the data was about

20 percent. The discrepancy could have been due to the grid

resolution, chemistry, or turbulence models. Further study is

required to evaluate the effects of each parameter.

Significance

Efficient hypersonic combustors are essential to the design of airbreathing hypersonic vehicles. Flow physics complexity and

efficient mixing and combustion require an accurate and robust

computational tool.

Future Plans

Three-dimensional combustor flows with the effects of chemical

and turbulence models will be studied at various flow conditions.

(a)

(b)

Mach number contours for a hypersonic vehicle at Mach 8.

(a) Vehicle external flow Mach number contours. (b) Engine

internal flow temperature contours.

Aeronautics/Propulsion 87
Chimera Domain

Kurt F. Weber, Principal Investigator

Co-investigator: Dale W. Thoe

General Motors Corporation, Allison Gas Turbine Division

Research Objective

To port the advanced compressor analysis code COM3D to the

Intel iPSC/860 and evaluate its performance on a distributed

memory, multiple-instruction/multiple-data (MIMD) stream

architecture.

Decomposition Applied to Turbomachinery Flow

Approach

The purpose of this project was to apply the idea of domain

decomposition using Chimera grid embedding to couple the

calculated solutions of local flow fields in gas turbine fans and

compressors. The equations were solved using Pulliam's

diagonal version of the implicit approximate factorization

algorithm of Beam and Warming. A composite mesh for

complex flow fields was generated using the Chimera grid

embedding technique. The code worked with the output from

the multiple grid data management code PEGSUS, which

established the interpolation between individual grids. The flow

field was geometrically decomposed and grids were generated

for the individual elements of a gas turbine compression system.

PEGSUS was run to set up the interpolations for mesh boundary

communication. For parallel processing, data decomposition can

then be applied and the parallel processors distributed so the

solution on each mesh is calculated concurrently.

Accomplishment Description

The best approach for applying a parallel Chimera code

to turbomachinery was to base it on the parallel code,

POVERFLOW. The POVERFLOW host program, the main

program for the nodes, and the coding to update the Chimera

boundary points were documented. Ten 32-node hours on the

Intel iPSC/860 were used to calculate a 2-grid simulation of a

GMA 3014 wing-nacelle configuration with the nacelle grid

embedded in the wing O-grid. Learning and documenting the

code and completing an initial calculation were necessary

before upgrading the code to calculate turbomachinery flow.

Significance

The use of three-dimensional flow codes for aerodynamic design

improvements in aircraft gas turbine engine compressors and

fans has become more widespread. Increased emphasis is on

extending the capability of analyses by coupling previously

isolated flow-field calculations to determine the effect of

coupling on component performance. Cost-effective parallel

computers allow the use of complex turbomachinery flow

calculations in the routine design processes.

Future Plans

POVERFLOW will be modified for turbomachinery calculations.

Multiple grid parallel capability for turbomachinery flow can

best be acquired by modifying the Chimera code, as opposed to

porting our in-house serial flow code.

Static pressure contours and particle traces for a wing and simplified GMA 3014 nacelle configuration.

88 Aeronautics/Propulsion
Turbofan Compression System Simulation Using Chimera Domain Decomposition

Kurt F. Weber, Principal Investigator

Co-investigator: Dale W. Thoe

General Motors Corporation, Allison Gas Turbine Division

Research Objective

To perform validation studies of a Navier-Stokes code that

simulates the flow through the low-pressure compression system

of a turbofan engine. Special emphasis was placed on the

simulation of the flow field for the GMA 3007 engine.

Approach

A fully viscous three-dimensional (3-D) finite difference code

was applied to the solution of compressor and fan flow fields.

The equations were solved using Pulliam's diagonal version of

the Beam-Warming implicit approximate factorization algorithm. Flow-field data were interpolated between grids using the

Chimera grid embedding technique. The code worked with the

output from the multiple grid data management code PEGSUS.

Accomplishment Description

A 3-D embedded grid Navier-Stokes code was applied to

simulate transonic compressor flows. Fully viscous solutions

with a very fine H-grid embedded in the clearance region were

calculated for several designs of the low-pressure compression

system of the GMA 3007 turbofan engine. Calculations included

the flow through the fan rotor downstream through the flow

splitter in the core and bypass ducts. The solutions show good

agreement with component rig test data. The figure shows static

pressure contours on the hub, fan rotor, and splitter surfaces.

This multiple grid composite mesh used five individual grids. A

typical calculation used 18 megawords of storage and required

18 Cray-2 hours for 2,500 iterations.

Significance

The use of 3-D flow codes for aerodynamic design improvements in aircraft gas turbine engine compressors and fans has

become more widespread. However, the level of accuracy has

not been high enough to reduce significantly the amount of

testing required. Major improvements in performance or

significant reductions in cost are unlikely unless the 3-D

analyses focus on adequate resolution of the flow field and

accuracy of the calculations. Accurate modeling of the flow

field within compressor blade rows will provide the information

needed to improve compressor performance.

Static pressure contours for the GMA 3007 fan with a flow-field

splitter; blue = low pressure, red and magenta = high pressure.

Future Plans

An analysis for turbomachinery flow using Chimera domain

decomposition that will run on massively parallel systems will

be developed.

Aeronautics/Propulsion 89
Multifunction Propulsion 'Systems

Perry A. Wooden, Principal Investigator

Co-investigator: Timothy P. Nobel

LTV, Aircraft Division

Research Objective

To investigate the performance of multifunction propulsion

systems including cascade array thrust reversers, pivot-door

thrust reversers, and long-cowl mixed flow nozzles in order to

isolate a suitable technique for accurately predicting component

performance.

Approach

A subscale model of a pivot-door thrust-reversing system was

analyzed using computational fluid dynamics (CFD) methods

and statically tested in a wind tunnel. The test hardware was a

40-percent-scale representation of an 18 degree sector of a

high-bypass-ratio turbofan engine. The model instrumentation

provided data comparisons with CFD predictions. Contiguous

block interfacing eliminated linear interpolation errors.

Accomplishment Description

A full three-dimensional pivot-door analysis using PARC3D

was completed. The figure shows a case with the pivot-door

deployed at a 30 degree angle measured from the vertical

position. Inflow conditions were set to an average total pressure

and temperature was obtained during testing. The outflow

condition was set to ambient, sea level condition at zero

velocity. Particle traces were initiated at several positions along

the fan duct. Large amounts of separated flow existed at the

juncture of the lower fan duct wall and the fully deployed pivotdoor. The particles released along the centerline traveled the

complete door length, but the off-centerline traces revealed

spanwise flow over the external portion of the door. The three

GRIDGEN blocks generated for this problem were assembled

into a single block to eliminate input/output and take advantage

of the computer memory. The single block contained 1.3 million

grid points requiring 32 megawords of memory. At a Courant

number of 0.8, the solution took _5,000 iterations to converge

four orders of magnitude. The computations required

40 Cray Y-MP hours to complete.

Significance

Thrust reversers represent the most common multifunction

propulsion system application that enhances the overall effectiveness of both commercial and military aircraft. The pivot-door

thrust-reverser concept is relatively new and offers weight

benefits and reduced complexity.

Future Plans

Navier-Stokes simulations on the pivot-door configuration will

be examined. Navier-Stokes calculations on a cascade array

thrust reverser design will begin. The multifunction propulsion

design investigation will be expanded to include a mixed-flownozzle experiment and computations.

Particle traces for a 30 degree pivot-door Navier-Stokes solution

with an inlet Mach number of 0.786.

90 Aeronautics/Propulsion
Numerical Study of Thrust Vectoring Characteristics

David T. Yeh, Principal Investigator

Rockwell International, North American Aircraft Division

Research Objective

To extend and validate the Navier-Stokes methodology to

predict the thrust-vectoring flow phenomena and to provide an

assessment of turning efficiency and effectiveness as functions

of paddle deflection and engine power settings.

Approach

The unified solution algorithm (USA) code was used to simulate

the thrust-vectoring flow characteristics. The USA code solves

the three-dimensional Reynolds-averaged Navier-Stokes

equations through an upwind, flux-difference, finite-volume

approach. A high-resolution total-variation-diminishing (TVD)

scheme is incorporated to achieve a high-order spatial accuracy

while maintaining numerical stability. The X-31 thrust-vectoring

system is the focus of this study. This configuration was chosen

because of the existence of a ground-test data base and the

ongoing flight-test program.

Accomplishment Description

A blocked-grid approach was exploited to model the X-31 thrust

vectoring system. A two-dimensional algebraic grid generation

technique was utilized at each cutting surface to model the

paddles with multiple deflection settings. The first figure

illustrates a typical cross-sectional grid on a cutting surface,

where a rectangular grid was created at the center to avoid a

potential convergence problem associated with a singular

axis. The study of the thrust deflection effectiveness was

accomplished by sweeping one of the paddles into and out of

the jet at constant deflection angles of the other two paddles.

Data for the F404-GE-400 was used to define the boundary

conditions at the nozzle throat for a prescribed power setting.

Numerical solutions revealed the formation of diamond wave

patterns throughout the exhaust jet. Numerical results show

that turning efficiency decreases as paddle deflection increases,

which is a result of stronger shocks at larger deflections. In the

second figure, the surface pressure contours and grid for the

X-31 thrust-vectoring system are illustrated. The top paddle was

deflected 20 degrees into the jet while the other two paddles

stayed at their neutral positions. A typical computation for

this configuration requires 20 megawords of memory and

25 Cray Y-MP hours.

Significance

The numerical study parallels the ongoing X-31 flight-test

program and establishes the computational fluid dynamics

capability for analysis of thrust-vectoring flow phenomena.

This study will have significant impact on current and future

advanced aircraft design.

Typical grid arrangement on a cutting surface.

Pressure contours for the X-31 thrust-vectoring system. Contour

levels are between 0.4 and 3.0.

Future Plans

The control effectiveness of the thrust-vectoring system varies

with altitude due to different nozzle pressure ratios. Numerical

simulation will be performed at the flight conditions to provide

a viable assistance to the in-flight control command.

Aeronautics/Propulsion 91
National Aero-Space Plane External

Shaye Yungster, Principal Investigator

ICOMP/NASA Lewis Research Center

Research Objectives

To provide a means for determining the sub-scale and full-scale

performance of external burning nozzles at transonic flight

conditions.

Approach

The Reynolds-averaged Navier-Stokes equations with finite-rate

chemistry were solved using two- and three-dimensional (2-D

and 3-D) multiblock, fully implicit codes based on the LU-SSOR

factorization scheme. The spatial discretization was based on a

second-order total variation diminishing scheme for the 2-D

code, and central-differences with second- and fourth-order

artificial damping for the 3-D code. The Baldwin-Lomax

algebraic turbulence model was used in all the calculations.

Most of the computations were carried out using a simple onestep combustion model.

Accomplishment Description

A computational model for the external fuel injection and

combustion processes was developed and applied to three

nozzle configurations: (1) a baseline cowl without a flame

holder; (2) an extended cowl without a flame holder, and (3) an

extended cowl with a flame holder. The numerical results were

compared with experiments. The computed pressure distributions along the expansion ramp agreed favorably with the

experimental data. Qualitative comparisons with experimental

infrared images, such as those in the figures, were also conducted. This study showed that nozzle flows that include

external burning can be adequately modeled using a 2-D

formulation. Using this simplified approach, different configurations were analyzed quickly and inexpensively. A typical

calculation required about 8 megawords of memory and 2 Cray

Y-MP hours. At transonic speeds and without external burning

3-D effects become important because of nozzle overexpansion;

thus a full 3-D analysis is required. For these cases, a multiblock

version of the RPLUS3D code was used.

Significance

Nozzles similar to those on the National Aero-Space Plane are

highly overexpanded at transonic speeds, resulting in large drag

increases. One way to eliminate this drag is to increase the

pressure along the afterbody by external combustion. Numerical

simulations can provide details of the flow field, identify wind

tunnel interference effects, determine scaling effects, and

analyze advanced designs.

Burning Nozzle Studies

Future Plans

Determination of the full-scale performance of external burning

nozzles at transonic flight conditions and investigation into the

effects of various parameters on the thrust and normal forces

acting on the expansion ramp and cowl are being investigated.

Experimental infrared image of external burning.

Computational results showing temperature contours:

Mach = 1.2, nozzle pressure ratio = 6, external burning fuel

pressure = 320 pounds per square inch.

92 Aeronautics/Propulsion


Aerodynamics and Acoustics of Rotorcraft

W. J. McCroskey, Principal Investigator

Co-investigators: J. D. Baeder, R. L. Meakin, V. Raghavan, and G. R. Srinivasan

NASA Ames Research Center

Research Objective

To develop and validate accurate, user-oriented viscous

computational fluid dynamics codes for three-dimensional (3-D)

unsteady aerodynamic flows about arbitrary rotorcraft configurations. Complex vortical wakes, shock waves, rotor-body

interactions, and acoustic waves are included.

Approach

Advanced 3-D unsteady multizone implicit Navier-Stokes codes

are used to simulate rotorcraft aerodynamics and acoustics under

flight conditions that have not previously been treated satisfactorily. High accuracy and stability are achieved for rotor blades by

using the transonic unsteady-rotor Navier-Stokes code, which

features full upwinding, enhanced accuracy, and high computational efficiency. Complete rotorcraft are treated using the

OVERFLOW code to solve the unsteady thin-layer Navier-Stokes

equations and the DCF3D code (a new domain connectivity

algorithm) to dynamically establish domain connectivity among

the systems of fixed and rotating overset grids. In addition,

special solution-adaptive grid clustering and wave fitting

techniques are used to capture low-level radiating acoustic

waves.

Accomplishment Description

A hypothetical, but realistic, set of flight conditions for the

V-22 Osprey aircraft was established to facilitate rigorous testing

of DCE3D, and to carry out an overset grid proof-of-concept

tiltrotor simulation. Relative motion and interference effects

between the V-22 airframe and rotor-blades were directly

simulated within the context of an unsteady, thin-layer

Navier-Stokes computation (see figure). The domain connectivity algorithm was verified to perform at rates equal to or greater

than those realized previously for store-separation-like applications. The feasibility of carrying out unsteady Navier-Stokes

analyses of rotorcraft problems has been demonstrated. The

noise from five different rotor blades was examined in hover for

a wide range of tip Mach numbers, including a detailed analysis

of the nonlinearities that plague conventional acoustic theories.

This study provided useful insight into the noise-generating

mechanisms of the various blade shapes and the near-field wave

front formations, and helped to define the limitations of the

linear acoustic analogy.

Significance

The ability to simulate accurately the aerodynamics and

acoustics of rotorcraft will allow quieter and more efficient

vehicles to be designed at lower cost and less risk. Analysis of

the flow separation on the V-22 Osprey aircraft has helped to

avoid costly design errors.

Future Plans

More accurate and efficient simulations of rotorcraft aerodynamics and acoustics, with emphasis on blade-vortex interaction

noise and rotor-body interference on tiltrotor aircraft will be

developed.

Unsteady simulation of the V-22 Osprey tiltrotor aircraft aerodynamics in low-speed forward flight (helicopter mode). Small matrices

of particles were released aft of an inboard and outboard blade-tip in a post-process flow visualization computation based on flowfield data from every 15th time step of the simulation. The particles are colored by time and highlight the structure of the rotor wake.

pAGE_._._4 INTENTIONALLYBL_,NK

Aeronautics/Rotorcraft 95
Calculations of High-Performance Rotorcraft

W. J. McCroskey, Principal Investigator

Co-investigators: J. D. Baeder, E. P. N. Duque, S. K. Stanaway, and G. R. Srinivasan

NASA Ames Research Center

Research Objective

To compute the viscous three-dimensional (3-D), unsteady flow

field around advanced helicopters. Particular emphasis is placed

on the transonic aerodynamics of high-performance rotor-blade

tips, vortical wake structure, and rotor-body aerodynamics

interference.

Approach

The unsteady 3-D Euler/Reynolds-averaged Navier-Stokes

equations were solved by structured- and unstructured-grid

methods, either in blade-fixed or inertial reference systems. The

codes were validated through detailed comparisons with

experimental data.

Accomplishment Description

The Reynolds-averaged thin-layer Navier-Stokes code

OVERFLOW was applied to the Comanche fuselage. The first

figure illustrates the surface pressure distribution over the front

part of the aircraft. Areas in gray were not included in the

calculation, thus they have no influence on the computed flow

field. Significant progress was made in coupling an unstructuredgrid rotor code to a structured-grid rotor code. In this case, a

structured grid optimized for viscous-dominated flows is used for

the near-body flows, while the unstructured grid is used for the

inviscid flow fields away from the body. The two grid systems

overlap and interpolate flow quantities at their boundaries. The

second figure illustrates the structured/unstructured-grid system

for a rotor in hover.

Significance

The Comanche helicopter is the Army's newest aircraft. Viscous

fuselage calculations provide detailed flow-field information

unattainable by other methods. These calculations can be used

to help evaluate and improve aircraft performance. The hybrid

structured/unstructured method incorporates the best features of

each individual technology. Generally, structured grids perform

much better for near body flows because the grid can cluster

near the walls without severe time-step limits. The unstructured

grids can refine around flow features more readily than the

structured grids. This feature potentially allows more accurate

wake calculations.

Surface pressure maps for the Comanche helicopter free stream;

Mach = 0.26, Reynolds number = 14,000,000, angle of attack =

0.0.

8-

Unstructured/structured hovering rotor grid system.

Future Plans

The unsteady flow field of the Comanche, including the rotor

hub and blades in forward flight, will be computed and compared with future experiments. The structured/unstructured

methodology will be validated for hover and forward flight and

used to improve Chimera grid boundary update and connectivity. Dynamic grid adaption of the unstructured mesh will be

implemented and demonstrated.

96 Aerona utics/Rotorcraft
Rotorcraft Drag Prediction

J. C. Narramore, Principal Investigator

Co-investigators: A. G. Brand, J. J. Shillings, and D. W. Axley

Bell Helicopter Textron, Inc.

Research Objective

To develop the capability to predict forces and moments

including drag on complete rotorcraft vehicles.

Approach

Three-dimensional Navier-Stokes codes were used to perform

steady and unsteady analyses of rotor blades and fuselages in

flight and correlate these results with existing wind tunnel data.

Both code and grid topology investigations were carried out.

Accomplishment Description

Research pertaining to the calculation of forces on threedimensional rotor blades, fuselages, and other components, with

special emphasis on the calculation of drag levels, was performed. Work was focused in four areas: developing an actuator

disk model, modeling inlet flows, Navier-Stokes solutions to the

flow about anhedral rotor tip shapes, and helicopter fuselage

drag computations. To acquire methods to design advanced antitorque devices for helicopters, a ducted tail rotor actuator disk

model based on Navier-Stokes techniques was developed. A

typical solution required 8 Cray Y-MP hours and 10 megawords

of memory. The figure illustrates the computed flow. Thrust was

generated for this zero-angle-of-attack case moving from the port

side of the disk to the starboard side aft of the disk (shown by the

particle traces).

Significance

Computational fluid dynamic methods offer high potential for

improved rotorcraft design and performance capabilities. This

new technology should reduce vehicle costs, improve performance, reduce development times, increase aerodynamic

responsiveness, and improve the quality of rotorcraft vehicles.

Requirements for modern high-speed helicopters include fast

rates of climb, high degrees of maneuverability, and improved

performance. Accurate assessment of blade and fuselage drag in

maneuvers, transonic flight, or at high angle of attack is crucial

to rotorcraft predictive and design methodology. Results from

detailed modeling of a ducted fan concept indicate that

Navier-Stokes methods may be useful for the design of

advanced helicopter antitorque devices. However, because

of slow convergence at low Mach numbers, an incompressible

Navier-Stokes methodology is also being pursued.

Future Plans

Chimera grid techniques will be incorporated into these

modeling techniques. Wind tunnel data, including pressure

and force comparisons, will be correlated for an incompressible

Navier-Stokes method. Navier-Stokes codes will be used to

investigate unsteady flows in a duct. Drag increments due to

fuselage shape perturbations will also be computed.

Thrusting actuator disk inside a faired duct with 6.8 pounds per

square foot disk loading.

Aeronautics/Rotorcraft 97
Dynamic Adaption for Three-Dimensional

Roger C. Strawn, Principal Investigator

Co-investigators: Rupak Biswas and Michael Garceau

RIACS/Stanford University

Research Objectives

To solve problems in rotary-wing aerodynamics where rotorwake systems are extremely complex. Unlike a fixed wing,

where the wake is rapidly convected downstream, a rotor

operates in close proximity to its own wake. In order to accurately model these effects, computational methods must minimize any numerical dissipation that can artificially diffuse the

vortices in the wake.

Unstructured Grids

Approach

Numerical dissipation of the wake was minimized by using

dynamic solution-adaptive grids and flow solvers that were

appropriate for the resulting unstructured meshes. The solutionadaptive approach adds and removes points locally to provide

higher resolution for moving flow features such as rotor-wake

vortices. This local increase in grid resolution provides an

excellent mechanism for minimizing numerical dissipation as

long as the mesh points can be efficiently added and deleted

from the mesh.

Mesh adaption for a 3-D NACA 0012 wing with inviscid

sidewalls, M_ = 0.85, _ = 1.0 degree.

Accomplishment Description

An efficient procedure was developed for the simultaneous

coarsening and refinement of three-dimensional (3-D) unstructured tetrahedral meshes. The mesh-adaption algorithm was

implemented in C and used an innovative data structure

consisting of a series of dynamically allocated linked lists.

These lists allowed the mesh connectivity to be reconstructed

rapidly when individual points were added or deleted. The

data structure was based on edges of the mesh rather than the

tetrahedral elements themselves and resulted in an efficient

data structure and facilitated anisotropic mesh refinement and

coarsening.

Significance of Accomplishment

The key to success for dynamic mesh adaption is the ability to

add and delete points from the grid efficiently. For an unsteady

flow, this coarsening/refinement step must be completed every

few time steps, so its efficiency must be comparable to that of

the flow solver. It must also have a reasonable memory requirement. The new mesh-adaption scheme has all these attributes.

In addition, it can be applied to a variety of important NASA

problems in areas other than rotorcraft aerodynamics.

Side view of the final mesh after three refinement and two

coarsening levels.

Future Plans

The new mesh-adaption scheme is currently being used to

compute large problems in rotor aerodynamics and acoustics.

It is also being modified to run concurrently with the flow solver

on the CM-5 parallel computer.

Publication

Biswas, R.; and Strawn, R. C.: A New Procedure for Dynamic

Adaption of Three-Dimensional Unstructured Grids. AIAA 31 st

Aerospace Sciences Meeting, Reno, Nev., AIAA Paper 93-0672,

Jan. 1993.

98 Aeronautics/Rotorc raft
Simulation of a Complex Three-Dimensional

H. Tadghighi, Principal Investigator

Co-investigator: A. A. Hassan

McDonnell Douglas Corporation

Research Objective

To develop a predictive capability to analyze accurately and

optimize efficiently complex internal flows for rotorcraft

advanced antitorque systems.

Approach

The three-dimensional, Reynolds-averaged, zonal Navier-Stokes

solver used for this study was NASTD. The solver algorithm is a

time-dependent finite-volume implicit scheme based on upwind

differencing. Baldwin-Lomax and k-_ turbulence models were

employed in this study. To examine the complex internal flow

created by the various components of the no-tail-rotor (NOTAR)

concept antitorque system (such as highly curved inlet geometry,

fairing, centerbody, and multibladed high-speed fan), analysis

was performed for a range of mass flow rates with and without

the influence of the swirl induced into the flow by the fan. The

accuracy of the flow-field analysis was assessed using measured

data.

Accomplishment Description

Several mass flow rates were evaluated using an actuator disk

model to represent the existence of the fan. Analyses were

performed with two turbulence models to investigate their

effectiveness in improving the numerical simulation of the

NOTAR antitorque system internal flow. Comparisons between

measured data and the predictions indicated good agreement for

a range of mass flow rates. The solution using the k-£ turbulence

model improved the prediction of the flow-field features by more

accurately capturing the separated flow regions observed in the

measured data. The inlet-duct centerbody-fairing grid was

comprised of 3 grid zones and contained approximately

0.7 million points. A typical solution with this mesh required

48 megawords of memory and up to 15 Cray-2 hours.

Internal Flow Field

Future Plans

The NOTAR antitorque system internal flow will continue to be

optimized by reducing interference effects between the inletfairing centerbody-fan components and optimizing duct surface

contours. To reduce the intensive processing requirement for the

k-£ model, the Baldwin-Barth one-equation turbulence model

will be investigated for computational efficiency.

c_

l O 822

io 140

NOTAR antitorque system internal flow prediction at

Mach = 0.2.

Significance

Accurate analysis of the complex flow field associated with the

internal flow using the NASTD solver lends confidence in relying

on a computational fluid dynamics technique for applications to

problems of a similar nature. This provides valuable insight into

the details of the flow and will furnish the designer with a viable

tool to enhance the flow characteristics and the performance

efficiency of the NOTAR antitorque system. The capability of

NASTD to capture the features of the system's complex internal

flow field is encouraging.

Aeronautics/Rotorcrafl 99



Transition in a Highly Disturbed Environment

David E. Ashpis, Principal Investigator

Co-investigator: Philippe R. Spalart

NASA Lewis Research Center/Boeing Commercial Airplane Group

Research Objective

To understand the physical mechanisms of nonlinear stability

and bypass transitions to provide guidance for engineering

model developers.

Approach

Direct numerical simulations of controlled disturbances in the

boundary layer were performed. Spatial simulations were

performed using a Spalart's "fringe code." Numerical experiments simulating single frequency excitation (vibrating ribbon)

and broad band excitation (blowing pulse) were performed. The

disturbances were introduced on the wall by specifying localized suction/blowing. The flow field, frequency, and wavenumber spectra were computed for incompressible Blasius

boundary layer flow.

Accomplishment Description

The original fringe code was modified to allow specification of

time-dependent wall boundary conditions. Series of two- and

three-dimensional runs were performed for a range of disturbance frequencies, amplitudes, and Reynolds numbers.

Emphasis was on moderate and high levels of disturbances

in the subcritical region. Computing resources varied according

to the domain size, grid resolutions, and the length of simulation. Typical runs required 4-32 megawords of memory and

4-20 Cray Y-MP hours. Observations included some elongated

streamwise structures in the pulse case, and a peculiar multiperiodic behavior in the low-frequency ribbon excitation case.

Significance

A highly disturbed environment prevails in flow in the gas

turbine (for example, at the entrance to the turbine). Hence,

conclusions obtained from the simulations of the associated

nonlinear and bypass stability mechanisms are valuable for

developers of models for applied engineering calculations.

Future Plans

Post-processing of the vast data acquired will be completed.

Effects of factors such as pressure gradients and curvature will be

studied. The studies will be extended to flows with heat transfer. Flow structures generated by a blowing pulse in a Blasius

boundary layer. Shown from top to bottom are u, v, w disturbance velocities in a horizontal x-z plane. Pulse is generated at

a Reynolds number of 68,000. Domain range is from Reynolds

numbers 30,000 to 305,000.

Aeronautics/Turbulence 103
High-Speed Turbulent Reacting Flows

J. Philip Drummond, Principal Investigator

Co-investigators: Peyman Givi, Cyrus K. Madnia, Steven H. Frankel, and Virgil Adumitroaie

NASA Langley Research Center/State University of New York at Buffalo

Research Objective

To develop subgrid scale models for large-eddy simulations

(LES) of turbulent reacting flows to make accurate predictions in

practical engineering flows.

Approach

The investigation was based on probability density function

(PDF) methods and results obtained from direct numerical

simulations (DNS) were used for model validation. Computations of both incompressible and compressible flows in two- and

three-dimensions were performed in homogeneous isotropic

turbulent flows and spatially developing mixing layers. The

numerical schemes were based on pseudospectral and hybrid

spectral/finite-difference algorithms. The DNS-generated data

were used in both a priori and a posteriori modes to validate

LES predictions.

Future Plans

Improved models for mixing and reacting in turbulent flows are

currently being developed for use as subgrid scale models in

LES. The use of such models in compressible reacting flows is

also under investigation.

Accomplishment Description

Because of the demonstrated capabilities of PDF closures in

homogeneous flows, their use was suggested as a subgrid scale

model in LES of reacting flows. The subgrid scale models

developed involved the use of a Smagorinsky-based, oneequation hydrodynamics model and an assumed Dirichlet PDF

for the thermochemistry. The procedure required the solution of

additional modeled transport equations for the subgrid turbulent

kinetic energy and the subgrid species covariance. Results were

obtained for both two-dimensional (2-D) and three-dimensional

homogeneous turbulence and a 2-D mixing layer. A typical

result for the shear layer is shown in the figure, which depicts

contour plots of the chemical product formed in the reaction

A + B --_ P. The results suggested that better models were needed

for the subgrid moments and for the formation of the joint scalar

PDF. DNS requires 25 Cray Y-MP hours and 23 megawords of

memory, whereas LES requires on the order of minutes of

Cray Y-MP time.

Significance

DNS of high-Reynolds-number engineering flows with complex

chemical kinetics is not possible in the foreseeable future.

LES and PDF modeling offer an attractive, economically and

computationally feasible means of making accurate engineering

predictions. This work is one of the first attempts at incorporating

PDF methods for LES of turbulent reacting flows.

(a)

(b)

Contour plots of product mass fraction from (a) LES and

(b) filtered DNS results.

104 Aeronautics/Turbulence
Simulation of a Turbine-Blade Cooling Passage

Kai-Hsiung Kao, Principal Investigator

Co-investigator: Meng-Sing Liou

ICOMP/NASA Lewis Research Center

Research Objective

To predict the flow-field and heat transfer coefficients in a

branched-duct test section that includes several significant

features of a turbine blade cooling passage. Higher engine

efficiency can be obtained at higher operating temperatures.

Higher turbine inlet temperatures often require internal cooling

in the blades and vanes. Effective design of internal cooling

passages is greatly enhanced by the availability of dependable,

validated computer codes.

Approach

A computer code based on a time-accurate, three-dimensional

(3-D), finite-volume, high-resolution scheme for solving the

compressible full Navier-Stokes equations was developed. The

numerical formulation used a new advection upstream splitting

method that improved the accuracy and efficiency of the flow

solver. A 3-D Chimera grid-embedding technique was incorporated with the Navier-Stokes solver, subdividing into regions the

physical domain of the flow field to accommodate an easily

generated grid. The strategy from Chimera overset grids provides

great flexibility in grid generation for complex configurations.

Accomplishment Description

The present Navier-Stokes solver was developed and validated

for various two- and three-dimensional complex configurations.

For the turbine blade cooling passage problem, a primary coarse

grid system was generated using 18 structured grid blocks.

Different grid topologies were used to resolve specific geometries. Because the overset meshes used were independent of

each other, they could readily be embedded in arbitrary

orientations. Consequently, grid points within solid boundaries

were blanked or excluded from other overset grids. Communication through boundary interfaces between the separate grids was

carried out using trilinear interpolation. The grid used for the

simulation had about 160,000 points. Preliminary results clearly

displayed the basic features, including the recirculation and

separation regions. Typical runs required 3.5 megawords of

memory and 3.7 x 10 -5 Cray Y-MP seconds per grid point for

each iteration.

Significance

This simulation provided insight into the physics of flow

structures and examined heat transfer behaviors in a turbine

blade cooling passage. The results enhance the accuracy of the

design process and reduce the cost of experimental tests.

Future Plans

A finer mesh system is required to enhance the accuracy of the

simulation. Various turbulence models will be considered to

predict turbulent flow regions. Numerical results and experimental data for the velocity field and surface heat transfer coefficients

will be compared.

Temperature distributions in a turbine blade cooling passage with M_ = 0.4, Reynolds number = 45,000. Maximum temperature

(white) = 0.75 K, minimum temperature (blue) = 0.65 K.

Aeronautics/Turbulence 105
Numerical Simulation of Turbulence

J. J. Kim, Principal Investigator

Co-investigators: Robert D. Moser, R. S. Rogallo, Nagi N. Mansour, and J. R. Chasnov

NASA Ames Research Center/Center for Turbulence Research

Research Objective

To investigate the physics of turbulence in simple model flow

situations and provide otherwise inaccessible turbulence data for

use in modeling and other studies.

Approach

Direct numerical simulations and large-eddy simulations of

turbulence in simple flow situations (currently homogeneous

turbulence and turbulence in a plane channel) were performed

by solving the incompressible Navier-Stokes equations using

spectral numerical methods. The flow solutions were processed

to extract relevant statistical data and to answer questions

regarding the physics of turbulence.

Accomplishment Description

Several direct numerical simulations of homogeneous turbulence

were performed. The decay of isotropic turbulence was simulated to determine the Reynolds number dependence of the

decay power law. The results modified the k-_ turbulence

model to reflect the observed dependence. Isotropic turbulence

simulations were used to obtain Lagrangian velocity correlations.

Large-eddy and direct simulations of homogeneous turbulence

were used to study the convection of scalars, buoyancy driven

flows, and simple combustion processes. These simulations were

used to study self-similar turbulence decay at high Reynolds

numbers, the "collapse" of eddies in stratified flows, and the

behavior of probability density functions in reacting flows. Largeeddy simulations were used to address behavior of the largest

scales in turbulent flows. Finally, simulations of nonisotropic

turbulence were performed to provide better resolution in these

cases. Two direct simulations of plane channel flow were

performed. First, a channel at a Reynolds number of about 6,600

was simulated to study scalar transport in wall-bounded flows.

Second, simulation of a channel at a Reynolds number of

approximately 13,000 was begun.

Future Plans

Simulations to provide turbulence data at higher Reynolds

numbers and in new flow situations are planned. New algorithmic developments for parallel machines made as part of this

project and the new Paragon hardware will increase our

capability to study these important flows.

.6

.4

S

.2

. e.

._ee _ DNS data • Experimental

I I I I I I

5 10 15 20 25 30

Rk

Velocity-derivative skewing; Reynolds number versus experimental data.

Significance

In many flow situations turbulence is the controlling phenomenon, yet efforts to predict or control the effects of turbulence

have had only limited success. These simulations provide

important information needed to improve our understanding of

and our ability to predict and control turbulence. Results from

these simulations are being used to develop improved turbulence models and control strategies.

106 Aeronautics/Turbulence
Receptivity, Transition, and Turbulence Phenomena

Linda D. Kral, Principal Investigator

Co-investigators: William W. Bower and John F. Donovan

McDonnell Aircraft Company

Research Objective

To investigate the physical processes in fluids through simulations of transition mechanisms or turbulence phenomena

occurring in incompressible to supersonic flows.

Approach

Three areas were investigated: (1) boundary layer receptivity,

(2) boundary layer transition, and (3) compressibility effects on

turbulent boundary layers. To study boundary layer receptivity,

the linearized Navier-Stokes equations were recast in spectral

form and solved. For the studies in boundary layer transition and

turbulence, direct and large-eddy simulation techniques were

used. Different subgrid-scale models were investigated for the

simulation of both transitional and turbulent boundary layers.

Future Plans

The physics of the compressible, boundary layer transition

process using direct numerical simulations and large-eddy

simulations with the dynamic subgrid-scale model are being

studied. Boundary layer transition over a curved surface will be

simulated. The influence of localized wall heating to control

transition is also under investigation.

Accomplishment Description

The focus was on large-eddy simulations of supersonic turbulent

boundary layers and boundary layer receptivity. Suppression of a

Tollmien-Schlichting wave through the motion of a compliant

surface and a springbacked membrane-compliant surface were

simulated. Large-eddy simulations of supersonic, wall-bounded,

turbulent flows were performed. Two compressible subgrid-scale

models based on the Favre-filtered equations of motion for an

ideal gas were tested. Higher order boundary conditions and

adjustments in the computation of the viscous terms resulted in a

more stable numerical scheme. The compressible Smagorinsky

model and the structure function subgrid-scale model results

agreed well with experimental data. In the first figure, a

Smagorinsky model constant density surface is shown.

Horseshoe-shaped structures similar to those found in incompressible boundary layers are visible. The second figure shows

isovelocity surfaces of the streamwise velocity. The lowmomentum fluid near the wall indicates low-speed streaks and

shows regions of these streaks lifting away from the wall near the

horseshoe-like structures. Typical supersonic large-eddy

simulations required 10 Cray Y-MP hours and 6 megawords of

memory when starting in the late stages of transition.

Significance

A better understanding of supersonic, turbulent boundary layers

can lead to improved turbulence models.

Constant density surface from a large-eddy simulation of a

supersonic boundary layer.

Isovelocity surface of streamwise velocity from a large-eddy

simulation of a supersonic boundary layer.

Aeronautics/_urbulence 107
Turbulence Modeling for Three-Dimensional

Linda D. Kral, Principal Investigator

Co-investigators: John A. Ladd, Mori Mani, and John F. Donovan

McDonnell Aircraft Company

Research Objective

To investigate turbulence models used in numerical simulations

of complex viscous flows for critical aircraft components such as

the forebody, wing, nozzle, inlet, diffuser, and afterbody.

Approach

A three-dimensional (3-D) zonal Navier-Stokes code, NASTD,

with a two-equation turbulence model was used to predict

turbulent flow fields around flight vehicles. An implicit, approximately factored, upwind scheme was employed to solve the

3-D, compressible, Favre-averaged Navier-Stokes and energy

equations together with the k-E two-equation turbulence model.

In addition, a compressibility correction for free shear layers was

implemented.

Accomplishment Description

The k-E model was compared with other complex turbulent flow

models. The compressibility correction worked well for highspeed mixing layers; however, when this correction was invoked

for a high-speed boundary layer, the skin friction was significantly lower. Investigation of turbulent flow about an F/A-I 8

forebody at high angle of attack showed that the Jones-Launder

k-_ model predicted flow features similar to the Baldwin-Lomax

algebraic model. Investigation of a supersonic combusting shear

layer showed that proper ignition delay was best predicted

using the Chien low-Reynolds-number model. Calculations

of a transonic supercritical airfoil showed that the algebraic

Baldwin-Lomax and the two-equation model predicted the

shock location and surface pressure distribution, but the results

were extremely grid dependent. Calculations of a 3-D highly

offset diffuser showed the ability of the Chien low-Reynoldsnumber model to predict accurately the flow-field features.

A comparison of the total pressure at the diffuser centerline and

engine face station is shown in the figure. The primary region of

total pressure loss was similar, but the deficit was greater and

more clearly defined in the algebraic model prediction since

there was no eddy viscosity modeled in this region and the

losses did not diffuse as quickly as in the two-equation model

prediction. The algebraic model also predicted a spurious

region of secondary flow. The two-equation model better

predicted recovery and distortion parameters at the engine face.

The calculation of the 3-D diffuser required approximately

I 0 Cray Y-MP hours and 14 megawords of memory for a

converged solution using the two-equation turbulence model.

Flow Fields

Significance

The data used to develop turbulence models were obtained from

canonical flows and it is not clear how these turbulence models

will perform for more realistic geometries. Successful simulation

of turbulent flows on complex aircraft would provide a valuable

design tool.

Future Plans

The performance of the code will be enhanced by improving

the initialization process of the turbulence quantities and

developing methods to handle local problem areas in the flow

using an improved time-stepping procedure. Also, the So and

Huang-Coakley k-E low-Reynolds-number models and the

Menter-blended two-equation model will be evaluated. The

one-equation model of Baldwin-Barth will be evaluated for

these complex flows. Finally, additional corrections for curvature

and compressibility will be evaluated.

Comparison of total pressure contours along the centerline and

at the engine face for a highly offset 3-D diffuser; (top) BaldwinLomax algebraic model prediction, and (bottom) Chien twoequation model prediction. Engine face total pressure contours

are to the right of the figure.

108 Aeronautics/Turbulence
Drag Reduction Mechanism by Riblets

Parviz Moin, Principal Investigator

Co-investigators: John Kim and Haecheon Choi

Center for Turbulence Research/Stanford University/NASA Ames Research Center

Research Objective

To perform direct numerical simulations of turbulent flows over

riblets, to analyze the resulting flow data base, and to educe the

mechanism of drag reduction by riblets. Such an understanding

can potentially lead to the design of riblet configurations that

yield even higher drag reduction.

Approach

Direct numerical simulation is used to simulate turbulent

flows over riblet-mounted surfaces. A fully implicit finitedifference method is used to solve the unsteady incompressible

Navier-Stokes equations in generalized coordinates.

by the streamwise vortices above the wall or by weakening the

strength of the streamwise vortices. Active and passive control

strategies, aimed at influencing the streamwise vortices with the

combinations of blowing and suction on the riblet surfaces, may

be more effective in reducing drag.

Future Plans

We plan to investigate further other riblet configurations and

active control strategies to achieve more drag reduction.

Accomplishment Description

The computed drags on the riblet surfaces are in good agreement

with the existing experimental data. The calculations with riblet

spacing S = 40 and S = 20 with ridge angles of 60 degrees show

12 percent drag increase and 6 percent drag decrease, respectively. The mean-velocity profiles show upward and downward

shifts in the log-law for drag-decreasing and drag-increasing

cases. Turbulence statistics above the riblets are computed and

compared with those above a flat plate. Differences in the meanvelocity profile and turbulence quantities are found to be limited

to the inner region of the boundary layer. Velocity and vorticity

fluctuations as well as the Reynolds shear stresses above the

riblets are reduced in drag-reducing configurations. Quadrant

analysis indicates that riblets mitigate the positive Reynoldsshear-stress producing events in drag-reducing configurations.

From examination of the instantaneous flow fields, a drag

reduction mechanism by riblet is proposed--riblets with small

spacings reduce viscous drag by restricting the location of the

streamwise vortices above the wetted surface so that only a

limited area of the riblets is exposed to the downwash of highspeed fluid that the vortices induce. The total time per time step

on the Cray Y-MP was 60 seconds. Thus, for each case with

S = 20, approximately 200 Cray Y-MP hours were required to

integrate 4,000 viscous time units, and 10 megawords of core

memory and 40 megawords of disk scratch space were required.

Streamwise vorticity evolution time sequence above riblets

(S = 40, drag-increasing case) in coordinates moving with the

structures. Flow is into the page. A strong streamwise vortex is

formed in between riblets.

Significance

The proposed drag reduction mechanism indicates that the skin

friction on the wall can be significantly reduced by reducing the

wetted surface area exposed to the downwash motion induced

Aeronautics/Turbulence 109
Turbulent Flow Over a Backward

Parviz Moin, Principal Investigator

Co-investigator: Hung Le

Center for Turbulence Research

Facing Step

Research Objective

The objectives of this research are to perform a direct simulation

of turbulent flow over a backward facing step with inflow and

outflow boundary conditions, to generate a database for modeling,

and to investigate the physics of turbulent reattachment.

Approach

A flat-plate boundary layer mean velocity profile was applied

at a distance-10 step height prior to a step with superimposed

random fluctuations. The fluctuations were derived using

broadband spectra and were scaled to match all Reynolds stress

components of a boundary-layer flow. Convective boundary

condition was applied at the exit. The Reynolds number based

on the step height was 5100. The simulation used approximately

] 3 megawords of memory and required 21 CPU seconds per

time step on the Cray C-90 with 770 x ] 94 x 66 grid points.

3.0 x/h=4 x/h=6 x/h=10 x/h=19

2.5

2.0

_, 1.5

1.0

.5

0

0 0 0 0 .5 1.0

U/Uo

Comparison of streamwise velocity profiles; solid line = computation; dotted line = experiment.

Accomplishment Description

The statistical results (coefficient of friction, velocity, and

turbulence intensity profiles) were in excellent agreement with

the concurrent experiment at NASA Ames Research Center. The

first figure shows a comparison of the simulated and experimental streamwise velocity. The mean reattachment length is 6h

(h = step height). The second figure illustrates the high threedimensionality of the flow near the reattachment with several

imbedded streamwise vortices. Because of the low Reynolds

number effects, a strong negative skin friction is seen in the

recirculation region. At 20h downstream of the separation, the

presence of the free-shear layer is still strong. In wall coordinates, the mean velocity profiles fall below the universal

log-law, indicating that the flow has not fully recovered.

Significance

The simulations constitute a successful development and

implementation of a method for direct computation of complex

turbulent flows with inflow and outflow conditions. A comprehensive database has been archived for Reynolds averaged

modeling and has up to third-order statistics.

Future Plans

The database will be analyzed to gain understanding of the

physics of turbulent reattachment.

0

Instantaneous velocity vectors in the cross-flow plane at two

step heights before reattachment. Streamwise velocity;

blue = negative, magenta = positive.

110 Aeronautics/Furbulence
Turbulence Simulation on the Connection Machine

J. Blair Perot, Principal Investigator

Co-investigator: Paul Malan

Stanford University

Research Objective

To use direct numerical simulation to investigate the fundamental

fluid and heat transfer physics of wall-turbulence interactions.

Approach

The unsteady, incompressible Navier-Stokes equations and a

passive scalar equation (for heat transfer) are solved using a

second-order, finite-volume staggered-mesh approach in

primitive variables. Time advancement is performed using a

semi-implicit fractional-step method, and all the relevant scales

of turbulent motion are resolved.

Accomplishment Description

A number of different boundaries were studied using grid

sizes over 32 million nodes and obtaining turbulent Reynolds

numbers as high as 375. Performance over a GFLOP was

obtained on the CM-2, and initial results of over 500 MFLOPS

were obtained using the CM-5. Detailed near-wall statistics,

including Reynolds stresses, turbulent heat fluxes, and transport

equation budgets, were calculated. Statistics exist for solid walls,

free surfaces, permeable walls, and moving solid walls. The

near-wall structures associated with kinematic effects, viscous

effects, and energy transfer were fully identified. The understanding of these structures has lead to the development of a number

of improved near-wall turbulence models.

Significance

Understanding and modeling near-wall turbulence is of vital

importance to almost any engineering application which

involves fluid mechanics. This work probes the fundamental

physics of the wall-turbulence interaction and brings closer the

development of rational near-wall turbulence models.

Future Plans

The effects of small mean shear on near-wall turbulence are

under investigation. There are plans to evaluate large-eddy

turbulence models and extend these studies to curved and

rotating walls.

Publications

1. Perot, J. B.: Direct Numerical Simulation of Turbulence on

the Connection Machine. Parallel Computational Fluid

Dynamics 1992, Elsevier Publishing Co., 1993.

2. Perot, I. B.; and Moin, P.: A Near-Wall Model for the

Dissipation Tensor. Eleventh Australasian Fluid Mechanics

Conference, Hobart, Tasmania, Dec. 14-18, 1992.

\

Velocity vectors and temperature contours in several planes

extending perpendicularly from a shear-free wall.

Aeronautics/Turbulence 111
Direct Simulations of Airfoil

Man Mohan Rai, Principal Investigator

NASA Langley Research Center

Flows

Research Objective

To develop a methodology for performing direct simulations

of compressible flow over complex geometries and to use this

methodology to compute transitional/turbulent flows over a

turbine airfoil subjected to moderate-to-high levels of freestream turbulence.

Approach

High-order-accurate finite-difference methods were developed

to perform direct simulations and large-eddy simulations of compressible transitional/turbulent flows over complex geometries.

Accomplishment Description

A computer program to compute transitional/turbulent flows

over an airfoil cross section was constructed. Computations

using this code (for subsonic Mach numbers and a prescribed

level of free-stream turbulence) have been performed on a very

coarse grid. Transitional/turbulent flow over a flat plate was

previously computed using a direct simulation procedure. Freestream turbulence to trigger the flows to transition was obtained

numerically. The power spectrum obtained numerically only

roughly approximated the Von Karman spectrum. The power

spectrum from the experiment very closely approximated the

Von Karman spectrum. Additionally, the power spectrum of the

free-stream turbulence in the experiment closely approximates

the Von Karman spectrum. In anticipation of the direct simulation of the flow about an airfoil, most of the computing resources

allocated to the project were used to obtain a spectrum that is

much closer to the Von Karman spectrum. The first figure shows

the spectrum obtained earlier compared to the Von Karman

spectrum, and the second figure shows the improved spectrum

obtained through a better choice of inlet parameters. The

recently obtained spectrum is in much better agreement with

the Von Karman spectrum.

Significance

This computation represents a first-of-a-kind simulation of flow

transition on an airfoil. The computation will yield a wealth of

information that can be used to understand the transition process

and develop turbulence models. The finite-difference methodology developed can be used for direct simulations of flow over

other complex geometries.

Future Plans

A refined grid airfoil computation will be completed.

Von Karman spectrum

101 F [] Computation

E

_, 100

10-1

101 --

10o

100

(Top) Computed power spectrum for the streamwise velocity

component (used in a flat-plate simulation). (Bottom) Computed

power spectrum for the streamwise velocity component

(generated for the airfoil simulation).

112 Aeronautics/Turbulence
Forced Plane Mixing Layers

Michael M. Rogers, Principal Investigator

Co-investigators: Robert D. Moser, S. Scott Collis, and Chris Rutland

NASA Ames Research Center/Stanford University/University of Wisconsin, Madison

Research Objective

To study the effects of two-dimensional (2-D) forcing on

structures, statistics, and chemical reactions in incompressible

fully turbulent plane mixing layers.

Approach

Direct numerical simulations of temporally evolving incompressible turbulent plane mixing layers with various levels of 2-D

forcing were generated using previously simulated turbulent

boundary layers to provide a realistic turbulent initial condition.

Pseudospectral numerical methods were used to solve the fully

three-dimensional Navier-Stokes equations.

Accomplishment Description

Three fully turbulent mixing layers with Reynolds numbers up

to 20,000 were simulated. A passive scalar quantity was solved

to permit the study of "fast" chemical reactions. In one case,

no 2-D forcing (beyond what is present in the initial boundary

layers) was used. The second case included mild levels of 2-D

forcing, and the third flow was strongly forced. The first two

layers evolved self-similarly, with mean velocity and Reynolds

stress profiles that collapsed when scaled by the linearly

increasing layer thickness. The third case was approximately

self-similar, although the collapse in scaled coordinates was not

as good as the other two cases. The evolution of the strongly

forced case was different from the other two. The flow was

dominated by large spanwise 2-D rollers with pronounced

braid regions separating them in the streamwise direction (see

figure). These rollers undergo a classical "pairing" process, and

the large incursions of free-stream fluid engulfed in the layer

result in mixed fluid with a fairly uniform scalar concentration

across the layer. In contrast, the unforced and mildly forced

cases do not exhibit such organization in the self-similar period.

There are no "clean" braid regions separating rollers, and the

mixed fluid scalar concentration across the layer resembles the

mean scalar profile. The eddies in the flows provided good

turbulence statistics. The profiles of all the terms in the Reynolds

stress balance equation were computed to provide data for

further model development. The simulations required up to

512 x 210 x 192 modes. The two forced simulations required

approximately 900 CPU hours and used up to 80 megawords

of memory on the Cray Y-MP.

Significance

These simulations were the first turbulent free-shear-flow direct

numerical simulations to achieve self-similarity. The strong effect

of forcing on the layer structure helps explain discrepancies in

experimental observations by different investigators.

Future Plans

Similar simulations of turbulent plane wakes will be generated.

!T '

Concentration (red = high, blue = low) of passive scalar quantity

in the self-similar period for the unforced layer (top), mildly

forced layer (middle), and strongly forced layer (bottom). Views

are at typical spanwise locations, with the streamwise direction

horizontal.

Aeronautics/Turbulence 1 1 3
Development of a Turbulent Spot

Bart A. Singer, Principal Investigator

Co-investigator: Ronald D. Joslin

High Technology Corporation/NASA Langley Research Center

Research Objective

To identify and understand the mechanisms involved in the

formation and growth of a turbulent spot in a boundary layer.

Approach

Spatial direct numerical simulation (DNS) of the incompressible

Navier-Stokes equations was used to generate the data for this

study. Data visualization was performed using the Flow Analysis

Software Tool.

Accomplishment Description

Previously, hairpin vortices were generated by a variety of

different localized wall disturbances. In one case, the hairpin

vortex grew, elongated, and spawned a secondary vortex head

upstream of the primary vortex and a subsidiary vortex developed beneath the hairpin vortex legs. Work continued the DNS

into the early stages of a turbulent spot. As the flow became

more complicated and smaller scales developed, the resolution

requirements rose from 5.8 to 8.9 million grid points, increasing

the core memory requirements for the code from about 125 to

200 megawords. Approximately 500 Cray-2 hours were required

in the computation. Multiple additional vortices developed in

the flow. These vortices fell into three categories: (1) new hairpin

vortex heads, (2) horseshoe vortices forming beneath other

vortex legs, and (3) quasi-streamwise vortices. Examples of each

can be seen in the first figure (low-pressure regions are used to

mark the vortex cores). Beneath the hairpin vortex in the

upstream portion of the flow is a small horseshoe vortex. The

elongated structures in the figure consist largely of streamwise

vorticity. The vortices seem to form the core region of a developing turbulent spot. In the second figure, a plan view of vertical

vorticity contours reveals the arrowhead shape of a turbulent

spot.

Significance

This work provides a better understanding of the way in which

coherent flow structures evolve from a single localized disturbance to a group of structures that form the early stages of a

turbulent spot. Clues are provided regarding the types of

structures that dominate the flow and how these structures

interact and regenerate, resulting in a more turbulent flow.

These insights should result in more realistic models for predicting flows in the transition region.

Future Plans

The calculation will be continued until a fully developed

turbulent spot is obtained. Data will be analyzed to understand

how the various types of vortices form, interact, and regenerate.

As new data become available, the role that the vortices play in

engulfing new fluid into the turbulent region will be studied.

Transition region modeling information will be obtained.

Plan view of low-pressure regions in flow. Flow goes from left to

right.

Plan view of vertical vorticity magnitude. Vorticity levels:

red = high and blue = low.

. _ .

1 14 Aeronautics/Turbulence
Effect of Suction Holes

Philippe R. Spalart, Principal Investigator

Boeing Commercial Airplane Group

Research Objective

To understand and predict better the fluid dynamics associated

with suction holes used in laminar flow control to prevent the

transition of a boundary layer to turbulence.

Approach

Direct numerical simulations (DNS) were used to produce

solutions to the full Navier-Stokes equations in boundary

layers under conditions characteristic of the flow over a wing.

The "swept Hiemenz flow," a classical exact solution of the

Navier-Stokes equations, was considered with focus on stationary crossflow disturbances. The suction holes were represented

by a prescribed nonzero, non-uniform normal velocity at the

wall. Single rows and arrays of holes were treated. The Reynolds

number, row location, hole diameter, hole spacing in each

direction, and flow rate led to a large parameter space.

Accomplishment Description

About 20 converged solutions were obtained and grid-refinement

studies were conducted. Linearity was checked when appropriate. In combination, these solutions allowed the calculation of

the "receptivity coefficient," which gave the initial amplitude of

the crossflow vortex created by the suction holes. The enhanced

receptivity when the crossflow vortex was aligned with the holes

was computed. The figure shows the streamwise direction and

u' is the amplitude of the crossflow vortices, normalized by the

average suction velocity. Pure crossflow vortex eigenmodes

became unstable near x = 250. For x less than about 400, the

response was dominated by the local disturbances identified

with each row; their spacing was 10. Beyond x = 400 an

unstable mode overshadowed the local disturbances and a fairly

smooth growth curve was seen. The curve with square symbols

had a low suction rate, which allowed rapid growth of the

instability. The curve with diamonds had a much higher suction

rate, which considerably reduced the rate of growth without

completely suppressing it. The normalized amplitudes were very

close around x = 300; thus the receptivity coefficient was not

strongly affected by the suction level. The central processing unit

time was several hundred hours for fully turbulent solutions and

the central memory was up to 10 megawords.

Significance

Receptivity theories are making rapid progress, but it is essential

to obtain quantitative results by other means in order to validate

them. It is also necessary to find out when theories fail because

of nonparallel, nonlinear, or similar effects. The conjunction of

DNS and more theoretical approaches substantially improves the

production of competitive designs for laminar flow control

suction systems.

Future Plans

Related instability issues in three-dimensional boundary layers

are being explored using a pressure distribution more realistic

than in Hiemenz flow.

"O

2

E

1000 -- _ Curve with lower

-- suction rate

-- _ Curve with higher

suction rate

100

10

250 300 350 400 450 500

Streamwise coordinates (x)

Crossflow vortices created by arrays of holes.

PAGE_ INTENTIONALLYBLANK

Aeronautics/Turbulence 115



Space Station Flow Analysis

Paul McConnaughey, Principal Investigator

Co-investigators: Eric Stewart, Bruce Vu, and Lee Kania

NASA Marshall Space Flight Center/Sverdrup Technology, Inc.

Research Objective

To facilitate the design of the air distribution system aboard

Space Station Freedom. By performing computational fluid

dynamics (CFD) analyses of the internal ventilation flows, the

ability of the air distribution system to satisfy performance

requirements will be assessed and any shortcomings exposed.

The results will be used by the environmental control and life

support system (ECLSS) design team to devise any necessary

system modifications.

Approach

The CFD analyses were performed with the INS3D code. The

code was modified to permit the use of blocked grids, and it

exhibited only a small increase in the characteristic time. The

flow-field analyses were also conducted using an algebraic

turbulence model specifically calibrated for internal ventilation

flows. The model proved to be an acceptable compromise

between computational efficiency and accuracy. A steady-state

flow solution was obtained with volumetric flow rates specified

at each of the ventilation registers.

Accomplishment Description

Grids were generated for a generic module and resource node

and were comprised of 1.88 and 3.3 million points, respectively.

The resource node contained eight supply registers and seven

return registers. A converged, turbulent flow field was obtained

in 11,000 iterations and required 160 Cray-2 hours and

57 megawords of memory.

Significance

The air distribution system produced average velocities in the

range of 40 feet per minute, thus satisfying the ECLSS design

requirement in limited portions of the node only. There were

regions of low-speed flow and short-circuit ventilation that

indicated the ventilation registers were not optimally located.

These flow-field data will be used in conjunction with test data

to devise modifications to increase system efficiency.

Future Plans

The node and module elements will be coupled via a berthing

vestibule and the flow-field analysis will be repeated. Additionally, the flow fields will be used for performing contaminant

tracking analyses in the design of the fire protection systems.

Resource node velocity magnitude contours; supply registers are green, return registers are red.

PI_[£[O__ PAGE BLANK NOT F|L_,ED

Ill

PAGE [ [_ tNTENTIO,tlAt.LY 8LArlK

General/Astronautics 119
Parallel Integration of N-body Gravitational Systems

Jeffrey D. Scargle, Principal Investigator

Co-investigators: Jeffrey N. Cuzzi, Anthony Dobrovolskis, Luke Dones, Robert Hogan, Creon Levit, Mark Showalter, and Karl Young

NASA Ames Research Center

_J

Research Objective

By simulating the long-term dynamical evolution of the satellites

and ring particles of the planet Saturn, many questions were

addressed: How did Saturn's ring structure originate? How stable

is the configuration now observed? What is the connection

between the shepherd satellites, Prometheus and Pandora, and

the narrow, kinky F Ring that lies between them?

Future Plans

Test particles will be distributed over the entire ring system.

Estimates of the strengths of resonant and nonresonant interactions between satellites and test particles suggest that the

structure of Saturn's main ring system is a result of dynamical

effects. Computations will be done to learn more about the

origin, evolution, and stability of the rings.

Approach

Direct numerical integrations were completed for the Newtonian

equations of motion for a system with an oblate Saturn, its

18 most massive satellites, and a large number of massless

particles. Accurate orbital trajectories and the largest local

Lyapunov exponent of the system were computed to detect

chaotic dynamics. Orbital integrations were carried out with a

parallelized Bulirsch-Stoer integrator.

Accomplishment Description

A series of runs following the evolution of 8,000 moonlets

uniformly distributed in the vicinity of the orbits of the shepherd

satellites Prometheus and Pandora was completed. The moonlets

underwent forced motions on short and long time scales and

achieved orbital eccentricities on the same order as the F Ring.

Their semimajor axes were redistributed much like the structure

of the F Ring. Shepherd motion was not chaotic, but the local

Lyapunov exponent was occasionally positive for some moonlets, indicating the presence of mildly chaotic, bounded excursions of these particles. A connection for observable chaotic

variations in the orbital elements was established, allowing

interpretation of relatively short integrations (thousands of years)

in terms of long effects (millions to billions of years).

Significance

Orbits of the shepherding satellites are quite regular and will

not evolve significantly over time equivalent to a large fraction

of the age of the solar system. The motion of massless particles

exhibits a mild form of chaos in which the orbits wander

pseudo-randomly within a bounded region. Thus the system is

chaotic, but also stable, against large-scale disruption over

relatively long times.

Evolution of orbital elements of an ensemble of 8, 000 test

particles surrounding Saturn's shepherd satellites Prometheus

and Pandora. Upper panels based on the best estimates for the

masses of Prometheus, and lower panels for masses twice as

large. Left panels: initial semimajor axis of each particle plotted

against its semimajor axis. Color indicates the eccentricity of the

particle (red = circular orbit, blue = largest eccentricity). The

X-shaped patterns are particles trapped in "horse-shoe" orbits.

The reddish bar near the center is comprised of particles largely

undisturbed by the shepherds. Right panels: semimajor axis of

each particle plotted against its eccentricity; color indicates the

inclination of the orbit relative to the main orbital plane of

Saturn.

120 General/Astronomy
Numerical Experiments in the Formation and Evolution of Galaxies

Bruce F. Smith, Principal Investigator

Co-investigators: Richard H. Miller and Thomas Y. Steiman-Cameron

NASA Ames Research Center/University of Chicago/University of California, Santa Cruz

Research Objective

To develop numerical experiments to study the important

dynamical processes at work in the formation and subsequent

evolution of galaxies.

Approach

The numerical experiments were based on the time development of a fully three-dimensional, self-gravitational, and selfconsistent particle system. The codes followed the motions of

105-10 _ particles that represent the stellar and gaseous components of galaxies. The gravitational potential and forces were

calculated on a 256 _ grid.

Accomplishment Description

The particle codes were used on the Cray-2 and the Cray Y-MP

to investigate several important problems in galactic dynamics.

Observations and numerical experiments indicated that significant noncircular motions are present in disk galaxies. Global

galactic oscillations were identified in galaxy models. The

experiments suggested that normal-mode oscillations may be

present in nearly all galaxies, including the Milky Way, at

considerably higher amplitude than previously thought. The

effects of these oscillations may be most prominent in disk

galaxies. Recently, experiments were conducted on the dynamics of spiral galaxies embedded in oscillating, dark-matter halos.

The results indicated the development of ring structures that can

become offset relative to their center and the development of

bars that are often observed.

Significance

Observations of galaxies give us intriguing snapshots of complex, evolutionary sequences. The numerical experiments

proved to be an effective approach to understanding the

dynamic sequences in galactic evolution. This computational

capability is necessary to understand the wealth of new astronomical data becoming available from the Hubble space

telescope and other new observational platforms. The results on

the nature of non-steady-state motions in galaxies have a

profound effect on our understanding of the morphology and

dynamics of galaxies.

Future Plans

Additional physics will be incorporated into the codes used for

the numerical experiments. The gas dynamics representing the

gaseous component in galaxies was developed, and preliminary

runs are under way. High-resolution studies of the interplay

between stellar and gas dynamics are now possible, allowing

galaxy formation studies and studies of galaxy interactions with

highly stimulated star formation.

Various views of the warping of a disk galaxy after it has been captured by a larger nearby galaxy.

General/Astronomy 121
Numerical Simulations of Baroclinic

Jeffrey R. Barnes, Principal Investigator

Oregon State University

Research Objective

To investigate nonlinear baroclinic instability and transient

baroclinic eddies in the atmospheres of Mars and Earth through

numerical simulation.

Approach

A three-dimensional, spectral, primitive-equation model of

atmospheric circulation was employed to simulate the nonlinear

evolution of transient baroclinic eddies. This model offered

computational efficiency via a semi-implicit time scheme, high

spatial resolution, flexibility in configuration, and on-line energy

diagnostics.

Accomplishment Description

Synthetic basic states were developed for use in the nonlinear

Mars simulations. Several states were fully global, but they

proved to be computationally troublesome. A large set of singlewave experiments were performed using a hemispheric basic

state (representing the winter flow). The simulations yielded

several interesting results. There was great variation in structure

between the longer (waves 1 and 2) eddies and the relatively

short eddies (wave 4). The long eddies were extremely deep and

latitudinally broad and exhibited maximum amplitudes at very

high levels (above 20-30 kilometers altitude), while the short

eddies were much shallower and more latitudinally confined.

Wave 3 was an intermediate case, and it tended to reach the

largest amplitudes in the simulations. Wave 4 had a close

balance between energy conversions producing growth and

decay, and thus the wave amplitude (and structure) stayed

nearly constant. The longer eddies tended to have a "life cycle"

behavior--the waves grew to large amplitudes and then decayed

to relatively small amplitudes.

Significance

Transient baroclinic eddies are of great importance for the

Earth's atmospheric circulation in middle latitudes. Cyclonic

weather systems are manifestations of these eddies, which

transport large amounts of heat poleward. Major NASA programs

are directed at an enhanced understanding of the atmospheric

heat balance and the global climate system. Spacecraft observations show that transient baroclinic eddies are present in the

atmosphere of Mars, and observations and modeling demonstrate that they are of considerable importance for the atmospheric circulation and the climate system. The terrestrial and

Martian atmospheres are similar in many ways; thus they

constitute two "laboratories" where the basic dynamics of

transient baroclinic eddies can be investigated.

Instability

Future Plans

Mars Observer data should soon yield a great deal of information

on the structure of the transient eddies and the structure of

atmospheric basic states. Mars Observer will also obtain very

accurate topographic measurements. These new data should

make it possible to launch a much more detailed modeling study

of the dynamics of the transient eddies, with extensive examination of their seasonal, hemispheric, and interannual variations.

Publication

Barnes, J. R.; and Young, R. E.: Nonlinear Baroclinic Instability

on the Sphere: Multiple Life Cycles with Surface Drag and

Thermal Damping. J. Atmos. Sci., vol. 49, 1992, pp. 861-878.

8o

70

60

SO

e

40

3o

20

lO

o

o

-f'lO

10 20 30 40 50 00 70 80 90

Latitude

•iO -;O ._0 -is o *s 30 50 g'o 1io

Latitude-height cross section of the north-south heat flux by a

zonal wave 1 disturbance in a nonlinear simulation of baroclinic

instability in the Mars atmosphere.

122 General/Atmospheric Science
Orographically Forced Oscillations

Christian L. Keppenne, Principal Investigator

Jet Propulsion Laboratory

in the Martian Atmosphere

Research Objective

The intraseasonal time scales of the terrestrial atmosphere's

variability are dominated by oscillations induced by barotropic

instability over orography. Since the irregularities of Earth's and

Mars' surface reliefs compare similarly to their respective scale

heights, orographically forced modes should also be expected in

the Martian atmosphere. A model is used to search for such

oscillations.

Approach

A high-resolution barotropic model with realistic Martian

topography was run in several 6,000-sol (1 sol = 1 Martian day)

integrations. The model was based on the same set of dynamical

equations as terrestrial general circulation models, but its

formulation did not account for thermodynamic processes such

as the cycles of water on Earth or of carbon dioxide on Mars. As

a result, barotropic instability over orography is the main cause

of variability. The atmospheric angular momentum (AAM) of the

last 5,000 sols of each run were submitted to maximum entropy

spectral analysis. Extended empirical orthogonal function

analysis was applied to the model's pressure field in order to

analyze the spatial structure of the variability identified in the

AAM time series.

Accomplishment Description

The code was multitasked on the Cray Y-MP. The prognostic

equations were integrated in spherical harmonic space using a

semi-implicit scheme, but the nonlinear terms were computed

on a latitude-longitude grid. At high resolution, performance

came close to 1 GFLOP using all 8 processors. At more

modest resolution, performance decreased to 500 MELOP

(0.2 second/sol) and memory use was about 8 megawords.

A robust oscillation with a mean period near 85 sols was

identified. Two other, shorter period oscillations with mean

periods near 3 and 6 sols were also present. The power spectra

of local-pressure time series correspond well to those of the

Viking Lander I and II data. This data suggests that orographically forced instability can play a significant role in triggering

the Martian atmosphere's observed variability.

Significance

The oscillations identified are of sufficient amplitude to dominate the Martian atmosphere's intraseasonal variability. Not only

does the study lead to advancements in Martian climatology, but

an improved understanding of the mechanisms involved should

have repercussions in the meteorology of Earth and other

planets.

Future Plans

The model will be used to develop a sophisticated data assimilation scheme for Mars Observer data. This will result in nearoptimal weights to combine observations and model data in

order to obtain a better picture of the Martian atmosphere than

could be provided by either the observations or model alone.

o.o o.1 o.12 _| 0.3 o,4

-aO

180_180

Geographical distribution of the frequency corresponding to the

highest spectral density in the model's local-pressure time series.

The latitude and longitude axes are labeled in degrees and

frequencies are expressed in cycles/sol.

Generai/Atrnospheric Science 123
Dynamics of the Martian Atmosphere

James B. Pollack, Principal Investigator

Co-investigator: Robert Haberle

NASA Ames Research Center

Research Objectives

To simulate the present and past climate regimes on Mars. The

tool used to do these simulations was a general circulation

model (GCM) that predicted the time-evolving three-dimensional

wind and temperature fields and surface pressure, temperature,

and carbon dioxide ice abundance.

Approach

The Mars GCM was based on the primitive equations of

meteorology used for terrestrial weather prediction. Uniquely

Martian physics, such as the condensation and sublimation of

carbon dioxide in the polar region, were incorporated into the

Mars GCM. The Mars GCM was interfaced with an aerosol

physics model to simulate the transport of dust by winds and the

impact of absorption of sunlight by dust on temperatures and

winds.

Accomplishment Description

Analyses of GCM simulations at different seasonal dates and

dust Ioadings were carried out to define the properties of several

major components of the atmospheric circulation. Eastward

propagating high- and low-pressure systems develop at middle

latitudes of the winter hemisphere. Their simulated properties

compare favorably with meteorological measurements. Simulations required 250 Cray-2 hours, with a typical job using

5 megawords of memory.

Significance

From an atmospheric dynamical perspective, the Martian

atmosphere represents the closest analogue to the Earth's

atmosphere. By testing climate models extensively used for the

Earth in other planetary atmospheres, we can assess the validity

of these models and our understanding of basic dynamical

processes. The traveling high- and low-pressure systems of the

Martian atmosphere are produced by the same basic mechanism

as the baroclinic instability that causes middle latitude weather

on Earth. The Mars GCM predicts that the traveling lows and

highs on Mars are more periodic than the ones on Earth, and

it is in agreement with spacecraft data.

Future Plans

Improved simulations of the general circulation of the Martian

atmosphere will be carried out using higher spatial resolution

and introducing a parameterization for drag due to vertically

propagating gravity waves.

Publication

Pollack, J. B.; Haberle, R. M.; Murphy, J. R.; Schaeffer, J.; and

Lee, H: Simulations of the General Circulation of the Martian

Atmosphere. Two Seasonal Pressure Variations. J. Geoph. Res.,

vol. 98, 1993, pp. 3,149-3,181.

2

(a)

9O° N

6O° N

30° N

O*

60 ° 5

9O_ S

180° W 135° W 90 ° W 45 ° W 0 ° 45 ° E 9O ° E 135° E leO° E

LONGITUDE

(b)

Surface pressure maps corrected for topography. The dark red

and dark blue regions have pressures of 7.0 millibar and

+0.45 millibar, respectively, relative to the zonal average.

(a) Time = 984 hours; (b) time = 1,008 hours. Note the eastward

displacement of highs and lows in (b), which is one Martian

day later than (a).

124 General/Atmospheric Science
Simulation of Volcanic Aerosol Clouds

Richard E. Young, Principal Investigator

Co-investigators: O. B. Toon and J. B. Pollack

NASA Ames Research Center

Research Objective

To numerically simulate the behavior of the El Chichon and

Mt. Pinatubo volcanic aerosol clouds in the stratosphere in order

to better understand stratospheric transport and aerosol microphysical processes. An associated goal is to assess the climatic

impact of such large volcanic eruptions on stratospheric wind

and temperature fields.

Approach

A three-dimensional (3-D), spectral, primitive equation model

was used to compute wind and temperature fields in the

stratosphere. The model was coupled to a 3-D aerosol transport

and microphysical model, which computed the dispersion of the

volcanic aerosol cloud using the computed winds from the

circulation model.

Accomplishment Description

Both noninteractive and interactive tracer simulations for the

Mt. Pinatubo (eruption latitude 15 ° north) volcanic aerosol

cloud have been conducted for the first 3 months following the

eruption. Each 3 month simulation required about 6 Cray Y-MP

hours and 16 megawords of memory. The results were compared

to the satellite, airborne, and ground-based data sets that exist for

the aerosol cloud. Most of the aerosol mass resided in the

stratosphere above 20 kilometers (km) altitude. The simulated

Mr. Pinatubo cloud spread rapidly southward as observed only

when interactive heating of the aerosol cloud was accounted for,

reaching 20 ° south latitude within weeks after the eruption. The

interactive model simulations reproduced these characteristics.

The figure compares the simulated Mr. Pinatubo volcanic gas

and aerosol cloud after it had circumnavigated the globe once

with and without interactive heating. The noninteractive cloud

(top) basically stays north of the equator; the interactive simulation (bottom) produces a cloud that spreads rapidly southward.

Initially the cloud was a column 20-25 km over the Philippine

Islands, but eventually it spread out because of wind shear.

Significance

Unique opportunities to better understand stratospheric

transport and aerosol microphysical processes occurred when

the El Chichon and Mt. Pinatubo erupted--the two largest

volcanic eruptions this century in terms of aerosol material

injected into the stratosphere. A comprehensive climatology of

the subsequent dispersal of the volcanic clouds, gathered from

aircraft, satellite, and ground-based observations, has been

compiled for the years immediately following the eruptions.

By simulating the behavior of the volcanic clouds, theoretical

understanding of transport and aerosol processes in the stratosphere can be assessed and improved.

Future Plans

Multiyear simulations of both the Mt. Pinatubo and El Chichon

volcanic aerosol clouds will be conducted in passive-tracer and

interactive mode. Climatic feedback on the stratosphere will be

assessed.

Comparison of noninteractive simulation (top) with interactive

simulation (bottom) where heating of the volcanic cloud due to

infrared radiation from the troposphere is accounted for. Solar

heating is much smaller. Dotted lines represent 15° of latitude.

General/Atmospheric Science 125
Propane-Air Combustion Mechanism

Richard L. Jaffe, Principal Investigator

Co-investigators: Charles W. Bauschlicher, Harry Partridge, David Schwenke, Christopher Dateo, and Stephen P. Walch

NASA Ames Research Center/Eloret Institute

Research Objective

To support the development of a detailed propane-air combustion model, which is being used to optimize the design of the

propulsion system for the High-Speed Civil Transport (HSCT)

under the High-Speed Research Program (HSRP), by calculating

thermodynamic properties, transport coefficients, and reactionrate coefficients for intermediate species formed during the

combustion process.

Approach

Ab initio quantum chemistry methods were used to compute

potential energy surfaces for individual molecular species or

pairs of species. Collision cross sections and reaction-rate

coefficients were computed for binary collisions with intermolecular forces determined from the gradients of the potential

energy surface.

Accomplishment Description

There were four major accomplishments: (1) the accurate heat of

formation of hydroperoxy radical (HO2), universally the most

important combustion intermediate, was computed to reduce the

uncertainty in experimental determinations; (2) the thermodynamic properties of benzyl alcohol were computed to resolve a

factor of 20 discrepancy in conflicting thermodynamic property

data bases; (3) the potential energy surfaces were computed for

methanol decomposition (CH3OH) and for the prompt NOX

initiation reaction CH + N 2 --> N + HCN; and (4) the hightemperature rate constants for the important chain branching

reaction H + 02 --_ OH + O (see figure) were computed.

The calculations used 1,000 Cray-2 hours and required

15-24 megawords of memory and at least 300 megawords

of temporary disk space. The HO 2 calculations required more

than 1 gigaword of disk space.

Significance

These calculations provide input to detailed combustion models

used for predictions of engine performance. Particular emphasis

is placed on predicting the amount of NOX pollutants that

would be formed during engine operation. One of the major

goals of the HSRP is to design a nonpolluting supersonic aircraft

engine.

Future Plans

Work will continue on individual aspects of the combustion

mechanism for propane-air and propane-toluene-air systems.

Publications

1. Bauschlicher, C. W.; and Partridge, H.: Accurate Determination of the Heat of Formation of HO2. Chem. Phys. Letters,

vol. 208, 1993, p. 241.

2. Walch, S. P.: Theoretical Characterization of the Reaction

CH 3 + OH --->CH3OH --_ Products: The CH 2 + H20 ,

H2 + HCOH, and H 2 + H2CO Channels. J. Chem. Phys.,

vol. 98, 1993, p. 3,163.

- O-- k(T) = 1.1469 "10 x exp(-7228.8/T);

present work

k(T) = 1.6230 "10 x exp(-74740/T);

Du and Hessler

•.._... k(T) = 6.8250 "11 x exp(-5353.3/T);

Varandas DMBE IV

10-10

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0

10-12

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H + O, --_ 0 + OH rate coefficients.

126 General/Chemistry
Boost Phase Detection Studies

Stephen Langhoff, Principal Investigator

Co-investigators: Charles Bauschlicher, Richard Jaffe, Timothy Lee, and Harry Partridge

NASA Ames Research Center

Research Objective

To determine the radiation signature from the hard-body missile

bow shock layer during the assent phase of the trajectory. Also

studied was the chemistry occurring in the plume of the rocket

because radiation from the plume constitutes a large part of the

optical signature.

Approach

Computational chemistry methods were used to determine

radiative intensity factors, transport properties, electron impact

excitation cross sections, and reaction rate constants for air

species under thermal nonequilibrium conditions. These

methods were also used to determine the electronic structures

and vibrational frequencies of polyatomic molecules in the

plume.

Accomplishment Description

Electronic transitions in the nitrogen-oxygen (NO) molecule

constituted an important source of radiation in the bow shock

layer of hard bodies. Extensive theoretical calculations were

done to determine the band strengths of the Ogawa band system

of NO. To determine accurate electronic transition probabilities

required 4 Cray hours and about 20 megawords of memory at

each internuclear distance. Calculations were also carried out to

characterize the quintet states of carbon-oxygen (CO). These

states have not been characterized experimentally and they are

potential precursor states for populating the upper state of the

Cameron band system, which is prominently observed in

plumes. Calculations showed that the quintet pi state was

significantly bound. Furthermore, transitions between this state

and the low-lying triplet states explained perfectly the emission

spectrum between about 360-520 nanometers observed in a

matrix. The calculations showed, however, that this could not be

a mechanism for explaining the long-lived chemiluminescence

from the Cameron band system. The relevant potential energy

curves are shown in the figure.

Significance

Accurate transition probabilities are required to characterize the

radiation from the bow shock in front of hard bodies and from

the plume. The theoretical band strengths from this study are

very accurate and remove one source of error in modeling the

radiation. Calculations on the quintet states of CO have clarified

their role in the spectroscopy of this molecule.

Future Plans

The study of selected band systems will continue. The goal is

to characterize band systems that contribute to the radiation

signature at wavelengths less than 2,000 angstroms.

Publication

Bauschlicher, C. W.; Langhoff, S. R.; and Partridge, H.: Theoretical Study of the Lowest Quintet Pi and Sigma States of CO.

J. Chem. Phys., vol. 98, 1993, p. 8,785.

5000

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A

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-35,000

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Computed potential energy curves for the lowest quintet pi and

sigma states and the Klein Dunham curves for selected singlet

and triplet states of CO. Horizontal lines denote the positions of

the first three vibrational levels.

General/Chemistry 127
Computer Simulation of Astrophysical

Michael A. Wilson, Principal Investigator

Co-investigator: Andrew Pohorille

NASA Ames Research Center

Research Objective

To create a computer simulation of water ices formed under

astrophysical conditions.

Approach

A direct molecular dynamics (MD) computer simulation of

deposition events was completed. A low-density amorphous

(LDA) ice substrate was prepared by annealing a system of

343 water molecules from 300 K at a density of 0.94 grams per

cubic centimeter. This substrate was then used for deposition of

water and guest molecules from vapor.

Accomplishment Description

Computer simulations of hot (300 K) and cold (77 K) deposition

of water molecules onto a cold LDA ice substrate were performed. The structure of the emerging ice retained much of the

tetrahedral structure of the LDA ice, but the deposition experiments at 77 K were characterized by large void regions, whereas

the 300 K simulations yielded a more compact structure. This

compact structure occurred because of the larger net energy of

the 300 K molecules, which allowed greater surface reconstruction after deposition. Computer simulations of co-deposition

experiments of water with argon, krypton, and methane (in a

ratio of 20:1 ) were carried out at 77 K. The ices formed in the

co-deposition experiments were similar to the pure ices. The

guest species were found to stick with probabilities of nearly

one. Deposition simulations were also tried with neon, but it

did not bind to the water substrate at 77 K. Preliminary studies

showed that the microporous structure became denser upon

warming and some outgassing occurred. Individual runs required

about 15 Cray Y-MP hours and less than 8 megawords of

memory.

Significance

Laboratory experiments on water ices involve depositing hot

water molecules onto cold substrates. These conditions differ

from astrophysical conditions, where the impinging molecules

are at about the same temperature as the substrate. Consequently, the morphology of ices created in the laboratory might

be quite different from those formed in the outer solar system.

Ices

Future Plans

We will investigate the structural changes and outgassing that

occur upon heating and will calculate the thermal conductivity

of amorphous ices.

Surface of the system of water and krypton co-deposition

experiment after 400 depositions. Ratio of argon to water = 1:20.

Temperature of the impinging molecules and the substrate =

77 K. Neither the water nor the krypton atoms are particularly

mobile on the surface of the substrate after they have been

deposited. Periodic boundary conditions were used in the x

(horizontal, points out of the page) and y (vertical) directions.

128 General/Chemistry
Radiation

James P. Wolfe, Principal Investigator

Co-investigators: Robert Averback and Huilong Zhu

University of Illinois, Urbana/Champaign

Effects in Intermetallic Compounds

Research Objective

To investigate the atomistic processes that occur in intermetallic

compounds during irradiation with heavy ions and to determine

the number of defects produced and the amount of atomic

disordering that takes place in displacement cascades.

Approach

Molecular dynamics simulations were employed to follow

the atom trajectories in energetic displacement cascades

initiated by assigning one target atom an initial kinetic energy

of 10 kilo-electron volts/keV). Embedded atom method potentials are employed to represent the intermetallic compound,

nickel-aluminum (NiAI).

Accomplishment Description

Molecular dynamics computer simulations were used to

investigate the collision dynamics of energetic displacement

cascades in a B2 crystal structure, NiAI, for the first time. The

10 keV event was the highest energy cascade in an intermeta[lic

compound ever simulated by molecular dynamics. Careful

analysis of the degree of chemical and structural order throughout the history of the cascade event was performed. As shown

in the series of snapshots of atomic positions (see figure), crystal

structure is temporarily lost in a well-defined region of space,

giving the appearance of a local amorphous zone. Chemical

order is also seen to be lowered in this region. These simulations

made possible the determination of the number of atomic

defects, Frenkel pairs, and anti-site defects produced in the

cascade event. Each cascade event required 15 Cray-2 hours.

Significance

Stability of intermetallic compounds in a radiation environment

is important in the design of structural components to be used in

reactors or in space. The present molecular dynamics simulations provide understanding of the damage process in these

materials on an atomic level. This investigation is easily generalized to other systems. The results are useful in understanding the

point defects in crystalline matter and the stability of ordered

phases.

Future Plans

New simulations of cascades at metal surfaces and interfaces

are planned. The former is important for surface damage like

roughening and sputtering; the latter is important for radiation

effects involving second phase precipitates and artificially

structured materials.

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(c) 0.5 picoseconds, and (d) 6.0 picoseconds. Open dots = aluminum and solid dots = nickel.

General/Chemistry 129
Performance and Scalability of Parallel Graphics Libraries

Thomas W. Crockett, Principal Investigator

ICASE

Research Objective

To develop a three-dimensional (3-D) parallel graphics library

for message-passing architectures and to assess its performance

and scalability as a function of scene complexity and machine

size.

Approach

In previous work, a parallel polygon rendering algorithm suitable

for multiple-instruction/multiple-data message-passing architectures was developed. That work was extended by incorporating

an improved version of the algorithm in a parallel 3-D graphics

library with significant improvements in functionality and

performance. The parallel library was designed specifically for

inclusion in parallel application programs that employ a single

program multiple data programming style. Images generated

were compressed and transmitted across the network to the

user's workstation, where they were decompressed and displayed.

Accomplishment Description

A 3-D parallel graphics library (PGL) was developed to support

a standard graphics pipeline, including modeling transformations, lighting computations, 3-D clipping, perspective viewing,

Gouraud-interpolated rasterization, and z-buffered hidden

surface elimination. The library incorporated an improved,

span-based version of the Crockett/Orloff parallel rendering

algorithm. Preliminary results indicated that the new algorithm

was approximately twice as fast as the original on a standard

test scene• Parallel efficiency was doubled, reaching 60 percent

on the 128-processor iPSC/860 at NASA Ames. The compressed

image transmission scheme incorporated within PGL appeared

to be practical. Experiments using the Ames iPSC/860 from

NASA Langley demonstrated transmission times of about

three seconds per frame.

Test scene composed of 100,000 random triangles. The statistical properties approximate the assumptions of the analytical

performance models.

Significance

Applications that run on massively parallel supercomputers

produce massive output datasets that are often cumbersome to

move across the network for post-processing. Using the power

of the parallel machines to perform graphics and visualization

operations in place reduced the output to a manageable size.

Embedding graphics calls within parallel applications allows

users to take advantage of visual data representations for

debugging, execution monitoring, and interactive steering.

A more typical image. Load balancing becomes an issue because

the geometric complexity varies within the scene.

Future Plans

Performance evaluation of the span-based rendering algorithm is

under way. Of particular interest is the behavior as the number

of processors approaches or exceeds the number of scanlines in

the image• Additional work is needed to develop parallel

visualization algorithms.

130 General/Computer Science
Performance Monitoring of Parallel Programs

Jerry C. Yan, Principal Investigator

Co-investigators: Philip J. Hontalas, Catherine H. Schulbach, Pankaj Mehra, Sekhar R. Sarukkai, Melisa A. Schmidt, and Tarek S. Elaydi

Recom Technologies/NASA Ames Research Center/Lawrence Berkeley Laboratory

Research Objective

To develop tools and methodologies for performance evaluation

of multiprocessors.

Approach

The Automated Instrumentation and Monitoring System (AIMS)

is a software tool set that automatically instruments parallel

FORTRAN and C programs, monitors their execution on

multiple-instruction/multiple-data architectures, and analyzes

their performance. Dynamic variation of system parameters and

overall performance are automatically captured during execution

and displayed on a color screen.

Accomplishment Description

AIMS was deployed and extended. AIMS has three major

software components: the source-code instrumenter, the runtime monitoring library, and the visualization tool set. The

source-code instrumenter inserted software probes directly into

the application with little or no intervention by the user. The

run-time monitoring library traced program execution. The

visualization tool set processed the execution data gathered and

displayed it using graphical representations. AIMS also provided

source-code click-back--one click on an animated view will

bring up a window with the program text. Major enhancements

to AIMS included: (1) self-identifying trace files, (2) an "instrument enabling profile" to turn on/off instrumentation without

recompiling the application software, (3) development of an

alpha version of AIMS for the CM-5.

Significance

The development of tools and methodologies for High Performance Computing and Communication Program applications.

Future Plans

AIMS will be available to the public by the end of August 1993.

l

Example of graphical interface for trace control, performance data animation, statistical data visualization, and source-code

click-back.

General/Computer Science 131
Computational Electromagnetics for Massively Parallel Processors

Michael J. Schuh, Principal Investigator

Co-investigators: Alex C. Woo and John D'Angelo

NASA Ames Research Center/General Electric Corporation

Research Objective

To develop electromagnetic scattering prediction codes for

modeling large scatterers that cannot be accurately modeled

by optical approximations or by low-frequency methods. The

algorithms used, finite-volume time domain (FVTD), finitedifference time domain, and nodal-based finite-element method

(FE), must be scalable because of the intense computational

requirements of this resonant scattering problem.

Approach

The three-dimensional (3-D) FVTD code, ARCCEM, was

implemented using FORTRAN 90. It used a multiblock structured grid to model complex geometries. A 3-D FE code, RF3D,

for the radar cross-section prediction of complex objects was

implemented on the Intel i860 Hypercube. This FE code used

a completely unstructured mesh, high-order isoparametric

elements, and a quasi-minimal residual complex matrix solver.

The code used various finite-element types including hexahedral, tetrahedral, prism, and pyramid elements.

Accomplishment Description

ARCCEM was upgraded from a two-dimensional code to a 3-D

code. A significant effort was devoted to obtaining Cray C-90

performance from the CM-5. Early indications show that Cray

C-90 performance levels will only be obtainable for larger

problems. Currently, small problems require 5-16 times more

processing time on 32 nodes of the CM-5 compared to the Cray

C-90. For implementation of the FE code on the Hypercube, a

recursive, binary domain decomposition was used. The domain

decomposition of the finite-element mesh was performed on a

host computer. The Hypercube implementation showed

increased speed corresponding to the number of processors.

However, the speed improvement dropped significantly when

the number of unknowns per processor dropped to less than

1,000 per node. A large problem consisting of approximately

240,000 complex unknowns required 208 seconds

of computation.

Outline of finite-element model of a perfect electrical conducting wedge cylinder; X = 3.5 wavelengths, Y = 1.0 wavelengths,

and Z = 4.0 wavelengths.

Significance

Computational methods from serial and vector machines will be

adapted to massively parallel machines for production problems

larger than can currently be solved.

Future Plans

Further efforts will be made to get a higher level of performance

from the CM-5 with the ARCCEM code. The code will be

enhanced by adding materials and mismatched grid capabilities.

Currently, a version of RF3D using the parallel virtual machine

communication software is being examined.

Finite-element method solution showing the scattered Y component of the H field. The incident field was a plane wave traveling

in the X direction and polarized in the Y direction. The scale of

the contours are 0.942 (red) to -I. 10 (blue) with the incident

field having a magnitude of 1.0. The mesh contained 3,808

second-order elements with 16,607 nodes.

132 General/Electromagnetics
Time-Domain Computational Electromagnetics

Vijaya Shankar, Principal Investigator

Co-investigators: William F. Hall, Alireza Mohammadian, Chris Rowell, and Michael Schuh

Rockwell International Science Center/NASA Ames Research Center

Research Objective

To advance the state of the art in electromagnetic computations,

especially the prediction of radar cross section (RSC) of low

observable targets, by solving the time-domain Maxwell's

equations accounting for all types of material properties.

Approach

The time-domain Maxwell's equations cast in differentialconservation form were solved using an upwind-based finitevolume scheme. The interface fluxes were computed using an

appropriate Riemann solver taking into account material

property variations from cell to cell. A structured-grid multiblock

procedure was employed to model arbitrarily shaped scattering

targets. The transient near-field solution was processed through a

fast-Fourier transform algorithm to obtain the RSC information as

a function of frequency.

Accomplishment Description

The time-domain Maxwell's solver RSC3D runs efficiently

in a vector/parallel architecture such as the Cray C-90. On a

16 processor C-90, it performs in excess of 10 GFLOPS. The

code was successfully tested for RSC prediction on several twoand three-dimensional targets specified by the Electromagnetic

Code Consortium. A typical scattering computation for a

complete fighter geometry is shown in the figure. The many

applications of this electromagnetic code include modeling

antenna problems and printed microstrip circuits and biomedical

simulations.

Significance

Accurate prediction of RSC for low observable targets in the

frequency range 1 O0 megahertz 20 gigahertz is a technology

that is critically needed for many defense projects. This electromagnetic code development activity, by employing advanced

algorithms and supercomputer architectures, is truly advancing

the state of the art in computational simulation.

Future Plans

Work is progressing to adapt this technology to massively

parallel computing architectures with a goal to provide teraflop

perforniance in the near future.

Radar cross-section studies and scattered field contours for a VFY218 fighter aircraft at 250 MHz.

General/Electromagnetics 133
Vortex-Ring/Free-Surface Interaction

Samuel Ohring, Principal Investigator

Co-investigator: Hans J. Lugt

Naval Surface Warfare Center

Research Objective

To understand vortex-ring interaction with a free surface (the

deformation and connection of a vortex ring during encounter

with a free surface), which is subject to deformation. This

problem will help in the interpretation of surface signatures

of wake flows.

Approach

The three-dimensional Navier-Stokes equations for a viscous,

incompressible fluid with full-nonlinear free-surface conditions

were solved with boundary-fitted coordinates, adaptive gridding,

and artificial compressibility applied at each physical time step.

This technique solves seven field equations for the three velocity

components, pressure, and three spatial coordinates. The

number of grid points is of the order of one million.

Accomplishment Description

Numerical computations for various Froude, Reynolds, and

Weber numbers were made with a thick vortex ring that formed

an inclination angle of 45 degrees with respect to the free

surface. When the vortex ring approached the free surface, an

upwelling above the ring's center and a half-moon-shaped ridge

developed, with a deep trough in between. Near the outer part

of the trough a circular dimple, signaling the connection of the

vortex ring with the free surface, was expected. The computations, however, revealed a new connection process that

deformed the vortex ring to a cylindrical vortex sheet. This

connection process was first explained for a flat, free surface

and then analyzed for a deformed, free surface, which generated

vorticity tangential to the free surface. This additional vorticity

formed secondary vortex rings of signs opposite to the primary

vortex. The procedure is a combination of vortex stretching,

helical vortex filaments generation, surface deformation,

and secondary vorticity production. The average case took

70 Cray Y-MP hours and used up to 32 megawords of memory.

Significance

These findings contribute to the physical understanding of

vortex-ring/free-surface interaction--particularly the intricate

recollection process--and may help to obtain a complete picture

of wave signatures.

Future Plans

Vortex rings of thinner cores and larger Reynolds numbers will

be computed to find the transition process from circular filament

reconnection to cylindrical sheet reconnection. A trial simulation is planned for the reflection of vortex rings at a free surface

for inclination angles smaller than 23 degrees.

Vorticity lines for a vortex ring in orthographic projection. The

color spectrum for the vorticity line magnitudes ranges from blue

(2 x 10 -5) to red (10.0) and the symmetry plane spectrum ranges

from blue (3 x 10-5) to red(5.0).

134 General/Fluid Mechanics
Computational Hydrodynamic

Donald W. Davis, Principal Investigator

Co-investigator: Keith C. Kaufman

General Dynamics, Electric Boat Division

Research Objective

To investigate and quantify the effects of Reynolds number

scaling on crossflow separation over unappended bodies of

revolution at nonzero yaw angles using a Reynolds-averaged

Navier-Stokes (RANS) solver.

Approach

The three-dimensional (3-D), steady-state, incompressible RANS

solver NFC3D was used to solve for the turbulent flow over a

body of revolution. Four Reynolds numbers were considered

(4,900,000, 13,000,000, 100,000,000, and 1 ,t00,000,000).

Results from the two lower Reynolds number computations

were compared to experimental data. Results from the higher

Reynolds number computations were compared to predictions

from semi-empirical and singularity based techniques.

Accomplishment Description

Turbulent flow computations were completed for a 9.5 lift-todrag ratio body of revolution at a 15-degree angle of yaw. Half

the symmetric physical domain was modeled. Computations

were performed on two eight-block computational grids (lower

Reynolds numbers = 814,200 grid points; higher Reynolds

numbers = 854,910 grid points). The second figure shows

crossflow perturbation velocities and streamlines at four body

stations. Preliminary comparisons with experimental results

show good correlation with body separation patterns and with

the size and extent of the leeside crossflow vortex.

Significance

Geometry input simulators that reliably and efficiently predict

maneuvering performance under all operating conditions are

needed to achieve optimal designs. Accurate hydrodynamic

representations for forces and moments in terms of geometry are

required. For bodies of revolution at nonzero angles of yaw,

vorticity shed from the body accumulates in the downstream

direction and can cause large nonlinear forces and moments on

the body. The inflow to the stern control appendages is also

affected by shed vorticity. Shed vorticity can be ingested by stern

propulsion systems. Because only limited full-scale data exist,

independent studies of these effects using 3-D RANS techniques

are necessary.

Performance Evaluation

(a) (b) (c) (d)

Body stations: (a) 0.10, (b) 0.25, (c) 0.75, and (d) 0.90.

(a) (b)

(c) (d)

Crossflow perturbation velocities and streamlines for Reynolds

numbers 13,000,000 ((a) and (c)) and 1,100,000,000 fib) and

(d)). Axial locations: (a) 0.10, (b) 0.25, (c) 0.75, and (d) 0.90.

Future Plans

Current efforts will be expanded to include the effects of steady

turns on bodies of revolution with and without appendages.

Investigated parameters will be: rotation rate, rotation radius,

angle of attack, angle of yaw, and the effects of fins mounted at

the midbody and stern. More sophisticated turbulence models

will be used to better account for rotation effects on the turbulent stresses.

General/Hydrodynamics 135
Small Airway Fluid Dynamics

Jeffrey R. Hammersley, Principal Investigator

Co-investigators: Rama Reddy, Dan E. Olson, Boyd Gatlin, and Joe F. Thompson

University of Arkansas/Medical College of Ohio/Mississippi State University

Research Objective

To study the airflow mechanics within deep-seated and experimentally inaccessible small respiratory bronchioles. Confined

bifurcating flows were examined within single two-dimensional

(2-D) and three-dimensional (3-D) bifurcation models to

determine the effects of variations in local geometry and grid

design.

Approach

Multiblocked 2-D and 3-D structured grids, which accurately

mimic the anatomic features of small airway bifurcations, were

developed using the EAGLE grid system. The computational

representations of the airways were studied over a biologically

realistic range of Reynolds numbers using the multiblocked

INS3D and the Taylor-Whitfield ABRAINS flow solvers.

Accomplishment Description

All 2-D computational models failed to reproduce accurately the

experimental analyses done in expanded-scale 3-D physical

models of similar design. The models appeared unable to handle

a confined channel flow with secondary vortex development.

The 3-D computational grid development was made difficult by

the complex, multiply curved carinal (flow-splitting) surfaces at

branch points. These surfaces required extensive grid manipulation to maintain smoothly curved branching and controlled

cross-sectional expansion. A series of 3-D grids were constructed, and early grid design results closely followed experimental measurements. Wall-shear analyses indicated outer wallflow separation and zones of increased walt shear downstream

of the carina.

Two-dimensional flow simulation through a planar symmetrical

double bifurcation model with non-dimensional velocities

displayed from 0.0 (black) to 1.0 (white). Inspiratory directed

flow Re = 500.

Significance

Medical scientists contemplate the localized delivery of

aerosolized viral vectors for replacement gene therapy. They

hope to quantify particulate deposition during environmental

exposures by linking the minor changes in airway geometry that

produce major disabilities such as asthma and emphysema. This

work provided realistic biologic flow simulations of fluidic

mixing, particle transport, and shear forces on airway walls.

Three-dimensional models of biologic structures are needed to

achieve reasonable simulations and specifically designed flow

solvers are needed for biologic flow regimes.

Future Plans

Investigation of symmetrically bifurcating flows will be extended

to asymmetrical bifurcations and serially branching 3-D cascades.

These efforts will set the stage for the incorporation of wall

elasticity and particle-flow interactions.

Pressure development in a single small airway bifurcation.

Three-dimensional flow simulation with inspiratory directed flow

Re = 500. Nondimensional pressures displayed from 0.0 (blue)

to 1.0 (red).

136 General/Life Science
Computer Simulation of the Structure and

Robert D. MacEIroy, Principal Investigator

Co-investigators: Andrew Pohorille and Michael Wilson

NASA Ames Research Center

Research Objective

To provide a molecular level description of the structure and

function of the earliest living cells (protocells). It is generally

assumed that protocells were closed spheroidal structures

immersed in water, with walls composed of simple membranes.

The study goal is to explain how interfaces between protocellular membranes and water could have promoted the

development of two essential cell functions: energy and

catalysis of chemical reactions.

Approach

Small solute molecule behavior that could have occurred in

primitive catalysis and energy storage was simulated at the

interface between a simple membrane and water using molecular dynamics methods. In this approach, Newton's equations of

motion for all atoms in the system were integrated numerically.

Results included statistical and dynamical properties of the

system, which could be directly compared with experimental

results and a detailed microscopic description of the system.

Accomplishment Description

Surface fluctuations in simple, fluid membranes were analyzed.

These flu(tuations obeyed the capillary wave model--a fundamental model describing liquid interfaces. As a consequence,

deep thinning defects were occasionally formed in the membrane. Unassisted transport of ions (sodium and chloride) across

membranes proceeded through the thinning defects and were

associated with considerable water penetration into the membrane. As a result, ion penetration across fluid memhranes was

nlarkedly increased compared to rigid membranes. This allowed

the formation of ion gradients which can be used to store energy

for chemi(al reactions. Many small molecules had free energy

minima at water membrane interfaces. The barriers to conformational (hanges of these molecules could be substantially lowered

at the interfaces. This effe(t was characteristic of catalyzed

reactions.

Significance

This work provides critical benchmark information needed for

understanding the origins of cellular life. It demonstrates (1) the

mechanisnl of charge (ion) transport across membranes in the

absence of specialized proteins present in contemporary cells,

and (2) the role of water-membrane interfaces as primitive

catalysts.

Function of Membranes

Future Plans

Protocell structure and function studies will be continued by

(1) assessing the feasibility of proton transport across membranes

via the "proton wire" mechanism, (2) investigating a model

electron transfer reaction at the water-membrane interface, and

(3) studying simple peptides at interfaces for likely candidates of

the earliest enzymes.

Ray-traced image of a sodium (Na+) ion perpetrating a water/

glycerol-1-monooleate membrane. The Na+ is yellow, the

oxygens and hydrogens of water are red and white, respectively;

the membrane polar head groups are magenta and the hydrocarbon tail groups are blue.

General/Life Science 137
Visualizing Neurons in Three Dimensions

Muriel D. Ross, Principal Investigator

Co-investigator: Kevin Montgomery

NASA Ames Research Center/Sterling Software Systems

Research Objective

To automate fully three-dimensional (3-D) reconstruction of

biological neuronal networks from serial sections digitized

directly from a transmission electron microscope (TEM).

Approach

A video camera and a personal computer captured and digitized

images from a TEM and then automatically sent the images to

a workstation. A mosaic of the images was created using a

connection machine to reproduce the original sections. The

mosaics were then viewed, enhanced, and enlarged on the

workstation. Contour extraction was a manual process and was

conducted by tracing objects displayed on the monitor. Algorithms were developed for contour registration, tesselation,

3-D reconstruction, and smoothing by computer.

Accomplishment Description

This is the first demonstration of a computerized method for

complex object reconstruction directly from a TEM. Computerized reconstruction methods reduce time requirements by

approximately 75 percent and have improved registration

compared to manual approaches. Eliminating conventional

photography results in a savings up to $5,000 per reconstruction.

Benefits have already included revision of previously held

theories of gravity sensor architecture, uncovering the fundamental organization of the sensors, and production of the first twodimensional simulations of neuronal endings in the sensors. To

create the mosaics, 90 hours of connection machine time were

used.

Significance

Computerized reconstruction of neurons and thorough simulations based on true architecture have been goals of neuroscientists since advanced computers appeared. If this goal is achieved,

the software designed in this study will greatly advance the fields

of neuroscience and neurocomputation. At NASA, this tool

enhances research into the effects of altered gravity on gravity

sensors.

Future Plans

Software development will be improved until the method is fully

automated and simulations of activity are transparent to the

reconstruction process.

Computer generated reconstruction of a nerve ending (yellow) in

a rat gravity sensor with enclosed gravity detecting neurons

(type I hair cells are gray).

138 General/Life Science
Large-Scale Numerical

Robert T. Whalen, Principal Investigator

Co-investigators: Marcus G. Pandy and Frank C. Anderson

NASA Ames Research Center/University of Texas at Austin

Research Objective

To assess the feasibility of computing large-scale optimal control

solutions for human movement using massively parallel and

vector-processing supercomputers. An optimal control problem

for walking was developed for musculoskeletal loading during

movement.

Simulations of Human Motion

Approach

A detailed three-dimensional model of the human body was

developed to simulate human motion on Earth and in space. The

skeleton was represented as an 8-segment, 19-degree-of-freedom

linkage, actuated by 46 muscles (first figure). Each muscle was

represented as a 3-element, lumped-parameter entity in series

with tendons.

can alter the functional properties of muscle and bone tissue.

The combination of musculoskeletal modeling and optimal

control theory has emerged as a powerful tool for determining

musculoskeletal forces during movement. Unfortunately,

accurate simulation of human motion and convergence to an

optimal control solution is computationally expensive. It is now

possible to use highly dimensioned musculoskeletal models to

simulate human movement and specify musculoskeletal load

histories.

Future Plans

A fully parallel computational algorithm for solving large-scale

optimal control problems on MIMD supercomputers will be

developed.

Accomplishment Description

The computational expense of solving the optimal control

problem was evaluated using a conventional serial machine, a

parallel-vector-processing machine, and a multiple-instruction/

multiple-data (MIMD) parallel machine. Computing the optimal

control solution on serial machines was impractical since

convergence required approximately three months of dedicated

processing time. Using either a parallel-vector-processing

machine or a Cray, a solution can be found with about

80 processing hours. The MIMD parallel machine performed

best in the calculation of the derivatives. This performance was

due to the fact that calculation of the derivatives scaled almost

linearly with the number of processors used (second figure). In

contrast, the Cray performed best during parameter optimization

of the controls, executing the parameter optimization routine

37 times faster than the serial machine. The ideal computer

architecture for solving very-large-scale optimal control problems appears to be a hybrid system in which a vector-processing

machine is integrated into the communication network of a

MIMD parallel machine. About 12 megabytes of memory are

needed to compute the optimal control solution for walking.

?

"O

Or)

128

96

64

32

0

m

m

B

;"-i__l II I I I

0 32 64 96

Number of processors

I I

128

Speed-up factor for computing derivatives for the walking

problem (dashed line) compared to ideal scaling (solid line) as

the number of processors increases.

Significance

The ability to simulate human movement and accurately

compute musculoskeletal load histories is important where

exposure to different load environments or gravitational fields

General/Life Science 139
Multidimensional Burner-Stabilized

K. Kailasanath, Principal Investigator

Co-investigator: G. Patnaik

Naval Research Laboratory/Berkeley Research Associates

Research Objective

To develop a better understanding of the effects of a burner on

the structure and stability of premixed flames.

Approach

A detailed, time-dependent, multidimensional, multispecies,

numerical model was used to simulate flames stabilized on a

burner. The model included finite-rate, detailed chemical

kinetics coupled with algorithms for convection, thermal

conduction, viscosity, molecular diffusion, and external forces

such as gravity. Boundary conditions were incorporated to

account for the heat lost to the burner.

Accomplishment Description

The effects of systematically changing the inflow velocity on

the structure and stability of multidimensional flames above a

burner were investigated. For large inflow velocities, the flame

structure was similar to that of a freely propagating flame.

However, as the inflow velocity was decreased, heat loss to

the burner increased and the cellular structure of the flame

disappeared. For very low inflow velocities, nearly flat flames

Flames

were obtained. The impact of different boundary conditions

(isothermal or Hirschfelder) were also studied. The figure

shows the structure of an 11 percent hydrogen-air flame on

an isothermal burner for 4 different inflow velocities. A typical

calculation required 30 Cray Y-MP hours.

Significance

This is the first multidimensional, detailed numerical simulation

of burner-stabilized flames. The simulations indicate that

obtaining a flat flame by adjusting the inflow velocity (as is often

done in experiments) might distort the true structure of the flame.

Future Plans

The code will be used to simulate and understand a number of

other interesting flame phenomena observed on Earth and in

microgravity.

Publication

Patnaik, G.; and Kailasanath, K.: Simulations of Multidimensional Burner-Stabilized Flames. AIAA Aerospace Sciences

Meeting, Reno, Nev., AIAA Paper 93-0241, Jan. 1993.

Effect of inflow velocity on an 11 percent hydrogen-air flame stabilized on an isothermal burner.

140 General/Reactive Flow
Protoplanetary Particle-Gas Dynamics

Jeffrey N. Cuzzi, Principal Investigator

Co-investigators: Anthony R. Dobrovolskis, Joelle M. Champney, and Robert C. Hogan

NASA Ames Resear(h Center/University of California, Santa Cruz/Synernet, Inc.

Research Objective

To understand the turbulent two-phase fluid dynamical environment in which the first solid aggregates were formed in the

protoplanetary nebula.

Approach

Vertical structure was modeled at various radii in a disk-shaped

gaseous nebula at a stage when particles of centimeters to meters

in size have accumulated and settled into a concentrated layer

very near the nebula's midptane. Using a Prandtl approach, gas

turbulence was modeled in the shear flows induced by orbital

velocity differences between the particles and the pressuresupported gas. Using a spatial-spectral direct numerical simulation code, particles of various Stokes numbers were introduced

into three-dimensional turbulen(e. Objective sampling and

visualization techniques were developed, allowing us to follow

the details of the particle clustering structure.

Accomplishment Description

There were substantial mean radial flows of gas and particles,

which help explain planetesimal growth and mixing of hightemperature and lower-temperature mineralogical phases.

Models were implemented with an array of particle sizes

evolving simultaneously. Each run required several Cray-2 hours

to reach convergence and 5 megawords of central memory.

The structure and evolution of particle clumps that evolve in

turbulence were determined. Particles evolved into small,

Kolmogorov-scale clumps, and then stretched out in one or

more dimensions (see figure). The clumps were persistent on a

time scale longer than an eddy turnover time. The runs shown in

the figure required 1.5 Cray-2 hours and 16 40 megawords of

central memory.

Significance

These are the only models in existence of realistic two-phase

fluid dynamics in the environment of the very earliest planetesimal and protoplanetary accretion.

Fulure Plans

Multiple-particle-size codes for different size distributions and

nebula environments will be run. Clumping calculations will be

pushed to higher Reynolds numbers. Correlations of particle

clump evolution with various particle Stokes numbers, and

velocity, pressure, and vorticity regimes in the turbulence will

be studied.

Five particle clumps at tile be/4inning (left) and end (ri,_ht) of a sequence lasting six eddy turnover periods. Particles are

colored to indicate the clump they belong to.

General/Space Science 141
Magnetoelectrohydrodynamics of Microgravity Crystal Growth

George S. Dulikravich, Principal Investigator

Co-investigator: Vineet Ahuja

Pennsylvania State University

Research Objective

To develop analytical models and software for simulation of

solidification under the influence of separately applied magnetic

and electric fields.

Approach

Both electrohydrodynamic and magnetohydrodynamic models

incorporated Lorentz force, joule heating, latent heat using

enthalpy methods, and thermally induced buoyancy via an

extended Boussinesq approximation allowing for separate

temperature-dependent physical properties in the melt and in

the solid. Explicit central differencing on boundary-conforming

non-orthogonal computational grids was used with four-step

Runge-Kutta time integration.

Accomplishment Description

Depending on the mutual orientation of the gravity, electric,

magnetic, and thermal gradient fields it was possible to significantly alter the solid/liquid interface shape, the amount of solid

accrued, thermal gradients in the solid, and the charged particle

deposition pattern in the solid.

Solid accrued on the walls and the velocity profiles in a turnaround channel with a magnetic field applied in the elbow

region.

Significance

This research indicates a possibility for using relatively weak

magnetic and electric fields to actively control the distribution

pattern of impurities and dopants and magnitudes of residual

thermal stresses in crystals.

Future Plans

All codes will be converted to time-accurate versions capable of

analyzing g-jitter and rotation.

Publications

1. Dulikravich, G. S.; Ahuja, V.; and Lee, S.: Three-Dimensional

Control of Crystal Growth Using Magnetic Fields. SPIE

Paper 1916-07, Feb. 1993.

2. Dulikravich, G. S.; Ahuja, V.; and Lee, S.: Three-Dimensional

Solidification with Magnetic Fields and Reduced Gravity.

AIAA Paper 93-0912, Jan. 1993.

A deformed thoroidal melt motion in a container with solid

accruing from the top wall. There is a downward centrally

located jet and upward corner jets.

142 General/Space Science
Computational Structural

Susan W. Bostic, Principal Investigator

NASA Langley Research Center

Mechanics Applications

Research Objective

To develop an efficient algorithm for nonlinear design sensitivity

analysis (DSA) on shared-memory, high-performance computers.

Approach

Nonlinear finite-element analysis was performed before

computing DSA. Parallel-vector versions of Newton Raphson

(N-R), Modified Newton Raphson (M N-R), and

Broyden-Fletcher-Goldfarb-Shann (BFGS) algorithms

were developed and compared for the nonlinear analysis

applications. The adjoint variable method was used for the

design sensitivity analysis. These nonlinear analyses and

the DSA algorithms were optimized for parallel-vector

supercomputers. In the adjoint variable method, each design

variable is assigned to a processor that computes the partial

derivatives of the internal loads with respect to the design

variable in parallel and with no synchronization. Each violated

displacement constraint is assigned to a processor that performs

a back substitution to obtain the adjoint vector. Then, each

processor multiplies the adjoining vector by the computed

partial derivative. To complete the DSA, each processor

computes the derivative of the constraints with respect to the

design variable.

Accomplishment Description

Three parallel-vector algorithms (N-R, M N-R, and BFGS) are

compared in the first figure. The N-R method outperforms the

M N-R and BFGS methods. The N-R method requires forming a

new stiffness matrix and a complete solution of the system of

equations at each iteration. A previously developed parallelvector algorithm, PVSOLVE, and the nodal algorithm for

generation and assembly of the stiffness matrix were incorporated into all three methods. The N-R method gained the most

from the previous algorithms and outperformed the other two

methods (first figure). Since little synchronization is required in

DSA, the computation time is reduced as the number of processors increases (second figure). The nonlinear analysis using the

N-R method shows an efficiency of 86 percent, the DSA shows

an efficiency of 98 percent, and the total solution efficiency is

96 percent.

Future Plans

The DSA algorithm will be implemented on a massively parallel

computer.

30

.__ 25

>,

-- _ 20

10

P

0

z 5

I Modified Newton-Raphson

(301 million operations)

I BFGS (354 million operations)

IWB Newton-Raphson

(438 million operations)

1 2 4 8

Number of Cray Y-MP processors

Central processing unit time for nonlinear analysis methods on

the Cray Y-MP.

80

A

_60 "_

>,40

°_

=o_20

z_

I-'7 Design sensitivity analysis

Nonlinear analysis

1 2 4 8

Number of Cray Y-MP processors

Central processing unit time for design sensitivity analysis on the

Cray Y-MP.

Significance

The N-R method on parallel-vector supercomputers reduces

the analysis time significantly, although the operation count is

higher than those of other methods (first figure). A DSA algorithm

resulted in a scalable reduction of time. Total analysis and DSA

time is reduced as the number of processors used increases.

This increase reduces the design cycle time.

General/Structural Mechanics 143
Simulation of Vertical Tail

Osama A. Kandil, Principal Investigator

Old Dominion University

Buffet

Research Objective

To simulate the buffet problem of fighter aircraft vertical tails

using simple models. The basic concept is to generate an

unsteady, vortex-breakdown flow and to place a cantilevered

vertical tail downstream of the vortex-breakdown flow. This way,

the buffet problem is isolated from the aircraft and the computational resources are focused on a small region for high resolution.

Approach

The multidisciplinary problem was solved successively at each

time step using three sets of equations for the unsteady vortexbreakdown flow, aerodynamic loads, aeroelastic deflections,

and grid displacements. For the unsteady aerodynamic loads,

the full Navier-Stokes equations were solved using an implicit,

upwind, flux-difference splitting, finite-volume scheme. For the

aeroelastic deflections, the aeroelastic equations for bending and

torsional vibrations were solved using the Galerkin method and

a four-stage Runge-Kutta scheme. For the grid displacements,

interpolation equations were solved to update the grid coordinates due to the tail aeroelastic deflections.

Accomplishment Description

Two models simulated the tail buffet problem. The first model

had a configured duct in which an inlet, supersonic, swirling

flow was forced to break down through a generated shock wave

at the inlet region. Downstream of the break-down flow, a

vertical cantilevered tail was placed and allowed to deform in a

bending mode. The second model had a delta-win_vertical-tail

configuration. Flow conditions were selected so that the wing

primary-vortex core experienced vortex break down before

reaching the tail location. The first figure shows a delta-wing/

vertical-tail configuration. The wing and the rectangular vertical

tail each have an aspect ratio 1. The wing angle of attack is

35 degrees, the free-stream Mach number is 0.4, and the

Reynolds number is 10,000; the tail thickness is 0.01 and its

width is 0.5. An O-H grid of 1 x 258 x 584 points in the

wraparound, normal, and axial directions was used to obtain the

solution. The spiral saddle points of vortex break down and the

surface pressure and the total-pressure surfaces at time step (t)

17,1 O0 (At = 0.003) are shown.

Significance

Numerical simulation of fighter aircraft vertical-tail buffet

problems was accomplished for the first time using simple

models. The models enabled researchers to investigate the buffet

problem, identify the critically dominant parameters, and find

methods to control buffet efficiently.

Future Plans

The combined bending and torsional modes response of the

vertical tails for both models will also be investigated.

Top view and three-dimensional view of the delta-wing/verticaltail configuration showing critical points of vortex breakdown,

surface pressure, and total-pressure surfaces (0.65).

144 General/Structural Mechanics
Buckling and Post-Buckling of Multilayered Composite Panels with Cutouts

Ahmed K. Noor, Principal Investigator

Co-investigator: Jeanne M. Peters

NASA Langley Research Center/University of Virginia

Research Objective

To study the buckling and post-buckling responses of multilayered composite panels with cutouts subjected to combined

mechanical and thermal loads.

Approach

The computational model of the composite panels was based

on the geometrically nonlinear shallow-shell theory, with the

effects of both transverse shear deformations and anisotropic

material behavior included. Analytic sensitivity coefficients

were evaluated, measuring the sensitivity of the post-buckling

response to variations in material and lamination parameters.

The panels were subjected to combined applied-edge shear

and either edge compressive loading or edge displacement. An

efficient reduced-basis computational procedure was used for

determining the stability boundary, generating the post-buckling

response, and evaluating the sensitivity derivatives. The procedure allowed a significant reduction in the number of degrees of

freedom used in the initial discretization.

Accomplishment Description

Extensive numerical studies were performed to understand the

effects of lamination parameters and the size of the cutout on the

stability boundary and post-buckling response of flat, unstiffened

panels. The figure shows the effect of hole size on the stability

boundary for a 16-layer quasi-isotropic panel. Also shown are

the sensitivity coefficients of the total strain energy with respect

to material parameters and fiber angles in the post-buckling

range for a panel subjected to combined edge shear and uniform

temperature change.

Significance

Future high-speed and high-performance aircraft will be

subjected to combined mechanical and thermal Ioadings. This

study will help in the lamination selection and identify the

material systems required.

Future Plans

Panel response to combined thermal and mechanical Ioadings

will be studied. The effect of temperature dependence of

material properties will also be studied. Stiffened panels with

cutouts will be considered. The numerical simulations will be

coupled with laboratory experiments to identify failure mechanisms and develop panel failure criteria.

Nt21 [N12,=lr..x

dlL'_ q.l lq°,crlm.x

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EL

ET

GLT

c_ L

c_ T

q.llq .... l.,.x

1,0

.8

.6

.4

.2

0.

.8

.6

N,2 L

.2

0.0.4(

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-20 0

au L

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The effect of hole size on the stability boundary for a 16-layer quasi-isotropic panel.

General/Structural Mechanics 145
Nonlinear Analysis of Damaged Stiffened-Fuselage Structure

James H. Starnes, Jr., Principal Investigator

Co-investigator: Vicki O. Britt

NASA Langley Research Center

Research Objective

Many commercial aircraft are operated beyond their designed

service life because it is more economical to maintain longservice aircraft than to replace them. The possibility of damage

to long-service aircraft makes important the structural integrity

and the safety issues associated with maintaining an aging fleet.

Reliable structural analysis is needed to determine the residual

strength of aircraft with cracks in the fuselage skin and other

structural elements.

Approach

A nonlinear shell analysis capability for stiffened shells with

propagating cracks was developed and implemented using the

STAGS (structural analysis of general shells) analysis code.,

Hierarchical stiffened-shell models of a fuselage shell skin were

developed to predict accurate local stress and deflection

gradients caused by frames, stringers, joints, and local cracks.

Accomplishment Description

The hierarchical scheme had three levels of finite-element

models, beginning with the nonlinear analysis of a large

stiffened-fuselage section subjected to internal pressure and a

mechanical bending load. Structural details such as frames,

stringers, tear straps, shear clips, floor beams, and stanchions

were included. Damage was represented in the model by a

longitudinal skin crack and a broken tear strap. Deflection

gradients were shown as deformed structural geometry plots,

and the hoop stress gradients were represented by the color

contours (see figures). Global model analysis displacements

at varying crack lengths were applied as boundary conditions

to the second level of modeling (a 6-bay by 6-bay stiffenedfuselage crown panel). The local model had higher mesh

refinement and modeling detail to characterize more accurately

the local structural response. Displacements from the 6-bay by

6-bay stiffened-panel model were applied as boundary conditions to a more refined 2-bay by 2-bay local stiffened-panel

model centered around the crack. The finer mesh better represented the local stress and deflection gradients in the crack

region. Stress intensity factors from this local model helped

determine the residual strength of the aircraft. A typical nonlinear analysis used 4 megawords of memory and 30-60 minutes

central processing unit time.

Significance

This method makes possible in-depth arialysis of the behavior of

a fuselage shell with a crack. Properties such as stress intensity

factors at the crack tip are determined and are used to assess

residual strength and determine how long an aircraft can be

flown before its residual strength becomes a concern.

Future Plans

This method will be expanded to include mechanical loads such

as vertical shear and torsion loads. Adaptive mesh refinement

capabilities will be added to STAGS to allow for curved crack

growth. Algorithms will be added to STAGS to determine when

and in what direction a crack will grow.

Hoop stresses in a global stiffened-fuselage shell model with a

skin crack and subjected to internal pressure and a mechanical

bending load. Blue = minimum stress; green, yellow, and

red = maximum stresses.

Hoop stresses in a local 2-bay by 2-bay stiffened-panel model

with a skin crack and subjected to internal pressure and a

mechanical bending load. Blue -- minimum stress; green, yellow,

and red = maximum stresses.

146 General/Structural Mechanics
Index

Index by Research Sites

Page

Analytical Services and Materials, Inc.

Richard L. Gaffney, Jr. Interaction of Turbulence and Chemical Reaction ............................................................................. 70

(with Institute for Computer Applications in Science and Engineering

and NASA Langley Research Center)

Charles R. McClinton Base Pressurization Methods for Scram jet Combustors ...................................................................... 78

(with NASA Langley Research Center)

C. D. Pruett Validation of the Parabolized Stability Equation Method ................................................................... 43

(with High Technology Corporation)

Veer N. Vatsa Block-Structured Approach for High-Lift Applications ....................................................................... 26

(with NASA Langley Research Center)

Thomas A. Zang Modeling for High-Speed Transitional Boundary Layers .................................................................... 49

(with NASA Langley Research Center)

Bell Helicopter Textron

J. C. Narramore Rotorcraft Drag Prediction ................................................................................................................ 97

Berkeley Research Associates

K. Kailasanath Multi-Dimensional Burner-Stabilized Flames .................................................................................. 140

(with Naval Research Laboratory)

Boeing Commercial Airplane Group

David E. Ashpis Transition in a Highly Disturbed Environment ................................................................................. 103

(with NASA Lewis Research Center)

H. C. Chen Inverse Design of Installed Nacelle ................................................................................................... 66

(with NASA Langley Research Center and ViGYAN, Inc.)

Effect of Suction Holes .................................................................................................................... 115 Philippe R. Spalart

Boeing Company, The

Frank Caradonna Computation of Helicopter Rotor and Wake Flows ............................................................................. 8

(with U.S. Army Aeroflightdynamics Directorate, AVSCOM, Flow Analysis, Inc.,

and Woodside Summit Group, Inc.)

Carnegie Mellon University

Wei J. Chyu Airframe/Inlet Aerodynamics ............................................................................................................. 67

(with NASA Ames Research Center and Cornell University)

Tom I-P. Shih Flow in Turbine-Blade Coolant Passages ........................................................................................... 84

(with NASA Lewis Research Center)

Center for Turbulence Research

J. J. Kim Numerical Simulation of Turbulence .............................................................................................. 106

(with NASA Ames Research Center)

Parviz Moin Drag Reduction Mechanism by Riblets ........................................................................................... 109

(with Stanford University and NASA Ames Research Center)

Parviz Moin Turbulent Flow Over a Backward Facing Step ................................................................................. 110

pi_,Clil)tN _& PAGE BLANK NOT FILMED

149
CFD Research Corporation

Clifford E. Smith

Cornell University

Wei J. Chyu

Eloret Institute

Jean-Luc Cambier

Brian L. Haas

William D. Henline

Richard L. Jaffe

Grant Palmer

Flow Analysis, Inc.

Frank Caradonna

Page

Jets-ln-Crossflow Mixing ................................................................................................................... 85

(with NASA Lewis Research Center)

Airframe/Inlet Aerodynamics ............................................................................................................. 67

(with NASA Ames Research Center and Carnegie Mellon University)

Pulsed Detonation Wave Augmentor Concept .................................................................................. 62

(with NASA Ames Research Center)

Particle Simulation of Hypersonic Rarefied Flow ............................................................................... 33

(with Stanford University and NASA Ames Research Center)

Computation of Space Transport Vehicle Re-Entry Heating ............................................................... 34

(with NASA Ames Research Center)

Propane-Air Combustion Mechanism ............................................................................................. 126

(with NASA Ames Research Center)

Hypersonic Computational Studies ................................................................................................... 41

(with NASA Ames Research Center)

Computation of Helicopter Rotor and Wake Flows ............................................................................. 8

(with U.S. Army Aeroflightdynamics Directorate, AVSCOM,

Woodside Summit Group, Inc., and The Boeing Company)

General Dynamics, Electric Boat Division

Donald W. Davis Computational Hydrodynamic Performance Evaluation .................................................................. 135

General Dynamics, Space Systems Division

Thanh T. Phan Launch Vehicle Base Flow Simulations ............................................................................................. 83

General Electric Aircraft Engines

Richard D. Cedar CFD for Engine-Airframe Integration ................................................................................................. 64

General Electric Corporation

Michael J. Schuh Computational Electromagnetics for Massively Parallel Processors .................................................. 132

(with NASA Ames Research Center)

General Motors Corporation, Allison Gas Turbine Division

Kurt F. Weber Chimera Domain Decomposition Applied to Turbomachinery Flow ................................................. 88

Kurt F. Weber Turbofan Compression System Simulation Using Chimera Domain Decomposition .......................... 89

George Washington University

Rainald Lohner Finite-Element Euler Solver for a Distributed Memory Computer ....................................................... 19

150
Page

Georgia Institute of Technology

Suresh Menon Parallel Simulation of Unsteady Combustion .................................................................................... 80

(with NASA Ames Research Center)

Suresh Menon Unsteady Combustion in a Ramjet ....... : ............................................................................................ 81

Grumman Corporate Research Center

Michael J. Siclari Sonic Boom Predictions for High-Speed Civil Transport .................................................................... 24

(with Lockheed Engineering and Sciences Company)

High Technology Corporation

Analysis and Design of High-Lift Systems .......................................................................................... 52

(with NASA Langley Research Center)

Validation of the Parabolized Stability Equation Method ................................................................... 43

(with Analytical Services and Materials, Inc.)

Development of a Turbulent Spot ................................................................................................... 114

(with NASA Langley Research Center)

Anutosh Moitra

C. D. Pruett

Bart A. Singer

ICASE

Thomas W. Crockett

ICOMP

John J. Adamczyk

Chunill Hah

Kai-Hsiung Kao

Eli Turkel

Shaye Yungster

Imperial College

Rajiv Thareja

Performance and Scalability of Parallel Graphics Libraries .............................................................. 130

Simulation of Turbomachinery Flows ................................................................................................ 59

(with NASA Lewis Research Center and Sverdrup Technology, Inc.)

Three-Dimensional Transonic Compressor Stage Flows .................................................................... 72

(with NASA Lewis Research Center and WL/FIMM, Wright Patterson AFB)

Simulation of a Turbine-Blade Cooling Passage .............................................................................. 105

(with NASA Lewis Research Center )

Navier-Stokes Equations for Jet Acoustics ......................................................................................... 25

(with NASA Lewis Research Center)

National Aero-Space Plane External Burning Nozzle Studies ............................................................ 92

(with NASA Lewis Research Center )

Adaptive Unstructured Hypersonic Multigrid Solver ......................................................................... 46

(with Lockheed Engineering and Sciences Company, University of Wales,

and Massachusetts Institute of Technology)

Institute for Computer Applications in Science and Engineering

Richard L. Gaffney, Jr.

Iowa State University

Nateri K. Madavan

Interaction of Turbulence and Chemical Reaction ............................................................................. 70

(with Analytical Services and Materials, Inc. and NASA Langley Research Center)

Unsteady Multistage Turbomachinery Applications .......................................................................... 77

(with MCAT Institute and NASA Ames Research Center)

151
Page

JetPropulsionLaboratory

ChristianL.Keppenne Orographically ForcedOscillationsintheMartianAtmosphere...................................................... 123

LawrenceBerkeleyLaboratory

Jerryc.Yan Performance MonitoringofParallel Programs............................................................................131

(withRecomTechnologies andNASAmesResearchCenter)

LockheedEngineeringandSciencesCompany

JeffreyA.Catt EvaluatingNozzleDragUsingComputational FluidDynamics......................................................... 63

DougG.Howlett EvaluatingHigh-Angle-of-Attack InletCharacteristics ....................................................................73

LawrenceD.Huebner PoweredHypersonicAir-Breathing Configuration Studies...............................................................35

(withNASALangleyResearchCenter)

SteveL.Karman, Jr. Unstructured Grid/FlowSolverCalibration...............................................................................17

RamadasK.Prabhu Computational StudiesofNonequilibriumFlows........................................................................42

(withNASALangleyResearchCenter)

DavidM.Schuster WingswithControl Surfaces................................................................................................23

Michael J.Siclari SonicBoomPredictions forHigh-Speed CivilTransport .................................................................24

(withGrummanCorporateResearchCenter)

RajivThareja AdaptiveUnstructured Hypersonic MultigridSolver.....................................................................46

(withUniversityofWales,Massachusetts InstituteofTechnology,

andImperial College)

LTV,Aircraft Division

PerryA.Wooden MultifunctionPropulsionSystems..........................................................................................90

MassachusettsInstitute of Technology

c. s.Tan FlowPhenomena iTurbomachinery .....................................................................................86

RajivThareja AdaptiveUnstructuredHypersonic MuItigridSolver.....................................................................46

(withLockheedEngineering andSciencesCompany, UniversityofWales,

andImperial College)

MCAT Institute

JorgeE.Bardina

ThomasA.Edwards

KarenL.Gundy-Burlet

ScottULawrence

NateriK.Madavan

Three-Dimensional Compressible TurbulenceModeling................................................................30

(withNASAAmesResearchCenter)

Aerodynamic OptimizationfSupersonic Aircraft.......................................................................12

(withNASAAmesResearchCenter)

UnsteadyTurbomachinery Computations .................................................................................71

(withNASAAmesResearchCenter)

IntegratedHypersonic VehicleSimulation................................................................................37

(withNASAAmesResearchCenter)

UnsteadyMultistageTurbomachinery Applications.....................................................................77

(withNASAAmesResearchCenterandIowaStateUniversity)

152
Page

McDonnell Aircraft Company

Dharmanshu L. Antani CFD Applications to HSCT Design and Analysis ............................................................................... 55

(with NASA Ames Research Center)

David Halt F/A-18 Store Carriage Analyses ......................................................................................................... 15

Linda D. Kral Receptivity, Transition, and Turbulence Phenomena ....................................................................... 107

Linda D. Kral Turbulence Modeling for Three-Dimensional Flow Fields ............................................................... 108

McDonnell Douglas Aerospace

Alan B. Cain Simulation of Supersonic Jet Screech .................................................................................................. 6

McDonnell Douglas Corporation

Ramesh K. Agarwal Numerical Solution of Three-Dimensional Maxwell Equations ............................................................ 3

Pieter G. Buning Subsonic Transport Propulsion/Airframe Integration .......................................................................... 61

(with NASA Ames Research Center)

C. C. Lee Efficiency Improvements of Boundary Layer Transition Predictions ................................................... 38

H. Tadghighi Simulation of a Complex Three-Dimensional Internal Flow Field ...................................................... 99

McDonnell Douglas Research Laboratories

Ramesh K. Agarwal Flow Fields Including Chemical Kinetics ........................................................................................... 29

Medical College of Ohio

Jeffrey R. Hammersley Small Airway Fluid Dynamics ......................................................................................................... 136

(with University of Arkansas and Mississippi State University)

Mississippi State University

Jeffrey R. Hammersley Small Airway Fluid Dynamics ......................................................................................................... 136

(with University of Arkansas and Medical College of Ohio)

J. Mark Janus Computational Analysis of a Tip-Engine Configuration ...................................................................... 74

NASA Ames Research Center

Dharmanshu L. Antani CFD Applications to HSCT Design and Analysis ............................................................................... 55

(with McDonnell Aircraft Company)

Jorge E. Bardina Three-Dimensional Compressible Turbulence Modeling ................................................................... 30

(with MCAT Institute)

Timothy J. Barth A Three-Dimensional Implicit Unstructured Euler Solver .................................................................... 5

(with Sterling Software Systems)

Pieter G. Buning Subsonic Transport Propulsion/Airframe Integration .......................................................................... 61

(with McDonnell Douglas Corporation)

Jean-Luc Cambier Pulsed Detonation Wave Augmentor Concept .................................................................................. 62

(with EIoret Institute)

Neal M. Chaderjian High-Incidence Static/Dynamic Roll Computations ............................................................................. 9

(with Naval Post Graduate School and Stanford University)

C. L. Chen Computation of Separated Nozzle Flows .......................................................................................... 65

(with Rockwell International Science Center)

153
WeiJ.Chyu

JeffreyN.Cuzzi

ThomasA.Edwards

KarenL.Gundy-Burlet

BrianL.Haas

WilliamD.Henline

RichardL.Jaffe

J.J.Kim

StephenP.Klotz

StephenLanghoff

ScottL.Lawrence

NateriK.Madavan

RobertD.MacEIroy

W.J.McCroskey

W.J.McCroskey

SureshMenon

ParvizMoin

ParvizMoin

GrantPalmer

JamesB.Pollack

YehiaRizk

Michael M.Rogers

MurielD.Ross

Page

Airframe/Inlet Aerodynamics ................................................................................................67

(withCarnegieMellonUniversityandCornellUniversity)

Protoplanetary Particle-Gas Dynamics.....................................................................................141

(withUniversityofCalifornia, SantaCruzandSynernet, Inc.)

Aerodynamic OptimizationfSupersonic Aircraft.......................................................................12

(withMCATInsitute)

UnsteadyTurbomachinery Computations .................................................................................71

(withMCATInstitute)

ParticleSimulationofHypersonicRarefiedFlow..........................................................................33

(withEloretInstituteandStanfordUniversity)

Computation ofSpaceTransport VehicleRe-EntryHeating..............................................................34

(withEIoretInstitute)

Propane-Air Combustion Mechanism......................................................................................126

(withEloretInstitute)

Numerical SimulationfTurbulence......................................................................................106

(withCenterforTurbulenceResearch)

SimulationofFlowAboutanAirborneObservatory ......................................................................18

BoostPhaseDetectionStudies..............................................................................................127

IntegratedHypersonicVehicleSimulation.................................................................................37

(withMCATInstitute)

UnsteadyMultistageTurbomachinery Applications......................................................................77

(withMCATInstituteandIowaStateUniversity)

Computer SimulationoftheStructureandFunctionofMembranes ................................................. 137

Aerodynamics andAcousticsofRotorcraft .................................................................................95

Calculations ofHigh-Performance Rotorcraft ..............................................................................96

Parallel SimulationofUnsteadyCombustion..............................................................................80

(withGeorgiaInstituteofTechnology)

DirectComputation ofAerodynamic SoundGeneration.................................................................21

(withStanfordUniversity)

DragReductionMechanismbyRiblets....................................................................................109

(withCenterforTurbulenceResearch andStanfordUniversity)

HypersonicComputational Studies.........................................................................................41

(withEloretInstitute)

DynamicsoftheMartianAtmosphere ......................................................................................124

Computational FluidDynamicsofF-18FlowField.......................................................................22

ForcedPlaneMixingLayers.................................................................................................113

(withStanfordUniversityandUniversityofWisconsin, Madison)

VisualizingNeuronsinThreeDimensions.................................................................................138

(withSterlingSoftwareSystems)

154
Page

JeffreyD.Scargle Parallel IntegrationfN-bodyGravitational Systems ....................................................................... 120

LewisB.Schiff AircraftForebodyFlowControl Technology .................................................................................53

Michael J.Schuh Computational Electromagnetics forMassivelyParallel Processors .................................................. 132

(withGeneral ElectricCorporation)

VijayaShankar Time-Domain Computational Electromagnetics ............................................................................133

(withRockwell International ScienceCenter)

BruceF.Smith Numerical Experiments inheFormationandEvolutionofGalaxies............................................... 121

(withUniversityofChicagoandUniversityofCalifornia, SantaCruz)

Robert T.Whalen Large-Scale Numerical Simulations ofHumanMotion.................................................................... 139

(withUniversityofTexasatAustin)

Michael A.Wilson Computer SimulationofAstrophysical Ices.................................................................................128

JerryC.Yan Performance MonitoringofParallel Programs..............................................................................131

(withRecomTechnologies andLawrenceBerkeleyLaboratory)

RichardE.Young SimulationofVolcanicAerosol Clouds......................................................................................125

S.H.KonradZhu DSMCSimulationofHigh-AltitudePlumeInteraction.....................................................................50

(withRockwell International, Rocketdyne Division)

NASA Johnson Space Center

Fred W. Martin, Jr. Space Shuttle Flow Field ................................................................................................................... 20

NASA Langley Research Center

Harold L. Atkins Simulation of Acoustic Scatter ............................................................................................................. 4

Susan W. Bostic Computational Structural Mechanics Applications .......................................................................... 143

Richard L. Campbell Efficient Constrained Aerodynamic Design ......................................................................................... 7

H. C. Chen Inverse Design of Installed Nacelle ................................................................................................... 66

(with Boeing Commercial Airplane Group and ViGYAN, lnc.l

Charles E. Cockrell, Jr. Waverider and Reference Model Computational Study ..................................................................... 31

J. Philip Drummond High-Speed Turbulent Reacting Flows ............................................................................................ 104

(with State University of New York at Buffalo)

Richard L. Gaffney, Jr. Interaction of Turbulence and Chemical Reaction ............................................................................. 70

(with Analytical Services and Materials, Inc. and Institute for

Computer Applications in Science and Engineeringt

Peter A. Gnoffo Reentry Heating Effects on the Shuttle Orbiter ................................................................................... _2

Lawrence D. Huebner Powered Hypersonic Air-Breathing Configuration Studies ................................................................. 35

(with Lockheed Engineering and Sciences Companyl

Kenneth M. Jones Subsonic High-Lift Analysis ............................................................................................................... 51

(with ViGYAN, Inc.)

Ajay Kumar Computational Performance Enhancement of a Scram jet Inlet .......................................................... 36

Charles R. McClinton Base Pressurization Methods for Scramjet Combustors ...................................................................... 78

(with Analytical Services and Materials, Inc.)

155
Page

CharlesR.McClinton Performance Potential ofFull-ScaleInjectors.............................................................................40

CharlesR.McClinton SimulationofScramjet Combustor FlowField.............................................................................79

AnutoshMoitra AnalysisandDesignofHigh-Lift Systems..................................................................................52

(withHighTechnology Corporation)

AhmedK.Noor BucklingandPost-Buckling ofMultilayeredCompositePanelswithCutouts................................... 145

(withUniversityofVirginia)

JohnC.Otto Three-Dimensional Parallel MixingComputations ........................................................................82

RamadasK.Prabhu Computational StudiesofNonequilibriumFlows.........................................................................42

(withLockheedEngineering andSciencesCompany)

ManMohanRai DirectSimulations ofAirfoilFlows..........................................................................................112

R.ClaytonRogers FuelPlumeImaging/CFDComparisons ....................................................................................45

BartA.Singer Development ofaTurbulent Spot..........................................................................................114

(withHighTechnology Corporation)

JamesH.Starnes, Jr. Nonlinear AnalysisofDamagedStiffened-Fuselage Structure.......................................................... 146

RichardA.Thompson Aerothermodynamic Benchmark forCandidateAMLSVehicle.......................................................... 47

VeerN.Vatsa Block-Structured ApproachforHigh-Lift Applications ....................................................................26

(withAnalytical ServicesandMaterials, Inc.)

K.JamesWeilmuenster Navier-Stokes Simulations ofOrbiterAerodynamic Characteristics .................................................. 48

ThomasA.Zang ModelingforHigh-Speed Transitional BoundaryLayers..................................................................49

(withAnalytical ServicesandMaterials, Inc.)

NASA Lewis Research Center

Simulation of Turbomachinery Flows ................................................................................................ 59

(with Sverdrup Technology, Inc. and ICOMP)

Transition in a Highly Disturbed Environment ................................................................................. 103

(with Boeing Commercial Airplane Group)

Investigation of Scalability of Parallel Turbomachinery Codes ........................................................... 60

Three-Dimensional Transonic Compressor Stage Flows .................................................................... 72

(with ICOMP and WL/FIMM, Wright Patterson AFB)

Simulation of a Turbine-Blade Cooling Passage .............................................................................. 105

(with ICOMP)

Flow in Turbine-Blade Coolant Passages ........................................................................................... 84

(with Carnegie Mellon University)

Jets-ln-Crossflow Mixing ................................................................................................................... 85

(with CFD Research Corporation)

Navier-Stokes Equations for Jet Acoustics ......................................................................................... 25

(with ICOMP)

National Aero-Space Plane External Burning Nozzle Studies ............................................................ 92

(with ICOMP)

John J. Adamczyk

David E. Ashpis

Richard A. Blech

Chunill Hah

Kai-Hsiung Kao

Tom I-P. Shih

Clifford E. Smith

Eli Turkel

Shaye Yungster

156
Page

OsamaA.Kandil

OsamaA.Kandil

OregonStateUniversity

NASAMarshall SpaceFlightCenter

PaulMcConnaughey SpaceStationFlowAnalysis..................................................................................................119

(withSverdrupTechnology, Inc.)

Naval PostGraduateSchool

NealM.Chaderjian High-Incidence Static/Dynamic RollComputations ........................................................................9

(withNASAAmesResearchCenterandStanfordUniversity)

Naval ResearchLaboratory

K.Kailasanath Multidimensional Burner-Stabilized Flames................................................................................140

(withBerkeleyResearchAssociates)

CarolynR.Kaplan LiftedEthyleneJetDiffusionFlames.........................................................................................16

Naval SurfaceWarfare Center

Samuel Ohring Vortex-Ring/Free-Surface Interaction........................................................................................134

Northrop Corporation

RichardD.Crouse CavityAeroacoustic LoadsPrediction.......................................................................................10

Old Dominion University

SimulationofVertical TailBuffet............................................................................................144

TransonicFlowAroundaDeltaWing......................................................................................56

JeffreyR.Barnes Numerical Simulations ofBaroclinicInstability............................................................................122

PennsylvaniaStateUniversity

GeorgeS.Dulikravich Magnetoelectrohydrodynamics ofMicrogravityCrystal Growth...................................................... 142

RecomTechnologies

Jerryc. Yan Performance MonitoringofParallel Programs.............................................................................131

(withNASAmesResearchCenterandLawrenceBerkeleyLaboratory)

RIACS

RogerC.Strawn DynamicAdaptionforThree-Dimensional Unstructured Grids......................................................... 98

(withStanfordUniversity)

Rockwell International Science Center

C. L. Chen Computation of Separated Nozzle Flows .......................................................................................... 65

(with NASA Ames Research Center)

Vijaya Shankar Time-Domain Computational Electromagnetics .............................................................................. 133

(with NASA Ames Research Center)

Rockwell International, North American Aircraft Division

Philip B. Gingrich Design-By-Optimization Method ...................................................................................................... 14

Jong H. Wang Hydrocarbon Scram jet Combustor Flows .......................................................................................... 87

157
DavidT.Yeh

DavidT.Yeh

Rockwell International,

E. D. Lynch

S. H. Konrad Zhu

Rockwell International,

Daniel F. Dominik

K. Kurian Mani

Page

Numerical Study of Thrust Vectoring Characteristics ......................................................................... 91

Steady and Unsteady Flows at High Angles of Attack ........................................................................ 54

Rocketdyne Division

Turbulent Base Heating Computational Fluid Dynamics ................................................................... 76

DSMC Simulation of High-Altitude Plume Interaction ....................................................................... 50

(with NASA Ames Research Center)

Space Systems Division

High-Fidelity Space Shuttle Simulation ............................................................................................. 11

Flow Past a Dynamic Multiple Body System ..................................................................................... 39

Frederik J. de Jong

Stanford University

Neal M. Chaderjian

Science Applications International Corporation

Sanford M. Dash Three-Dimensional High-Speed Plume-Propulsive Flow Fields ......................................................... 68

Scientific Research Associates, Inc.

Turbine-Blade Tip Clearance Flows .................................................................................................. 69

High-Incidence Static/Dynamic Roll Computations ............................................................................. 9

(with NASA Ames Research Center and Naval Post Graduate School)

Brian L. Haas Particle Simulation of Hypersonic Rarefied Flow ............................................................................... 33

(with EIoret Institute and NASA Ames Research Center)

Parviz Moin Direct Computation of Aerodynamic Sound Generation ................................................................... 21

(with NASA Ames Research Center)

Parviz Moin Drag Reduction Mechanism by Riblets ........................................................................................... 109

(with Center for Turbulence Research and NASA Ames Research Center)

J. Blair Perot Turbulence Simulation on the Connection Machine ....................................................................... 111

Michael M. Rogers Forced Plane Mixing Layers ............................................................................................................ 113

(with NASA Ames Research Center and University of Wisconsin, Madison)

Roger C. Strawn Dynamic Adaption for Three-Dimensional Unstructured Grids ......................................................... 98

(with RIACS)

State University of New York at Buffalo

J. Philip Drummond High-Speed Turbulent Reacting Flows ............................................................................................ 104

(with NASA Langley Research Center)

Sterling Software Systems

Timothy J. Barth A Three-Dimensional Implicit Unstructured Euler Solver .................................................................... 5

(with NASA Ames Research Center)

Muriel D. Ross Visualizing Neurons in Three Dimensions ...................................................................................... 138

(with NASA Ames Research Center)

158
Page

Sverdrup Technology, Inc.

John J. Adamczyk Simulation of Turbomachinery Flows ................................................................................................ 59

(with NASA Lewis Research Center and ICOMP)

Jinho Lee Studies of Mixing and Combustion in Supersonic Flows .................................................................... 75

Paul McConnaughey Space Station Flow Analysis ............................................................................................................ 119

(with NASA Marshall Space Flight Center)

Synernet, Inc.

Jeffrey N. Cuzzi Protoplanetary Particle-Gas Dynamics ............................................................................................ 141

(with NASA Ames Research Center and University of California, Santa Cruz)

Teledyne Brown Engineering

R. P. Roger Jet Interaction Aero-Optic Effects on Hypersonic Interceptors ........................................................... 44

University of Arkansas

Jeffrey R. Hammersley

Jeffrey N. Cuzzi

Bruce F. Smith

University of Chicago

Bruce F. Smith

Small Airway Fluid Dynamics ......................................................................................................... 136

(with Medical College of Ohio and Mississippi State University)

University of California, Santa Cruz

Protoplanetary Particle-Gas Dynamics ............................................................................................ 141

(with NASA Ames Research Center and Synernet, Inc.)

Numerical Experiments in the Formation and Evolution of Galaxies ............................................... 121

(with NASA Ames Research Center and University of Chicago)

Numerical Experiments in the Formation and Evolution of Galaxies ............................................... 121

(with NASA Ames Research Center and University of California, Santa Cruz)

University of Illinois, Urbana/Champaign

James P. Wolfe Radiation Effects in Intermetallic Compounds ................................................................................. 129

University of Texas at Austin

Robert T. Whalen Large-Scale Numerical Simulations of Human Motion .................................................................... 139

(with NASA Ames Research Centerl

University of Virginia

Ahmed K. Noor

University of Wales

Rajiv Thareja

Buckling and Post-Buckling of Multilayered Composite Panels with Cutouts ................................... 145

(with NASA Langley Research Center)

Adaptive Unstructured Hypersonic Multigrid Solver ......................................................................... 46

(with Lockheed Engineering and Sciences Company,

Massachusetts Institute of Technology, and Imperial College)

University of Wisconsin, Madison

Michael M. Rogers Forced Plane Mixing Layers ............................................................................................................ 113

(with NASA Ames Research Center and Stanford University)

159
U.S. Army Aeroflightdynamics Directorate, AVSCOM

Frank Caradonna

ViGYAN, Inc.

H. C. Chen

Page

Computation of Helicopter Rotor and Wake Flows ............................................................................. 8

(with Flow Analysis, Inc., Woodside Summit Group, Inc.,

and The Boeing Company)

Inverse Design of Installed Nacelle ................................................................................................... 66

(with Boeing Commercial Airplane Group and NASA Langley Research Center)

Kenneth M. Jones Subsonic High-Lift Analysis ............................................................................................................... 51

(with NASA Langley Research Center)

WL/FIMM, Wright Patterson AFB

Datta Gaitonde Three-Dimensional Turbulent Double-Fin Interactions ...................................................................... 13

Chunill Hah Three-Dimensional Transonic Compressor Stage Flows .................................................................... 72

(with NASA Lewis Research Center and ICOMP)

Woodside Summit Group, Inc.

Frank Caradonna Computation of Helicopter Rotor and Wake Flows ............................................................................. 8

(with U.S. Army Aeroflightdynamics Directorate, AVSCOM,

Flow Analysis, Inc., and The Boeing Company)

16O
Form Approved

REPORT DOCUMENTATION PAGE OMeNoo7o4-o18e

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,

gathering and maintaining the data needed, and completing and reviewing the collection of information Send comments regarding this burden estimate or any other aspect of this

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for information Operations and Reports, 1215 Jefferson

Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3.

March 1994

4. TITLE AND SUBTITLE

NAS Technical Summaries, March 1992-February 1993

6. AUTHOR(S)

Numerical Aerodynamic Simulation Program

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

Ames Research Center

Moffett Field, CA 94035-1000

9. SPONSORING/MONITORING AGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

REPORT TYPE AND DATES COVERED

Reference Publication

5. FUNDING NUMBERS

536-01-11

8. PERFORMING ORGANIZATION

REPORT NUMBER

A-94015

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA RP-1 321

11. SUPPLEMENTARYNOTES

Point of Contact: Pat Elson, Ames Research Center, MS 258-5, Moffett Field, CA 94035-1000;

(415) 604-4463

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Unclassified -- Unlimited

Subject Category 99

Available from the NASA Center for AeroSpace Information,

800 Elkridge Landing Road, Linthicum Heights, MD 21090; (301) 621-0390

13. ABSTRACT (Maximum 200 words)

NASA created the Numerical Aerodynamic Simulation (NAS) Program in 1987 to focus resources on

solving critical problems in aeroscience and related disciplines by utilizing the power of the most advanced

supercomputers available. The NAS Program provides scientists with the necessary computing power to solve

today's most demanding computational fluid dynamics problems and serves as a pathfinder in integrating

leading-edge supercomputing technologies, thus benefiting other supercomputer centers in government and

industry. The 1992-93 Operational Year concluded with 399 high-speed processor projects and 91 parallel

projects representing NASA, the Department of Defense, other government agencies, private industry, and

universities. This document provides a glimpse at some of the significant scientific results for the year.

14. SUBJECT TERMS

Computational fluid dynamics, Supercomputing technology, Cray-2,

Cray Y-MP, High-speed processing, Parallel processing

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION

OF REPORT OF THIS PAGE OF ABSTRACT

Unclassified Unclassified

NSN 7540-01-280-5500

15. NUMBER OF PAGES

176

16. PRICE CODE

A08

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102

_U.S. GOVERNMENT PRINTING OFFICE: 1994-584-277