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NAS 

Technical 

Summaries 

Numerical Aerodynamic 

Simulation Program 

March 1991 - February 1992

- 1t f
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 this cover 

would be virtually incomprehensible.
10x^

 - /0 

National Aeronautics and W€t' 

Space Administration 

Ames Research Center 

Moffett Field, California 94035-1000 

(NASA-TM-109335) NAS TECHNICAL N94-15798 

SUMMARIES: NUMERICAL AERODYNAMIC 

SIMULATION PROGRAM, MARCH 1991 - 

FEBRUARY 1992 (NASA) 266 p Unclas 

G3/01 0191137 

NAS 

Technical 

Summaries 

Numerical Aerodynamic 

Simulation Program 

March 1991 - February 1992
0 Aerodynamics 

0 Atmospheric Sciences 

• Hydrodynamics 

[] Hypersonics 

E Physics 

Propulsion 

Structural Mechanics 

o Turbulence 

II Other 

cI 

C Sd.na 

• E1.d,t Engln..rb.ç 

• LH. Sd.nc. 

• Projects 

o Accounts 

NAS projects and accounts. 

87-88 88-89 89-90 90-91 91-92 

2000 

1000

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. 

This report contains selected scientific results from the 1991-92 

NAS Operational Year, March 4, 1991 to March 3, 1992, which 

is the fifth year of operation. During this year, the scientific 

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

The Cray-2, the first generation supercomputer, has four 

processors, 256 megawords of central memory, and a total 

sustained speed of 250 million floating point operations per 

second. The Cray Y-MP, the second generation supercomputer, 

has eight processors and a total sustained speed of one billion 

floating point operations per second. Additional memory was 

installed this year, doubling capacity from 128 to 256 megawords 

of solid-state storage-device memory. Because of its higher 

performance, the Cray Y-MP delivered approximately 77% of 

the total number of supercomputer hours used during this year. 

The DD40 disk drives used for scratch space were replaced by 

DD4Rs, which provide improved reliability and increased 

storage capacity. Software improvements were made, including 

CRI's Data Migration Facility to improve archiving and the NASdeveloped Session Reservable File System. The NAStore system

supported researchers' mass storage needs by providing the 

researchers with a virtually unlimited supply of very fast disk 

space. 

In the first three years of the Program, the number of research 

projects and the number of scientific researchers increased 

steadily. However, in the fourth year there was a shift in the 

resource allocation policy. That year, review committees 

approved "fewer but bigger projects," although the demand for 

NAS resources increased dramatically. This shift continued 

through this year. Increasing the amount of computation time 

allocated to specific projects allows scientists to direct research 

toward more computationally intensive efforts and encourages 

multidisciplinary research. 

By the end of the 1991-92 NAS Operational Year, 1,420 researchers conducted 391 research projects at 140 sites representing 

NASA, the Department of Defense, other Government agencies, 

private industry, and universities. 

The NAS Technical Summaries Report provides an overview of 

the significant scientific results for the year. Additional copies of 

the report for this year and previous years are available on 

request. Please address your request to 

NAS Documentation Center 

NASA Ames Research Center 

Mail Stop 258-6 

Moffett Field, CA 94035-1000 

(415) 604-4632

Project disciplines for NAS Operational Year 1991-92. 

PREcEDNG PAGE BLANK NOT FILMED

PAGE J\ ITETOALLY BLANK
Table of Contents 

Principal 

Investigator NAS Summary Page 

Ramesh K. Agarwal Numerical Solution of Three-Dimensional Maxwell's Equations.............................................................1 

Co-Investigators: Mark Shu and Mark R. Axe 

McDonnell Douglas Research Laboratories 

Ramesh K. Agarwal Helicopter Fuselage, Rotor, and Rotor/Body Interaction Flow-Field Calculations .................................... 2 

Co-Investigator: Jerry E. Deese 

McDonnell Douglas Research Laboratories 

Shreekant Agrawal Numerical Calculation of Vortical Flows ................................................................................................ 3

Co-Investigator: Brian A. Robinson 

McDonnell Aircraft Company 

Jassim A. Al-Saadi Transonic Reynolds Number Effects on Aircraft Configurations............................................................... 4 

Co-Investigators: Richard A. Wahls, William K. Londenberg, and Melissa B. Rivers 

NASA Langley Research CenterNiGYAN, Inc. 

K. Appa Hydrodynamic Ram Structural Response Analyses.................................................................................. 5 

Co-Investigator: Wendy S. Pi 

Northrop Corporation 

Jeffrey R. Barnes Nonlinear Baroclinic Instability .............................................................................................................. 6 

Co-Investigator: Richard E. Young 

Oregon State University/NASA Ames Research Center 

John T. Batina Euler Flutter Analysis of a Complex Aircraft Configuration ...................................................................... 7 

Co-Investigator: Russ D. Rausch 

NASA Langley Research Center/Purdue University 

Oktay Baysal Aerodynamic Shape Optimization.......................................................................................................... 8 

Co-Investigators: Mohamed E. Eleshaky and Greg W. Burgreen 

Old Dominion University 

Diane M. Bell Turbulent Boundary Layer with Heat Transfer......................................................................................... 9 

Stanford University 

Robert F. Bergholz Three-Dimensional Exhaust-Nozzle Flow Fields for High-Speed Civil Transport................................... 10 

Co-Investigator: William M. Turner 

General Electric Aircraft Engines 

Marek K. Bleszynski Electromagnetic Scattering from Large Three-Dimensional Targets ....................................................... 11 

Rockwell International Science Center 

William W. Bower Short Takeoff and Vertical Landing Aircraft Thermal/Acoustic Loads..................................................... 12 

Co-Investigator: Robert E. Childs 

McDonnell Douglas Research Laboratories/Nielsen Engineering and Research, Inc. 

James G. Brasseur Local Passive Scalar Dispersion in a Turbulent Boundary Layer .......... ................................................... 13 

Co-Investigators: P. K. Yeung, Samir Khanna, and Brian P. Moquin 

Pennsylvania State University/Georgia Institute of Technology 

Stephen H. Brecht Global-Venus/Solar-Wind Interaction ................................................................................................... 14 

Berkeley Research Associates

PAGE BLANK NOT flLMED fr_ BLANK 
Principal 

Investigator NAS Summary Page 

Richard C. Buggeln Hot-Gas-Manifold Flow Simulation ...................................................................................................... 15

Co-Investigator: Sang-Keun Choi 

Scientific Research Associates, Inc. 

Alan B. Cain Simulation and Modeling of Turbulence and Flow Acoustics................................................................ 16 

Co-Investigators: Dennis F. Fuglsang and Nagi N. Mansour 

McDonnell Aircraft Company/NASA Ames Research Center 

Richard L. Campbell Automated Transonic Wing Design ...................................................................................................... 1 7 

Co-Investigator: Leigh Ann Smith 

NASA Langley Research Center 

Brian J. Cantwell Simulation of a Time-Developing Incompressible Plane Wake ............................................................. 18 

Co-Investigators: R. Sondergaard, J. Soria, and J. H. Chen 

Stanford University/Sandia National Laboratory 

Frank Caradonna Prediction of Advanced Rotor Performance .......................................................................................... 19 

- Co-Investigators: John Bridgeman and K. Ramachandran 

U.S. Army Aeroflightdynamics Directorate, AVSCOfvVNASA Ames Research Center 

R. R. Chamberlain Aerothermal Analysis of Hypersonic Defense Interceptors .................................................................... 20 

Co-Investigators: Lawrence W. Spradley and Kenneth E. Xiques 

Adaptive Research Corporation 

Dean R. Chapman Boost-Phase Detection Study................................................................................................................ 21 

Co-Investigators: Forrest F. Lumpkin Ill, Robert W. MacCormack, and Stephane Moreau 

Stanford University/NASA Ames Research Center 

Denny Chaussee Transonic and Supersonic Flow Past Aircraft Configurations ................................................................. 22 

Co-Investigators: Jolen Flores and Eugene Tu 

NASA Ames Research Center 

C. L. Chen Numerical Analysis of Three-Dimensional Separated Juncture Flows.................................................... 23 

Co-Investigator: C. M. Hung 

Rockwell International Science Center/NASA Ames Research Center 

H. C. Chen Euler Analysis of Turboprop and Turbofan Integration .......................................................................... 24 

Co-Investigators: T. Y. Su, T. J. Kao, and D. A. Naik 

The Boeing CompanyNiGYAN, Inc ./NASA Langley Research Center 

Lee T. Chen Advanced Transonic Wing Concepts .................................................................................................... 2.5 

Co-Investigators: K. C. Chang, R. Pelkman, and A. Shmilovich 

Douglas Aircraft Company 

Robert E. Childs Turbulence-Model Development for Impinging Jet Flows ..................................................................... 26 

Co-Investigator: Laura C. Rodman - Nielsen Engineering and Research, Inc. 

Rodrick V. Chima Prediction of Turbine Endwall Heat Transfer......................................................................................... 27 

NASA Lewis Research Center 

Chuen-Yen Chow Compressible Taylor—Couette Flow....................................................................................................... 28 

Co-Investigators: Meng-Sing Liou and Kai-Hsiung Kao 

University of Colorado, Boulder/NASA Lewis Research Center

vi 
Principal 

Investigator NAS Summary Page 

Wei J. Chyu Airframe and Inlet Aerodynamics.......................................................................................................... 29 

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

NASA Ames Research Center/Cornell University/Carnegie Mellon University 

Charles E. Cockrell, Jr. Generic National Aero-Space Plane Forebody—Inlet Integration............................................................ 30 

Co-Investigator: Lawrence D. Huebner 

NASA Langley Research Center 

William B. Compton, Ill Three-Dimensional Afterbody Flow with Jet Exhaust............................................................................. 31 

Co-Investigator: Khaled S. Abdol-Hamid 

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

Raymond R. Cosner Analyses of F/A-18E/F Upgrades ........................................................................................................... 32 

Co-Investigators: F. Creasman, R. S. Dyer, T. D. Gatzke, J. A. Johnson, P. J. Malloy, 

W. W. Romer, and P. G. Willhite 

McDonnell Aircraft Company 

Andrew J. Crook Endwall and Casing Treatment Flow in a Transonic Fan Rotor.............................................................. 33 

General Motors Corporation, Allison Gas Turbine Division 

Richard D. Crouse Fighter Acoustic-Load Predictions......................................................................................................... 34 

Northrop Corporation 

M. I. Cruz Flexible Aerobrake Aerothermodynamic Study ..................................................................................... 35 

Co-Investigators: D. E. Ressler, T. P. Shivananda, and E. F. Zabrensky 

TRW, Federal Systems Division/TRW, Ballistic Missiles Division 

Jeffrey N. Cuzzi Particle-Gas Dynamics in the Protoplanetary Nebula: ........................................................................... 36 

Co-Investigators: Joelle Champney and Anthony Dobrovolskis 

NASA Ames Research Center/A.T.M., Inc./University of California, Santa Cruz 

Russell B. Dahlburg Dynamical Modeling of the Solar Atmosphere...................................................................................... 37 

Co-Investigators: S. K. Antiochos, Jill P. Dahlburg, and J. T. Karpen 

Naval Research Laboratory 

Sanford M. Dash Simulation of Turbulent Jets for Aeroacoustic Applications ................................................................... 38 

Co-Investigators: Neeraj Sinha, Brian I. York, and Robert A. Lee 

Science Applications International Corporation 

Sanford M. Dash Three-Dimensional High-Speed Plume/Propulsive Flow-Field Analysis ................................................ 39 

Co-Investigators: Neeraj Sinha, Brian J. York, Robert A. Lee, Ashvin Hosagadi, and 

Donald C. Kenzakowski 

Science Applications International Corporation 

Donald W. Davis Hydrodynamic Performance Evaluation................................................................................................ 40 

Co-Investigator: Keith C. Kaufman 

General Dynamics, Electric Boat Division 

Roger L. Davis Turbine "Hot Spot" Alleviation using Film Cooling............................................................................... 41 

Co-Investigator: Daniel J. Dorney 

United Technologies Research Center 

Russell G. DeAnna Transition over a Rough Surface ........................................................................................................... 42 

Co-Investigator: Eli Reshotko 

U.S. Army Propulsion Directorate, AVSCOM/Case Western Reserve University

vii 
Principal 

Investigator NAS Summary Page 

Jerry E. Deese Flow about Almost Complete Aircraft and Hypersonic Configurations.................................................. 43

Co-Investigators: Ramesh K. Agarwal, Thomas P. Gielda, Jerry C. Johnson, and Mark Axe 

McDonnell Douglas Research Laboratories/NASA Ames Research Center 

Robert G. Deissler Sensitivity of Turbulence to Initial Conditions....................................................................................... 44

Co-Investigator: Frank B. Molls 

NASA Lewis Research Center 

G. S. Deiwert Self-Adaptive Grid Code Applied to Complex Three-Dimensional Flows.............................................. 45

Co-Investigators: Carol B. Davies and Ethiraj Venkatapathy 

NASA Ames Research Center/Sterling Software/Eloret Institute 

A. 0. Demuren Complex Three-DimensiOnal Turbulent Flows ...................................................................................... 46 

Old Dominion University. 

Virendra K. Dogra Rarefied Hypersonic Condition Wake Structures................................................................................... 47 

Co-Investigators: Richard G. Wilmoth, James N. Moss, and Joseph M. Price 

ViGYAN, Inc./NASA Langley Research Center 

J. Philip Drummond Fuel—Air Mixing Enhancement by Jet—Shock Interactions ...................................................................... 48 

Co-Investigators: Peyman Givi, Cyrus K. Madnia, and Craig J. Steinberger 

NASA Langley Research Center 

C. T. Edquist Aerobra king Studies of Three-Dimensional Nonequilibrium Viscous Flow............................................ 49 

Co-Investigator: T. J. Galambos 

Martin Marietta Astronautics Group 

Thomas A. Edwards Aerodynamic Optimization of Supersonic Transport Aircraft ................................................................ 50 

Co-Investigator: Samson H. Cheung 

NASA Ames Research Center/MCAT Institute 

T. Alan Egolf Rotary-Wing Airload Performance Prediction ........................................................................................ 51 

Co-Investigator: Brian E. Wake 

United Technologies Research Center 

Dean R. Eklund Transpiration Cooling for Scramjet Combustor Flow Fields ................................................................... 52 

National Research Council/NASA Langley Research Center 

Larry L. Erickson Flow Solver for Euler Equations on Unstructured Tetrahedral Meshes ................................................... 53 

Co-Investigator: M. Jahed Djomehri 

NASA Ames Research Center/Eloret Institute 

William J. Feiereisen Discrete Particle Simulation of Compressible Flow ............................................................................... 54 

Co-Investigators: Jeffrey McDonald, Brian Haas, lain Boyd, Donald Baganoff, 

Michael Fallavollita, Terry Denery, and Avijit Goswami 

NASA Ames Research Center/Stanford University 

William J. Feiereisen Aeroassist Flight Experiment Flow Simulation ....................................................................................... 55

Co-Investigators: G. E. Palmer, E. Venkatapathy, D. S. Babikian, D. K. Prabhu, and E. E. Whiting 

NASA Ames Research Center 

Michael C. Fischer F16xL Supersonic Laminar Low-Control Experiment............................................................................. 56

Co-Investigator: Chandra S. Vemu ru 

NASA Langley Research Center

viii 
Principal 

Investigator NAS Summary Page 

Neal T. Frink Analysis of Slender-Wing Geometries using Unstructured Grids ...........................................................57 

Co-Investigators: Paresh Parikh and Shahyar Pirzadeh 

NASA Langley Research Center 

Dana Gaitonde Turbulent Flow Past a Complete Hypersonic Reentry Configuration ..................................................... 58

Co-Investigator: Joseph Shang 

WL/FIMM, Wright Patterson AFB 

Joseph L. Garrett Integrated Hypersonic-Propulsion Flow Paths....................................................................................... 59

Co-Investigators: Kevin Van Dyke and Balu Sukar 

Pratt & Whitney/General Electric Aircraft Engines 

Thomas B. Gatski Compressible Turbulent Flows.............................................................................................................. 60 

Co-Investigator: Joseph H. Morrison 

NASA Langley Research Center 

Peter A. Gnoffo Hypersonic Flows in Chemical and Thermal None quilibrium............................................................... 61 

NASA Langley Research Center 

Raymond Gordnier Unsteady Delta-Wing Flow................................................................................................................... 62 

WL/FIMM, Wright Patterson AFB 

Isaac Greber Numerical Calculation of a Three-Dimensional Separated Flow ........................................................... 63 

Case Western Reserve University 

Fernando F. Grinstein Spatially Evolving Reactive Jets............................................................................................................. 64 

Naval Research Laboratory 

W. L. Grose Three-Dimensional Atmospheric Simulation Model.............................................................................. 65 

Co-Investigators: W. T. Blackshear, R. S. Eckman, and R. E. Turner 

NASA Langley Research Center 

Karen L. Gundy-Burlet Rotor—Stator Interaction in Turbomachines ........................................................................................... 66 

Co-Investigator: Akil Rangwalla 

NASA Ames Research Center 

Guru P. Guruswamy Fluid and Structure Integration for Aerospace Applications................................................................... 67 

Co-Investigator: Shigeru Obayashi 

NASA Ames Research Center 

Karl E. Gustafson Airfoil Lift and Thrust Generation in Hover Mode................................................................................. 68 

Co-Investigator: Robert R. Leben 

University of Colorado, Boulder 

Edward J. Hall Unsteady Counterrotation of Ducted Prop fans...................................................................................... 69 

General Motors Corporation, Allison Gas Turbine Division 

David Halpern Satellite Data Assimilation and Ocean General Circulation Models...................................................... 70 

Co-Investigators: Y. Chao and C. Roberto Mechoso 

Jet Propulsion Laboratory/University of California, Los Angeles 

David Halpern Atmospheric General Circulation Model Sensitivity to Sea Surface Temperature Fields......................... 71 

Co-Investigators: C. Roberto Mechoso and Robert Haskins 

Jet Propulsion Laboratory/University of California, Los Angeles

ix 
Principal 

Investigator NAS Summary Page 

David W. Halt Transonic Analysis on Unstructured Grids ............................................................................................ 72

McDonnell Aircraft Company/McDonnell Douglas Research Laboratories 

H. Harris Hamilton Radiative Structure in Aerobrake Shock Layers ..................................................................................... 73

Co-Investigator: Robert B. Greendyke 

NASA Langley Research Center/ViGYAN, Inc. 

Jeff L. Hansen Incorporation of a Three-Dimensional Multistage Viscous Code into a Compressor Design System ......74 

Co-Investigator: John Adamczyk 

General Motors Corporation, Allison Gas Turbine Division/NASA Lewis Research Center 

Julius F. Harris Swept-Wing Leading-Edge Transition....................................................................................................75 

Co-Investigator: Venkit Iyer 

NASA Langley Research Center/VIGYAN Inc. 

Lin C. Hartung Radiation Transport around Axisymmetric Blunt-Body Vehicles ........ ..................................................... 76 

NASA Langley Research Center 

H. A. Hassan New Approach for Transitional-Flow Modeling .................................................................................... 77

Co-Investigators: T. Wayne Young and Eric S. Warren 

North Carolina State University 

James D. Holdeman Hot-Gas Ingestion by a Short Takeoff and Vertical Landing Aircraft in Ground Proximity ..................... 78

Co-Investigators: David M. Fricker and S. Pratap Vanka 

NASA Lewis Research Center/University of Illinois, Urbana/Champaign 

Yeu-Chuan Hsia Full Navier—Stokes Analysis of a Three-Dimensional Scramjet Inlet ...................................................... 79

Co-Investigators: Endwell Daso and Sukumar Chakravarthy 

Rockwell International, Rocketdyne Division/Rockwell International Science Center 

Carl T. Hsieh High Angle-of-Attack Missile Aerodynamics......................................................................................... 80

Naval Surface Warfare Center 

Thomas T. Huang Submarine-Appendage Design and Turbulence Modeling .................................................................... 81

Co-Investigators: Michael Griffin, Jeff Tsai, William Smith, and Dane Hendrix 

Naval Surface Warfare Center/Jason Associates/Scientific Research Associates, Inc./MCAT Institute 

Lawrence D. Huebner Generic National Aero-Space Plane Fuselage Configuration Study ....................................................... 82

Co-Investigator: Kenneth E. Tatum 

NASA Langley Research Center/Lockheed Engineering and Sciences Company 

Joseph W. Humphrey Hypersonic Scramjet and Detonation Flows ......................................................................................... 83

Co-Investigators: Darrell W. Pepper, Thomas H. Sobota, Frank P. Brueckner, Barry Dyne, 

and Juan C. Heinrich 

Advanced Projects Research, Inc./University of Arizona/NASA Langley Research Center 

Fazle Hussain Coherent-Structure Interactions with Turbulence .................................................................................. 84

Co-Investigator: Mogens V. Melander 

University of Houston/Southern Methodist University 

Fazle Hussain Chemically Reacting Free-Shear Flows ................................................................................................. 85

Co-Investigators: Ralph Metcalfe and Fernando Grinstein 

University of Houston/Naval Research Laboratory
Principal 

Investigator NAS Summary Page 

Scott T. Imlay Aero-Assisted Orbital Transfer Vehicle Flow Fields............................................................................... 86 

Co-Investigators: Donald W. Roberts and Moeljo Soetrisno 

Amtec Engineering, Inc. 

Gregory A. Intemann Wing—Nacelle—Pylon Installations......................................................................................................... 87 

Co-Investigator: Todd R. Michal 

Douglas Aircraft Company/McDonnell Aircraft Company 

Kenneth M. Jones Analysis of High-Speed Civil Transport Configurations ......................................................................... 88 

Co-Investigators: Victor R. Lessard and M. A. Takallu 

NASA Langley Research Center/ViGYAN, Inc./Lockheed Engineering and Sciences Company 

K. Kailasanath Structure and Dynamics of Multidimensional Flames............................................................................ 89 

Co-Investigator: G. Patnaik 

Naval Research Laboratory 

K. Kailasanath Unsteady Nozzle Flow Fields ............................................................................................................... 90 

Co-Investigator: R. Ramamurti 

Naval Research Laboratory 

Osama A. Kandil Simulation and Control of Slender Wing Rock...................................................................................... 91 

Co-Investigator: Ahmed A. Salman 

NASA Lewis Research Center/Old Dominion University 

Carolyn R. Kaplan Unsteady Ethylene Jet-Diffusion Flames................................................................................................ 92 

Co-Investigators: Elaine S. Oran and Seung W. Baek 

Naval Research Laboratory/Korea Advanced Institute of Science and Technology 

Steve L. Karman, Jr. Unstructured Grid-Generation/Flow-Solver Calibration ........................................................................ 93 

Co-Investigator: Gregory S. Spragle 

General Dynamics, Fort Worth Division 

George E. Karniadakis Turbulent Flow over Riblet-Mounted Surfaces ...................................................................................... 94 

Princeton University 

G. E. Karniadakis Turbulent Flow over a Backward-Facing Step ....................................................................................... 95 

Co-Investigator: S. A. Orszag 

Princeton University 

Jack H. Kennedy Aeronautical Vehicle Radar Cross Sections ........................................................................................... 96 

Co-Investigator: Daniel A. White 

General Dynamics, Convair Division 

Christian L. Keppenne Orographically Forced Oscillations of the Martian Atmosphere............................................................ 97 

Co-Investigator: Jean 0. Dickey 

Jet Propulsion Laboratory 

Suk C. Kim Low-Thrust Chemical Rockets............................................................................................................... 98 

Sverdrup Technology, Inc ./NASA Lewis Research Center 

S. W. Kim Fluid Flow of Jets in Cross Flow............................................................................................................ 99 

Co-Investigator: T. J. Benson 

NASA Lewis Research Center

XI 
Principal 

Investigator NAS Summary Page 

Kevin R. Kirtley Multistage Turbomachinery Flows ...................................................................................................... 100

Co-Investigators: John J. Adamczyk, Tim A. Beach, Mark L. Celestina, William A. Maul, 

and Rick A. Mulac 

- Sverdrup Technology, Inc./NASA Lewis Research Center 

W. Kollmann Turbulent Boundary Layers with Suction.............................................................................................101 

Co-Investigator: P. Mariani 

University of California, Davis 

Linda D. Kral Receptivity, Transition, and Turbulence Phenomena ..........................................................................102 

Co-Investigators: William W. Bowerand Alexander Pal 

McDonnell Douglas Research Laboratories 

Linda D. Kral Turbulence Modeling for Three-Dimensional Flow Fields................................................................... 103

Co-Investigator: John A. Ladd 

McDonnell Douglas Research Laboratories/McDonnell Aircraft Company 

Ajay Kumar Translating-Strut Scramjet Inlet ...........................................................................................................104 

Co-Investigator: Dal J. Singh 

NASA Langley Research Center 

Dochan Kwak High-Lift Aerodynamic Flow...............................................................................................................105 

Co-Investigators: Stuart Rogers and Lyn Wiltberger 

NASA Ames Research Center 

Dochan Kwak Space Shuttle Main Engine Flow .............................................................. . .......................................... 106 

Co-Investigators: S. Yoon, C. Kiris, L. Chang, and R. Williams 

NASA Ames Research Center 

H. T. Lai Hydrogen—Air Reacting Flow Fields in Drag Reduction External Combustion .....................................107 

Sverdrup Technology, Inc. 

Budugur Lakshminarayana Steady and Unsteady Turbomachinery Flow Fields .............................................................................108 

Co-Investigators: S. Fan and R. Kunz 

Pennsylvania State University/General Motors Technical Center 

Budugur Lakshminarayana Steady and Unsteady Viscous Flow ....................................................................................................109 

Co-Investigators: Yin-Hsiang Ho and Anton Basson 

Pennsylvania State University 

Scott L. Lawrence Integrated Hypersonic Vehicle Flow-Field Analysis ............................................................................ 110

Co-Investigators: Bradford C. Bennett and Gregory A. Molvik 

NASA Ames Research Center/MCAT Institute 

C. C. Lee Numerical Investigation of Vehicles Similar to the National Aero-Space Plane ................................... 111

Co-Investigators: William Bower, Shawn Hagmeier, Brad Hopping, Morzata Mani, 

Scott Van Horn, Charles Vaporean, Patrick Vogel, and Darrell Weber 

McDonnell Douglas Corporation 

Elizabeth M. Lee Subsonic/Transonic Flutter Boundaries ...............................................................................................112 

Co-Investigator: Mike Gibbons 

NASA Langley Research Center/Lockheed Engineering and Sciences Company 

xii
Principal 

Investigator NAS Summary Page 

Sang Soo Lee Resonant-Triad Interaction in an Adverse Pressure-Gradient Boundary Layer......................................113 

Co-Investigator: M. E. Goldstein 

Sverdrup Technology, Inc ./NASA Lewis Research Center

Yu-Tai Lee Unsteady Flows for Naval Applications .............................................................................................. 114 

Naval Surface Warfare Center 

Sanjiva K. Lele Skewed Compressible Mixing Layers .................................................................................................. 115 

Stanford University 

Sanjiva K. Lele Turbulence Effects on Stagnation-Point Flow ...................................................................................... 116 

Stanford University 

Clark H. Lewis Three-Dimensional Parabolized and Full Navier—Stokes Techniques .................................................. 117 

Co-Investigator: Bilal A. Bhütta 

VRA, Inc. 

C. S. Lin Electron-Beam Injections from the Space Shuttle ................................................................................ 118 

Co-Investigator: M. Muller 

Aurora Science, Inc./Southwest Research Institute 

S. J. Lin Turbopump Rotor/Stator Flows ........................................................................................................... 119 

Rockwell International, Rocketdyne Division 

James M. Luckring Advanced Computational Fluid Dynamics Applications for Complex Configurations.......................... 120 

Co-Investigators: Farhad Ghaffari and Brent L. Bates 

NASA Langley Research Center 

M. G. Macaraeg Ignition and Structure of a Diffusion Flame with Vortex Interaction .................................................... 121 

Co-Investigators: T. L. Jackson and M. V. Hussaini 

NASA Langley Research Center/Old Dominion University/ICASE 

Robert D. MacEiroy Computer Simulation of Water—Membrane Interfaces......................................................................... 122 

Co-Investigators: Andrew Pohorille and Michael A. Wilson 

NASA Ames Research Center 

M. R. Malik Supersonic Boundary-Layer Transition on a Cone at Incidence........................................................... 123 

Co-Investigator: P. Balakumar 

High Technology Corporation/NASA Langley Research Center 

M. R. Malik Compressible Boundary-Layer Transition............................................................................................ 124 

Co-Investigator: C. L. Chang 

High Technology Corporation/NASA Langley Research Center 

Nagi N. Mansour Turbulence in Compressible Fluids ..................................................................................................... 125 

Co-Investigator: Gary N. Coleman 

NASA Ames Research Center 

David L. Marcum Flow-Field Calculations using Unstructured Grids .............................................................................. 126 

Co-Investigators: Ramesh K. Agarwal and David W. Halt 

Mississippi State University/McDonnell Douglas Research Laboratories

XIII 
Principal 

Investigator NAS Summary Page 

Fred W. Martin, Jr. Space Shuttle Flow Field.....................................................................................................................127 

Co-Investigators: Pieter Buning, Steve Labbe, Ray Gomez, Jeff Slotnick, Steve Parks, 

and Max Kandula 

NASA Johnson Space Center/NASA Ames Research Center 

Dimitri J. Mavriplis Multigrid Solution of the Euler Equations ............................................................................................ 128 

ICASE/NASA Langley Research Center 

Charles R. McClinton High Mach-Number Mixing in a Shock Tunnel Environment.............................................................. 129 

Co-Investigator: Robert D. Bittner 

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

Charles R. McClinton National Aero-Space Plane Configuration Trade Studies ..................................................................... 130 

Co-Investigators: Arthur D. Dilley and Richard W. Hawkins 

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

Charles R. McClinton Scramjet Inlet Interaction .................................................................................................................... 1 31 

Co-Investigators: Thomas N. Jentink and Arthur D. Dilley 

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

Charles R. McClinton Scramjet Engine Design Optimization................................................................................................. 132 

Co-Investigators: Pradeep S. Kamath and Marlon Mao 

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

W. J. McCroskey Aerodynamics and Acoustics of Rotorcraft .......................................................................................... 133 

Co-Investigators: J. D. Baeder, V. Raghavan, G. R. Srinivasan, and S. K. Stanaway 

U.S. Army Aeroflightdynamics Directorate, AVSCOWNASA Ames Research Center 

W. J. McCroskey High-Performance Rotor-Blade Tips .................................................................................................... 1 34 

Co-Investigators: E. P. N. Duque, V. Raghavan, C. R. Srinivasan, and R. C. Strawn 

U.S. Army Aeroflightdynamics Directorate, AVSCOM/NASA Ames Research Center 

Suresh Menon Active Control of Ramjet Combustion Instability................................................................................. 135 

QUEST Integrated, Inc. 

Charles L. Merkle Coupled Navier—Stokes Maxwell Analysis for Microwave Propulsion ................................................. 136 

Co-Investigator: S. Venkateswaran 

Pennsylvania State University 

Parviz Moin Turbulent Flow Over Riblets............................................................................................................... 137 

• Co-Investigators: John Kim and Haecheon Choi 

Stanford University/NASA Ames Research Center 

Parviz Moin Turbulence Over a Backward-Facing Step .......................................................................................... 138 

Co-Investigators: John Kim and Hung Le 

NASA Ames Research Center 

Parviz Moin Aerodynamic Sound Generation......................................................................................................... 139 

Co-Investigators: Sanjiva K. Lele, Tim Colonius, and Brian E. Mitchell 

Stanford University/NASA Ames Research Center 

Parviz Moin Shock-Wave/Turbulent-Boundary-Layer Interaction........................................................................... 140 

Co-Investigators: Sanjiva K. Lele, Mahesh Krishnan, and Sangsan Lee 

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

xiv 
Principal 

Investigator NAS Summary Page 

Robert D. Moser Wall-Bounded Turbulent Flows ..........................................................................................................141 

Co-Investigator: John Kim 

NASA Ames Research Center 

James N. Moss Hypersonic Rarefied Flow about a Compression Ramp....................................................................... 142 

Co-Investigator: Joseph M. Price 

NASA Langley Research Center 

Jon S. Mounts Transonic-Shock/Boundary-Layer Interaction to Alleviate Separation.................................................. 143 

Co-Investigator: Thomas J. Barber 

United Technologies Research Center 

J. C. Narramore Rotorcraft Drag Prediction .................................................................................................................. 144 

Co-Investigators: A. C. Brand and D. W. Axley 

Bell Helicopter Textron, Inc. 

David Nixon Controlling Combustion in Free-Shear Layers ..................................................................................... 145 

Co-Investigator: Laurence Keef 

Nielsen Engineering and Research, Inc. 

Ahmed K. Noor Thermomechanical Buckling and Post-Buckling of Multilayered Composite Panels with Cutouts ....... 146 

Co-Investigator: Jeanne M. Peters 

NASA Langley Research Center/University of Virginia 

Grant Palmer Space Transportation Vehicle Aerothermodynamics ........................................................................... 147 

Co-Investigators: Susan Tokarcik, Ethiraj Venkatapathy, and William J. Feiereisen 

NASA Ames Research Center

Ajay K. Pandey Thermo-Viscoplastic Analysis of an Engine-Cowl Leading Edge........................................................... 148

Lockheed Engineering and Sciences Company/NASA Langley Research Center 

S. Paul Pao Afterbody Aerodynamics with Canted Pitch-Vectoring Twin Nozzles .................................................149 

Co-Investigator: K. S. Abdol-Hamid 

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

V. C. Patel Hydrodynamics of Self-Propelled Bodies ............................................................................................150 

Co-Investigator: F. Stern 

University of Iowa 

Darrell W. Pepper Hypersonic Scramjet Flow..................................................................................................................151 

- Co-Investigators: Frank P. Brueckner and Kevin L. Burton

Advanced Projects Research, Inc. 

Leonhard Pfister Wave-Induced Transports...................................................................................................................152 

NASA Ames Research Center 

Thanh T. Phan Rocket-Base Flow-Field Simulations ...................................................................................................153 

Co-Investigator: Richard J. Magnus 

General Dynamics, Space Systems Division 

Richard H. Pletcher Three-Dimensional Liquid-Sloshing Flows ..........................................................................................154 

Co-Investigators: Franklyn J. Kelecy and Babu Sethuraman 

Iowa State University

xv 
Principal 

Investigator NAS Summary Page 

James B. Pollack Martian Atmosphere General Circulation ........................................................................................... 155

Co-Investigator: Robert Haberle 

NASA Ames Research Center 

Greg D. Power Three-Dimensional Scramjet Combustors ........................................................................................... 156

Co-Investigator: Thomas J. Barber 

United Technologies Research Center 

Ramadas K. Prabhu Hypersonic-Flow Computations using Adaptive Unstructured Meshes................................................ 157

Lockheed Engineering and Sciences Company/NASA Langley Research Center 

C. D. Pruett Laminar Breakdown of High-Speed Boundary-Layer Flow.................................................................. 158

Co-Investigator: T. A. Zang 

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

R. Ramakrishnan Shock Impingement in Hypersonic Engine Inlets................................................................................. 159

NASA Langley Research Center 

Edward J. Reske Complex Three-Dimensional Flows in the Advanced Solid-Rocket Motor .......................................... 160

Co-Investigators: Dana F. Billings and Joni W. Cornelison 

NASA Marshall Space Flight Center 

David L. Rigby Leading-Edge Heat Transfer in a Flow with Spanwise Variations......................................................... 161

Sverdrup Technology, Inc. 

Charles L. Rino Scattering from Ocean Surfaces and Near-Surface Objects ................................................................. 162

Co-Investigators: Hoc D. Ngo, Thomas L. Crystal, and Carl E. Hess 

Vista Research, Inc. 

R. P. Roger Jet Interaction Aero-Optic Effects on Hypersonic Interceptors ............................................................. 163

Co-Investigator: S. C. Chan 

Teledyne Brown Engineering 

Michael M. Rogers Mixing and Reacting in Plane Mixing Layers....................................................................................... 164

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

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

R. Clayton Rogers Hypervelocity Mixing and Combustion in Pulse Facility Flows ........................................................... 165

Co-Investigators: David W. Riggins, Robert D. Bittner, and Glenn J. Bobskill 

NASA Langley Research Center/University of Missouri, Rolla/ 

Analytical Services and Materials, Inc. 

William C. Rose Hypersonic Inlet Flow Fields............................................................................................................... 166

NASA Ames Research Center 

William C. Rose National Aero-Space Plane Inlet Boundary-Layer Control................................................................... 167

NASA Lewis Research Center 

William C. Rose National Aero-Space Plane Inlet Flow Fields....................................................................................... 168

NASA Ames Research Center 

James C. Ross Trapped-Vortex Flows on Highly Swept Wing Configurations ............................................................. 169 

Co-Investigators: Paula Lovely and Todd Riddle 

NASA Ames Research Center 

xvi
Principal 

Investigator NAS Summary Page 

Karlin R. Roth Validation of a Short Takeoff and Vertical Landing Model ...................................................................170 

Co-Investigator: Stephen Chiu 

NASA Ames Research Center/California Polytechnic State University, San Luis Obispo 

Stephen M. Ruffin Coupled Rotation—Vibration—Dissociation Processes in Hypersonic Flows .......................................... 1 71 

Co-Investigators: Surendra P. Sharma and Ellis Whiting 

NASA Ames Research Center 

Christopher L. Rumsey High-Alpha Flow Fields ......................................................................................................................1 72 

Co-Investigators: James L. Thomas, Robert T. Biedron, Sherrie L. Krist, W. Kyle Anderson, 

Daryl L. Bonhaus, David L. Whitaker, Thomas W. Roberts, Gary P. Warren, Beyung Kim, 

and Robert P. Weston 

NASA Langley Research Center 

Jose M. Sanz Aerodynamic Inverse Design and Analysis for a Full Engine ...............................................................173 

NASA Lewis Research Center 

Lewis B. Schiff High Alpha Technology Program F- 18 Aerodynamics ........................................................................174 

Co-Investigators: Yehia Rizk, Ken Gee, and Scott Murman 

NASA Ames Research Center 

Lewis B. Schiff High Angle-of-Attack Vortex Flow Aerodynamics ............................................................................... 175 

Co-Investigators: Yehia Rizk, Neal Chaderjian, David Degani, John.Ekaterinaris, Yuval Levy, 

and Gabriel Font 

NASA Ames Research Center 

Balu Sekar Viscous Non-Reacting Flows in High-Speed Combustors .................................................................... 1 76 

Co-Investigators: Douglas Davis and Daniel Risha 

WL/FIMM, Wright Patterson AFB 

Vijaya Shankar Computational Fluid Dynamics Approach to Computational Electromagnetics ................................... 177

Co-Investigators: William F. Hall, Alireza Mohammadian, and Chris Rowell 

Rockwell International Science Center

Samuel P. Shanks Computational Fluid Dynamics of Store Separation ............................................................................178 

Co-Investigator: Jasim U. Ahmad 

JAI Associates, Inc. 

Tom I-P. Shih Wankel Engine Flow Fields.................................................................................................................179 

Co-Investigators: E. Steinthorsson, Z. Li, A. Karadag, E. A. Willis, and J. McFadden 

Carnegie Mellon University/NASA Lewis Research Center 

B. A. Singer Formation and Growth of Hairpin Vortices.........................................................................................180 

Co-Investigators: R. D. Joslin and S. P. G. Dinavahi 

NASA Langley Research Center 

Brian R. Smith Computational Fluid Dynamics Methods for Highly Maneuverable Aircraft........................................181 

Co-Investigators: C. L. Reed and A. Muyshondt 

General Dynamics, Fort Worth Division 

Bruce F. Smith Numerical Experiments in the Formation and Evolution of Galaxies ...................................................182 

Co-Investigator: Richard H. Miller 

NASA Ames Research Center/University of Chicago

xvii 
Principal 

Investigator NAS Summary Page 

Merritt H. Smith YAV-8B Aircraft Flow Simulation........................................................................................................183 

Co-Investigators: Kalpana Chawla and William R. Van Dalsem 

NASA Ames Research Center 

Robert E. Smith Grid Generation for Aerodynamic Configurations............................................................................... 184 

Co-Investigator: Michael J. Bockelie 

NASA Langley Research Center/GEOLAB 

K. Snyder National Aero-Space Plane High-Speed Combustion Experiment........................................................ 185 

Co-Investigators: J.-L. Cambier, D. K. Prabhu, and S. Tokarcik 

NASP Joint Program Office/NASA Ames Research Center 

Philippe R. Spalart Complex Turbulent Boundary Layers.................................................................................................. 186 

Boeing Commercial Airplane Group 

Fred Stern Free-Surface Effects on Boundary Layers and Wakes........................................................................... 187 

University of Iowa 

Eric T. Stewart Space Station Freedom Internal Flow Analysis .................................................................................... 188 

Co-Investigator: Lee A. Kania 

NASA Marshall Space Flight Center/Sverdrup Technology, inc. 

Olaf 0. Storaasli Parallel-Vector "Out-of-Core" Equation Solver.................................................................................... 189 

NASA Langley Research Center 

Michael W. Stortz Harrier YA V-8B Wing Aerodynamics.................................................................................................. 190 

Co-Investigators: Lie-Mine Gea and Gil W. Chyu 

NASA Ames Research Center 

Sundares V. Subramanian Viscous Reacting-Flow Applications to Scramjet Propulsion ............................................................... 191 

Co-Investigator: Richard J. Gaeta, Jr. 

General Electric Aircraft Engines 

Dennis Sullivan Microwave Hyperthermia Computer Modeling................................................................................... 192 

Stanford University School of Medicine

Chao-Ho Sung Computational Fluid Dynamics for Naval Applications ......................................................................193 

Co-Investigators: Michael J. Griffin, Chi-Wang Shu, and George T. Yeh 

David Taylor Research Center 

R. C. Swanson Simulation of Scramjet Flow Fields .....................................................................................................194 

Co-Investigator: E. Turkel 

NASA Langley Research Center 

Tsze C. Tai Low-Speed Aircraft Maneuvering Aerodynamics.................................................................................195 

Naval Surface Warfare Center 

Rajiv Thareja Unstructured Multigrid Euler Solver....................................................................................................196 

Co-Investigators: J. Peraire, J. Peiro, and K. Morgan 

Lockheed Engineering and Sciences Company/Imperial College/University College, Wales 

Gary E. Thomas Multi-Dimensional Simulations of Noctilucent Cloud Formation ........................................................197 

Co-Investigator: Eric J. Jensen - University of Colorado, Boulder/NASA Ames Research Center 

xviii 
Principal 

Investigator NAS Summary Page 

Owen B. Toon Stratospheric Ozone Destruction ........................................................................................................ 198 

NASA Ames Research Center 

Ban H. Tran Radar Cross-Section Studies ............................................................................................................ ... . 199 

Rockwell International, North American Aircraft Division 

Arnaud Trouvé Turbulent Premixed Combustion ........................................................................................................ 200 

Co-Investigator: Thierry Poinsot 

NASA Ames Research Center/Stanford University 

Eli Turkel Multigrid Solution of the Navier—Stokes Equations.............................................................................. 201 

ICASE/NASA Langley Research Center 

Walter 0. Valarezo Multi-Element High-Lift Concepts....................................................................................................... 202 

Co-Investigators: Vincent D. Chin and Dimitri J. Mavriplis 

Douglas Aircraft Company/NASA Langley Research Center/ICASE 

W. R. Van Dalsem Simulation of the Unsteady Flow about Transonic Cavities................................................................. 203 

Co-Investigators: C. A. Atwood, S. P. Klotz, and D. Sondak 

NASA Ames Research Center/MCAT Institute 

W. R. Van Dalsem Powered-Lift Computational Fluid Dynamics Project.......................................................................... 204 

Co-Investigator: K. Chawla 

NASA Ames Research Center/MCAT Institute 

Veer N. Vatsa High-Reynolds-Number Viscous Flow over Aircraft Components ....................................................... 205 

Co-Investigator: Eli Turkel 

NASA Langley Research Center 

Gururaja R. Vemaganti Two-Equation Turbulence Model Implementation .............................................................................. 206 

Lockheed Engineering and Sciences Company 

Bill J. Walker Tactical Missile Aero-Propulsion Interaction....................................................................................... 207 

Co-Investigators: C. D. Mikkelsen, K. D. Kennedy, P. F. Booth, and M. E. Vaughn, Jr. 

U.S. Army Missile Command 

Robert W. Walters Finite-Rate Chemistry Algorithms........................................................................................................ 208 

Co-Investigators: Bernard Grossman, Michael Applebaum, Andrew S. Godfrey, 

William D. McGrory, and David C. Slack 

Virginia Polytechnic Institute and State University 

Jong H. Wang Hydrocarbon Scramjet Combustor Flows............................................................................................ 209 

Rockwell International, North American Aircraft Division 

Kurt F. Weber Flow Simulation in Fan and Core Compressor Stages ..... ..................................................................... 210 

Co-Investigator: Dale W. Thoe 

General Motors Corporation, Allison Gas Turbine Division 

Kenneth J. Weilmuenster Winged Entry-Vehicle Computations .................................................................................................. 211 

Co-Investigators: Francis A. Greene and William L. KIeb 

NASA Langley Research Center 

Kenneth J. Weilmuenster Shuttle Leeside Temperature Study ..................................................................................................... 212 

Co-Investigator: William L. Kleb 

NASA Langley Research Center

xix 
Principal 

Investigator NAS Summary Page 

Douglas L. Westphal 

J. A. White 

Richard G. Wilmoth 

Chung-Jin Woan 

Chung-Jin Woan 

J. P. Wolfe 

Alex C. Woo 

Perry A. Wooden 

Richard F. Young 

Tropospheric Aerosols and Clouds .....................................................................................................213 

Co-Investigator: Owen B. Toon 

NASA Ames Research Center 

Supersonic Chemically Reacting Turbulent Flow................................................................................214 

Co-Investigators: J. P. Drummond, J. J. Korte, and M. H. Carpenter 

NASA Langley Research Center 

Shock Interactions in Hypersonic Rarefied Flows................................................................................215 

Co-Investigators: V. K. Dogra and J. N. Moss 

NASA Langley Research CenterNiGYAN, Inc. 

Supersonic Laminar Flow Control Concept.........................................................................................216 

Co-Investigators: Michael W. George and Philip B. Gingrich 

Rockwell International, North American Aircraft Division 

Nozzle Screech Phenomena ...............................................................................................................217 

Co-Investigator: Michael W. George 

Rockwell international, North American Aircraft Division 

Advanced Computational Materials Research .....................................................................................218 

Co-Investigators: R. Averback, J. Kogut, and P. Wolynes 

University of Illinois, Urbana/Champaign 

Computational Electromagnetics.........................................................................................................219 

Co-Investigators: Michael J. Schuh and Michael P. Simon 

NASA Ames Research Center 

Numerical Simulation of the PAMPA Jet Trainer.................................................................................220 

Co-Investigator: Ronald D. Lowe 

LTV, Aircraft Division 

Numerical Simulation of Submarine Propulsion..................................................................................221 

David Taylor Research Center 

Steady and Unsteady Flows at High Angle of Attack...........................................................................222 

Co-Investigator: Hiroshi ide 

Rockwell International, North American Aircraft Division 

Volcanic Aerosol Clouds in the Stratosphere.......................................................................................223 

Co-Investigators: Howard Houben and Owen B. Toon 

NASA Ames Research Center/Space Physics Research Institute 

Single Expansion Ramp Nozzle with External Burning........................................................................224 

ICOMP/NASA Lewis Research Center 

TransitionModeling............................................................................................................................225 

Co-Investigators: S. Dinavahi, U. Piomelli, R. Garg, and K. Ball 

NASA Langley Research Center/University of Maryland/University of Texas 

Cheng-I Yang 

David T. Yeh 

Shaye Yungster 

T.A. Zang

Indexby Research Sites ..................................................................................................................................................................... 227

xx
Numerical Solution of Three-Dimensional Maxwell's Equations 

Ramesh K. Agarwal, Principal Investigator 

Co-Investigators: Mark Shu and Mark R. Axe 

McDonnell Douglas Research Laboratories 

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 vehicle configurations could be considered and their 

performance evaluated, thus reducing the cost and time required 

in the design cycle. 

Approach 

The approach taken is to solve the three-dimensional Maxwell 

equations employing a mature computational fluid dynamics 

technology. Specifically, a compact higher-order finite-volume 

time-domain/frequency-domain method is used to compute the 

EM field about scattering bodies. 

Accomplishment Description 

The two-dimensional EM-scattering code MDRCS2D was 

extended to calculate the scattering from dielectric objects and 

was extended to three dimensions. Especially noteworthy are the 

trailing edge and transverse magnetic scattering calculations 

from an airfoil, an ogive, a slit, and the EM-scattering from a 

perfectly conducting sphere. 

Significance 

The proposed project represents the development of leadingedge technology in computational EM and will have substantial 

impact on the development of both fighter aircraft and missile 

programs. 

Future Plans 

To develop and validate a three-dimensional Maxwell solver and 

use it to calculate the radar cross-section of complete, realworld, three-dimensional configurations.

Publications 

1. Huh, K. S.; Agarwal, R. K.; and Widnell, S. E. "Numerical 

Simulation of Acoustic Diffraction of Two-Dimensional Rigid 

Bodies in Arbitrary Flows." AIAA Paper 90-3920, 1990. 

2. Huh, K. S.; Shu, M.; and Agarwal, R. K. "A Compact HighOrder Finite-Volume Time-Domain/Frequency-Domain 

Method for Electromagnetic Scattering." AIAA Paper 92-0453, 

1992. 

Amplitude of magnetic induction (B2) of a scattered field for the 

trailing-edge case about the airfoil NACA0012. Red represents 

maximum amplitude and blue represents minimum amplitude.
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Helicopter Fuselage, Rotor, and Rotor/Body Interaction Flow-Field Calculations 

Ramesh K. Agarwal, Principal Investigator 

Co-Investigator: Jerry E. Deese 

McDonnell Douglas Research Laboratories 

Research Objective 

To develop computational codes capable of predicting the 

transonic viscous flow about helicopter fuselages, rotors, and 

rotor-fuselage combinations in hover and forward (light. 

Approach 

For calculating turbulent flow about helicopter fuselages, the 

Reynolds-averaged Navier—Stokes (RNS) equations are solved in 

an inertial frame on a body-conforming curvilinear grid using a 

finite-volume Runge—Kutta time-stepping scheme. For calculating turbulent flow about rotor blades, the RNS equations are 

solved in a rotating coordinate system. Rotor-wake effects are 

modeled in the form of a correction applied to the geometric 

angle of attack along the blades. Two approaches are being 

examined for rotor/body interaction calculations. The zonal 

approach uses a grid fixed to the blade and a grid fixed to the 

fuselage with a buffer grid between that shears in time. The 

chimera approach uses grids fixed to and overlapping the blades 

and the fuselage. 

Accomplishment Description 

Flow fields about the MDX fuselage in both ascending and 

descending flight were calculated for validation of the no-tail

rotor concept. Zonal and chimera grids have been generated for 

a generic rotor/body configuration. Cross-sectional planes of this 

grid are shown in the accompanying figures. 

Significance 

Development of a code that can compute rotor/body interaction 

flow field will allow the designer to analyze the interference 

effects of the two components and optimize the configuration. 

Future Plans 

The primary focus is on the calculation of rotor/body interaction 

flow field. Calculations will be performed on multibladed 

helicopter rotor/fuselage configurations and will be compared 

with experimental data for code validation. Calculations will 

then be performed on new unconventional rotorcraft 

configurations. 

Publication 

Adams, F. P. "Decomposition Strategies for a Multi-Block TimeDependent Grid about a Helicopter Rotor/Fuselage Configuration." Master's Thesis, Mississippi State University, Dec. 1991. 

Cross-sectional views of a zonal grid for a generic two-bladed helicopter configuration. 

2
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cc

Numerical Calculation of Vortical Flows 

Shreekant Agrawal, Principal Investigator 

Co-Investigator: Brian A. Robinson 

McDonnell Aircraft Company 

Research Objective 

To use computational fluid dynamics (CFD) methods to accurately predict vortical flows and to develop a vortex breakdown 

criterion that allows rapid determination of breakdown location 

from a CFD solution. 

Approach 

The CFL3D Euler/Navier—Stokes code based on an implicit, finite 

volume, upwind scheme is used to calculate vortical flows over 

a flat plate 70 degree delta wing at high angles of attack. Very 

dense grids using grid embedding are used to resolve the vortical 

flow features. An integral formulation of the Rossby number is 

employed to develop a vortex breakdown criterion for the flow 

about this type of wing. The CFL3D code is also used to 

compute initial solutions on three different tangent ogive 

forebodies at high angles of attack. The TLNS3D code is also 

used to predict the highly separated flow over a wing upper 

surface which results in a separation vortex. 

Accomplishment Description 

The flat plate delta wing has been analyzed at M,, = 0.3 and 

Re = 1 million for several high angles of attack to examine 

leading edge vortex breakdown. An H-O grid topology 

(61 axial x 65 normal x 89 circumferential) with and without 

grid embedding has been used to examine effects of grid 

resolution. A vortex breakdown criterion based on the Rossby 

number was developed. A Rossby number of about 1.0 appears 

to indicate a burst vortex (shown in the figure). Mach number 

effects and solution instability were also investigated using the 

Rossby number. Each solution required about 10 Cray-2 hours 

and 40 megawords of central memory. Initial solutions on three 

tangent ogive forebodies (circular, elliptic, and chined cross 

sections) at high angle of attack have also been computed on 

sequenced grids. These initial solutions required about 2 Cray-2 

hours and 16 megawords of memory. Large scale separation and 

a separation vortex have been computed on a wing—fuselage 

geometry. This case required 10 Cray-2 hours and 95 megawords 

of memory.

Future Plans 

The Rossby-number vortex breakdown criterion will be applied 

to more complex geometries at different Mach numbers. Fine 

grid solutions will be obtained on the tangent ogive forebodies 

and solutions will be compared to wind tunnel data. Surface 

blowing will be investigated as a vortex control mechanism on a 

chined forebody configuration. 

Publications 

1. Agrawal, S.; Barnett, R. M.; and Robinson, B. A. "Numerica 

Investigation of Vortex Breakdown on a Delta Wing." 

AIAA Journal 30 (Mar. 1992): 584-591. 

2. Agrawal, S.; Robinson, B. A.; and Barnett, R. M. "Prediction 

of Vortex Breakdown on a Delta Wing." Presented at the 

Fifth Symposium on Numerical and Physical Aspects of 

Aerodynamic Flows, Long Beach, CA, Jan. 1992. 

3. Robinson, B. A.; Barnett, R. M.; and Agrawal, S. "A Simple 

Numerical Criterion for Vortex Breakdown." AIAA 

Paper 92-0057, Jan. 1992. 

4 r Angle of attack, degrees 

o 20 

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Dccurs 

_1 t I I I I 

0 .2 .4 .6 .8 1.0 

Fraction root chord 

Significance 

The Rossby-number vortex breakdown criterion identifies burst 

location as the CFD code is running, easing the burden on the 

aircraft designer to perform time consuming postprocessing to 

obtain the same information. As a result of the identification of 

the separation vortex from the highly separated wing, fences are 

being investigated as a potential fix.

Resolving vortical flows and locating burst using the Rossby 

number; 70 degree delta wing, M,, = 0.3, Re = 1 x 106, CFL3D 

laminar solutions. The Rossby number locates the vortex burst.
Transonic Reynolds Number Effects on Aircraft Configurations 

Jassim A. Al-Saadi, Principal Investigator 

Co-Investigators: Richard A. Wahls, William K. Londenberg, and Melissa B. Rivers 

NASA Langley Research CenterNiGYAN, Inc. 

Research Objective 

To predict Reynolds number effects on complex aircraft configurations at transonic and subsonic speeds through the application 

of computational fluid dynamics (CFD) techniques. 

Approach 

Various aircraft configurations are examined using established 

CFD codes. The aircraft under consideration correspond to 

models tested in the National Transonic Facility (NTF) under 

current research programs. The use of various CFD codes allows 

assessment of viscous modeling capabilities as well as the ability 

to treat complex aircraft geometries. 

Accomplishment Description 

Three-dimensional compressible Navier—Stokes solutions for the 

forebody of the X-29 research aircraft, including the nose boom 

and canopy, have been obtained for the flight Reynolds number 

at angles of attack of 30 and 40 degrees. The calculations are 

performed using the CFL3D code, with a typical solution 

requiring 18 megawords of memory and 6 Cray-2 hours. 

Comparison with experimental data is awaiting NTF testing. 

Three-dimensional wing/fuselage solutions for an advanced 

high-wing transport have been obtained using the unstructuredgrid Euler code USM3D. A typical solution requires 32 megawords 

of memory and 8.5 Cray-2 hours. Comparison is being made 

with existing NTF force and pressure data. The agreement with 

measured wing pressure coefficients is typical of inviscid 

calculations. 

Significance 

This research is part of the ongoing Reynolds-number scaling 

program. The use of Navier—Stokes codes is necessary for 

investigating the effects of the Reynolds number, while the 

treatment of complex geometries is simplified through use of 

unstructured-grid codes, which are currently Euler methods. 

The current approach allows code selection to be tailored to the 

given problem while providing a base of experience with such 

codes. The results obtained will help lead toward reliable 

prediction of Reynolds number effects and indicate the strengths 

and weaknesses of current CFD codes in such applications.

Future Plans 

The USM3D code will be used to analyze a transport configuration that includes underwing engine nacelles. Existing NavierStokes codes will be used to analyze High-Speed Civil Transport 

configurations at subsonic and transonic speeds. 

Upper surface pressure contours from an unstructured-grid Euler 

calculation of a high-wing transport wing/fuselage configuration 

at transonic cruise conditions.
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Hydrodynamic Ram Structural Response Analyses 

K. Appa, Principal Investigator 

Co-Investigator: Wendy S. Pi 

Northrop Corporation 

Research Objective "-•" t = 0 initial state 

To predict the hydrodynamic ram pressures and the structural t 6.10 msec (classic solution) 

responses of fuel tanks subjected to ballistic threats based on -°- t = 6.09 msec (current CFD method using 8 x 2 x 200 grid) 

computational fluid dynamics (CFD) methods. Dianhraam 

Approach 

The computational scheme employs finite-element methods for 

representing the fluid domain as well as the fuel-cell structures. 

Euler equations with numerical damping are used to describe the 

motion of the fluid due to impacts and traversal of the fuel cells 

by the projectiles. 

Accomplishment Description 

The finite-element based CFD code for hydrodynamic ram 

analysis was numerically exercised for two one-dimensional 

unsteady-gas dynamics problems: (1) the Riemann problem (a 

long shock tube separating two initial gas states) which simultaneously contains a shock wave, a contact discontinuity, and an 

expansion fan, and (2) the problem of two interacting blast 

waves which involve multiple interactions of strong shocks and 

rarefactions and with contact discontinuities. The computed flow 

quantities compare very well with the theoretical results for both 

cases. The average job run for the test cases needed only a 

minimal amount of memory and Cray hours. 

Significance 

The ability to predict the hydrodynamic ram pressures and the 

accompanying structural response of fuel tank panels is of great 

importance in designing aircraft for survivability against ballistic 

threats. There is no generally accepted and validated method by 

the United States aircraft survivability/vulnerability community 

to predict hydrodynamic ram pressures generated by highly 

explosive projectiles that detonate in the fuel, the liquid/structure 

interaction, and the structural response of the fuel tank. 

Future Plans 

The current code will be developed to include a dynamic model 

for a high velocity projectile traversing the fuel tank.

120,000 

100,000 

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80,000 

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60,000 

40,000 

0 2 4 6 8 10

X (m) 

Correlation of flow quantities in a shock tube (Riemann problem).

5
Nonlinear Baroclinic Instability 

Jeffrey R. Barnes, Principal Investigator 

Co-Investigator: Richard E. Young 

Oregon State University/NASA Ames Research Center 

Research Objective 

To investigate nonlinear baroclinic instability and transient 

baroclinic eddies in the atmospheres of Earth and Mars through 

numerical simulation. 

Approach 

A three-dimensional, spectral, primitive-equation model of 

atmospheric circulation is employed to simulate the nonlinear 

development and evolution of transient baroclinic eddies. This 

model offers computational efficiency through a semi-implicit 

time scheme, high spatial resolution, considerable flexibility in 

configuration, and on-line energy diagnostics. 

Accomplishment Description 

A study of the influence of thermal damping and surface drag on 

the nonlinear evolution of baroclinic eddies in Earth's atmosphere was completed with the performance of a final set of 

simulations and the computation of model diagnostics. The Earth 

simulations, typically requiring about 1-2 Cray Y-MP hours and 

10 megawords of memory, examined the fundamental roles of 

thermal damping and surface drag in the evolution of baroclinic 

eddies beyond an initial life cycle of growth and decay. These 

experiments demonstrated the crucial importance of surface 

drag for the eddy development through its effect on the mean 

circulation. The experiments also revealed novel aspects of 

wave/mean-flow interaction associated with a critical surface in 

the subtropical upper troposphere. A circulation model suitable 

for simulations of the Martian atmosphere was developed and 

tested in several initial experiments. It offers very high vertical 

resolution (the model top is 80 km altitude) and high spatial 

resolution and demands about 5-10 Cray Y-MP hours and 

20 megawords of memory per simulation. An important part of 

the Mars model development effort was devoted to the creation 

of a realistic global basic state. The initial numerical experiments 

were to determine the optimal values for the internal model 

diffusion and to examine the effect of the model's upper 

boundary. 

Significance 

Transient baroclinic eddies are important for the Earth's atmospheric circulation in middle latitudes. Cyclonic weather systems 

are manifestations of these eddies and transport large amounts of 

heat poleward. Because of this transport, baroclinic eddies are 

key players in the global climate system. Major NASA programs 

are directed at an enhanced understanding of the 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 Mars climate system. The 

transient baroclinic eddies in the terrestrial and Martian atmospheres are fundamentally similar and constitute two "laboratories" in which to investigate the basic dynamics of eddies. 

Future Plans 

Simulations of nonlinear baroclinic instability in the Martian 

atmosphere will be performed to investigate the role of strong 

thermal damping in eddy evolution and scale selection, effects 

of large topography, and the influence of basic-state structure on 

eddy dynamics. Simulations of Earth's atmosphere will also be 

conducted to examine the manner in which baroclinic eddies 

act to regulate the equator-to-pole temperature gradient through 

their heat and momentum transports. 

Publication 

Barnes, J. R. and Young, R. E. "Nonlinear Baroclinic Instability 

on the Sphere: Multiple Life Cycles with Surface Drag and 

Thermal Damping." J. Atmo. Sd. (May 15, 1992). 

70 - 

600 

z 

5: 000^ 

0 30 60 90 120 150 180 

Longitude 

III — 

05 .3 2 15 3 35

Isentropic potential vorticity at a level near the tropopause in a 

simulation of nonlinear baroclinic instability in Earth's atmosphere. Red/blue colors denote high/low values of this conserved 

dynamical quantity, evidencing "wrapping up" and mixing 

processes both pole ward and equatorward of the midlatitude 

jetstream.
Euler Flutter Analysis of a Complex Aircraft Configuration 

John T. Batina, Principal Investigator 

Co-Investigator: Russ D. Rausch 

NASA Langley Research Center/Purdue University 

Research Objective 

To demonstrate a time-marching aeroelastic analysis procedure 

for complex aircraft configurations. 

Approach 

Modifications were made to a three-dimensional, upwind-type 

Euler code that used unstructured grids of tetrahedra, including 

the structural equations of motion for their simultaneous timeintegration with governing fluid-flow equations. A novel aspect 

of this capability is that the solution is obtained using an 

unstructured grid made up of tetrahedra. The code also includes 

a deforming mesh algorithm capable of treating general aeroelastic motions of complete aircraft configurations. 

Accomplishment Description 

A calculation was performed for a supersonic transport configuration at a free-stream Mach number (Mj of 0.907 and at 

0 degrees angle of attack. The configuration consists of a rigid 

fuselage and a flexible clipped delta wing with two identical 

slender under-wing bodies to simulate engine nacelles. The 

bottom and side views of the surface mesh are shown in the 

accompanying figure and show that cells have been clustered 

near the wing tip and around the simulated nacelles. The 

complete mesh for the configuration contains 323,818 tetrahedra 

and 59,429 nodes. An aeroelastic time-marching calculation

was performed at a value of dynamic pressure (q) that was found 

experimentally to correspond to flutter. The generalized displacements for the first three structural modes are shown in the 

figure. The near-neutrally stable transient of the first generalized 

displacement indicates that the computed aeroelastic transient is 

near the flutter point. Therefore, the computed value of dynamic 

pressure at flutter is in good agreement with the experimental 

value. 

Significance 

The result is believed to be the first three-dimensional flutter 

calculation obtained using the unstructured-grid methodology 

for a complex aircraft configuration. 

Future Plans 

To continue the development of a highly accurate and efficient 

solution algorithm for the Euler and Navier—Stokes equations for 

aeroelastic analysis of complex aircraft configurations, and to 

perform additional calculations to further validate the method. 

Publication 

Rausch, R. D.; Batina, J. T.; and Yang, H. T. Y. "ThreeDimensional Time-Marching Aeroelastic Analyses using an 

Unstructured-Grid Euler Method." AIAA Paper 92-2506, 

Apr. 1992. 

Euler flutter analysis of a complex aircraft configuration using unstructured-grid methodology.
Aerodynamic Shape Optimization 

Oktay Baysal, Principal Investigator 

Co-Investigators: Mohamed E. Eleshaky and Greg W. Burgreen 

Old Dominion University 

Research Objective 

To develop a new aerodynamic shape method to optimize the 

nozzle of the National Aero-Space Plane (NASP) and to optimize 

the control surfaces and propulsion integration of the HighSpeed Civil Transport (HSCT). 

Approach 

A computational fluid dynamics (CFD) algorithm and an 

optimization algorithm are directly coupled by using a discrete 

sensitivity analysis methodology. The optimizer is used to 

determine the shape. To avoid restricting the shape to any class 

of surfaces, the aerodynamic shape is defined by the computational surface grid. The flow-field solution is obtained by a CFD 

analysis method when significant shape changes are made 

during the evolution toward the optimum shape. A flow-field 

extrapolation method called flow prediction is used when the 

shape changes are small. 

Accomplishment Description 

The optimization method is implemented in a new computer 

program known as Aerodynamic Design Optimization using 

Sensitivity (ADOS) analysis. The ramp shape of the NASP nozzle 

was optimized for maximum axial thrust coefficient. Starting 

from three completely different and arbitrary initial shapes, 

identical optimum nozzle shapes were obtained resulting in 

significant increases (up to 50%) of wall shape dependent thrust 

components. The ADOS code required 3.5 Cray Y-MP hours in 

each of these cases and used 18 megawords of memory. In order 

to account for viscous effects, the shape optimization methodol-

_142 2 6 10 14 18 

Normalized axial distance (x/H)

ogy was extended to the thin-layer Navier—Stokes equations. The 

method was then proven by successfully optimizing the shape of 

a transonic airfoil at 0 degree angle of attack to achieve a 

minimum drag while obtaining a lift above a specified value. 

This required 7 Cray Y-MP hours and used 35 megawords of 

memory. 

Significance 

The methodology contributes to state-of-the-art aerodynamic 

design optimization and will be directly applied to the design of 

the HSCT. 

Future Plans 

The two-dimensional sensitivity analysis computer program will 

be improved to reduce memory requirements and then will be 

expanded to three dimensions. The shape optimization methodology will be extended to handle complex geometries through 

domain decomposition procedures. The overall optimization 

method will be applied to the design of HSCT control surfaces. 

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1.0

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0.20

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M,=0.8,Re=0.5x106,a=0 

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Nondimensional axial distance (x/c) 

Optimum shape 

2

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2 —2

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Comparison of the final nozzle ramp shapes. Airfoil shape and distribution of surface pressure coefficients. 
Turbulent Boundary Layer with Heat Transfer 

Diane M. Bell, Principal Investigator 

Stanford University 

Research Objective 

To construct data bases for the heated turbulent boundary layer 

in an unperturbed environment and in the presence of freestream turbulence (FST). The data bases will be used to 

(1) quantify the poorly understood scalar transport mechanisms, 

(2) determine how FST interacts with the momentum and 

thermal boundary layers, and (3) improve existing turbulence 

models. 

Approach 

Direct numerical simulations (DNS) of the unsteady threedimensional incompressible Navier—Stokes equations are used to 

compute spatially developing turbulent flat-plate boundary 

layers with an arbitrary number of passive scalars. The fully 

spectral code uses no turbulence models. 

Accomplishment Description 

This simulation has a momentum thickness Reynolds number 

range of 300-700, scalar solutions for Prandtl numbers of 0.1, 

0.71, and 2.0, and an unperturbed free stream. The previously 

computed statistically stationary coarse-grid (256 x 55 x 96) 

solution was stabilized on two intermediate meshes 

(384 x 61 x 160 and 600 x 65 x 216) and on the fully resolved 

fine grid (720 x 70 x 256). This required 85, 150 and 250 singleprocessor CPU hours respectively. Approximately 10% of the 

fine-grid statistical sample was also collected. Since a fine-grid 

sample spanning three eddy turnover times requires approximately 350 CPU hours, the cost of the entire simulation is 

estimated at 1,000 CPU hours. The memory requirements for the 

three grids range from 4-12 megawords, with 35-60 megawords 

of solid-state storage-device memory. In general, the preliminary 

results agree well with experiment and boundary-layer theory. 

However, inner wall scaling of the mean profiles suggest 

different functional forms of Coles' low-Reynolds-number 

corrected hydrodynamic wake parameter and of the thermal 

wake term. Also, as shown in the first figure, the normal 

turbulence flux differs significantly from wind tunnel data. The 

similarity of the computed normal flux in air and the Reynolds 

stress is consistent with the similarity of their respective correlation coefficients and with the similarity of the axial momentum 

and thermal intensities. Since these turbulence quantities also 

compare well with measurements, we conclude that the 

experimental normal turbulence flux is too low. This low flux 

results in an overprediction of the experimental turbulent Prandtl 

number. In the second figure, the steep rise of the turbulent 

Prandtl number predicted from experimentally derived turbulence models is not found in the boundary-layer DNS, or in 

Kim's channel-flow DNS.

Significance 

Turbulence models for scalar transport are only accurate for 

simple geometries and restricted flow parameter ranges; no 

existing models adequately predict the effects of FST. The DNS 

results provide the detailed information necessary to investigate 

the flow physics and to develop predictive tools. 

Future Plans 

The collection of fine-grid statistical samples for the unperturbed 

case will be completed. Three papers are planned covering 

mean flow and turbulence statistics analysis, structural analysis, 

and turbulence modeling. The coarse-grid stabilization of two 

flows with FST will also be completed. 

Normal turbulent flux 

1.00

- Ree Pr 

- 1500 

Turbulent Prandtl number 

2.00 ReØ Pr 

925 - 0.1 and 0.7 1.75 • 1500 -- 2.0 

\ : 1950 0.1 and 0.7 Kim 

1.50 / .. • 2150 2.0 Kim 

.50 .

(b) I.

I 

0 .2 .4 .6 .8 1.0 1.2 

y/ö

(a) Normal turbulence flux. (b) Turbulent Prandtl number.

9 
(a) 

Two-dimensional convergent-divergent ejector/suppressor 

exhaust nozzle. (a) Core primary nozzle—vertical centerplane. 

(b) Mixing section—axial planes.

Three-Dimensional Exhaust-Nozzle Flow Fields for High-Speed Civil Transport 

Robert F. Bergholz, Principal Investigator 

Co-Investigator: William M. Turner 

General Electric Aircraft Engines

exhaust-plume analyses will be conducted and the results 

compared to hot-flow exhaust-plume laser-velocimetry data. 

Preliminary aero-design studies of new nozzle concepts are 

under way. 

Research Objective 

To support the large-scale, three-dimensional aerodynamic 

design of advanced High-Speed Civil Transport (HSCT) exhaust 

systems with acoustic suppression. The focus is on the analysis 

of complex internal-nozzle and exhaust-plume flows for 

multistream ejector and fluid-shield nozzles. The technical 

challenge is to achieve acceptable acoustic and thrust performance in the suppressed mode. This requires a detailed understanding of the dominant flow features (such as shocks, viscous 

mixing, and separation) for various classes of nozzles. 

Approach 

Currently available three-dimensional viscous-flow solvers 

(PA133D, CFL3D, PARC3D) are being applied in the analysis of 

full scale HSCT exhaust-nozzle designs and in validation studies 

using recently acquired subscale model test data. Candidate 

nozzle designs have both two-dimensional convergent-divergent 

(2DCD) and axisymmetric configurations, and may incorporate 

lobed suppressor elements, acoustic liners, flow splitters, or 

dual-stream fluid shields. Computational fluid dynamics (CFD) 

analyses are being performed to study the aerodynamics of mixer 

lobe geometry, secondary flow pumping, three-dimensional 

internal-mixing behavior, exhaust-plume decay, and nozzle 

performance loss mechanisms. 

Accomplishment Description 

The PA133D code, with a high-Reynolds-number two-equation 

turbulence model, was applied in a large scale validation study 

of a 2DCD suppressor/ejector nozzle. The analysis was made for 

a nozzle pressure ratio of 4.0, an ejector mixing section area 

ratio of 1 .2, and a tunnel free-stream Mach number of 0.27. The 

multiblock grid consisted of 1 zone/1 block in the primary core 

(196,725 points) and 6 zones/1 2 blocks (1,136,995 points) in th 

secondary inlet and mixing section. The first cycle of calculations required approximately 18 megawords of memory and 

65 Cray-2 hours. The two accompanying figures show the flow 

field in a central vertical plane through the supersonic primary 

core and in successive cross-stream planes through the mixing 

section. Notable flow-field features are the internal supersonic 

expansion, the recompression shock and boundary-layer 

interaction near the nozzle exit, and the downstream development of the rti( jI nrfl 

Significance 

The HSCTe'iijtunderstanding the internal shock structure and three-dimensienii 

mixing behavior of high-flow ejector nozzles under static and 

flight conditions. CFD analysis tools have been successfully 

applied in the generation of complex three-dimensional lobed 

mixer shapes now undergoing wind tunnel testing. 

Future Plans 

Further validation studies of 2DCD ejector nozzles and more 

recent fluid-shield nozzles are planned. Three-dimensional

TOTAL TEMPERATURE( R)

MIXING SECTION 

ft 1ff .

- 

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soojo 
Electromagnetic Scattering from Large 

Marek K. Bleszynski, Principal Investigator 

Rockwell International Science Center

Three-Dimensional Targets

Re Xs 

Research Objective 

To extend and validate computational capabilities of the 

adaptive integral method (AIM) code based on a numerical 

solution of Maxwell's equations in the integral form. 

Approach 

The AIM solver designed for large-scale three-dimensional 

electromagnetic computations was developed at Rockwell 

International during the last two years. The code incorporates 

modeling of general radar absorbing materials and structures 

encountered in actual aircraft design. The AIM solver is characterized by an approximately linear increase of memory and 

computational time with the number of computational elements 

N. The order of N (0(N)) performance is achieved with a 

combination of iterative solution, fast multilevel convolution 

algorithm, and elements of spectral technique enhanced by 

locally applied conventional method of moments. 

Accomplishment Description 

Approximately 20 hours were used to carry out validation of the 

new elements of the code involving local boundary conforming 

corrections to the spectral technique and new multitasking 

algorithms for fast Fourier transforms. Another 80 hours were 

spent on applications to actual targets of interest in aircraft 

design. Each calculation took 90 Cray Y-MP seconds using a 

single processor. 

Future Plans 

An automatic generation of a periodic uniform grid in the 

object's interior and a locally deformed grid at its boundaries 

and at material interfaces will be developed, implemented, and 

validated. Subsequently, an extensive and versatile set of new 

basis functions, which will improve the accuracy and extend the 

applicability of the method to a wider frequency range, will be 

used. Calculations involving a full aircraft will be attempted. 

Utilizing the AIM code optimization capabilities, a variety of 

optimization problems involving leading- and trailing-edge 

design and surface wave propagation will be carried out. 

Significance 

The ability to predict electromagnetic scattering and radiation of 

large three-dimensional structures is of considerable practical 

importance in aircraft design and biomedical applications. 

Computations based on the AIM code facilitate the task of 

finding the optimal target parameters which lead to desired 

electromagnetic signatures. 

Publication 

Bleszynski, M. and Jaroszewicz, T. "A New 0(N) Integral 

Equation Solver for Large-Scale Electromagnetic Computations." 

1992 URSI Conference. Boulder, CO, 1992.

j

k in 

r( 

/k 

Distribution of the real part of the current component along 

the polarization vector of the incident wave on a rhomboidal 

5.5A x 77.5A (a) perfectly conducting plate and (b) perfectly 

conducting plate with resistive edges, illuminated by a grazing, 

transverse magnetic polarized plane wave. The axes of the 

reference system x 1 , x2 are taken along the plate sides.

IV
Short Takeoff and Vertical Landing Aircraft Thermal/Acoustic Loads 

William W. Bower, Principal Investigator 

Co-Investigator: Robert E. Childs 

McDonnell Douglas Research Laboratories/Nielsen Engineering and Research, Inc. 

Research Objective 

To predict the dynamic behavior and thermal/acoustic loads 

associated with single and multiple supersonic impinging jets. 

Approach 

The calculations of unsteady flows are based on large-eddy 

simulation. With this approach, the mean flow plus the dominant turbulent and acoustic modes are resolved and the smallscale turbulence that cannot be resolved is modeled. The 

simulations are based on the compressible Navier—Stokes 

equations with a simple subgrid-scale turbulence model, which 

are solved with explicit Runge—Kutta time integration coupled 

with fourth-order finite-volume discretiza-tion. Nonlinear 

artificial dissipation is used to capture shocks. Characteristic 

nonreflecting boundary conditions are employed to pass 

acoustic and vortical modes out of the computational domain. 

Accomplishment Description 

The goal was to extend the axisymmetric supersonic free and 

impinging jet flow analysis to three dimensions. A preliminary 

simulation has been run for a jet issuing from a round nozzle 

with choked flow at a nozzle pressure ratio of 1,265, static 

temperature ratio of 0.833, expanded-jet Mach number of 1.19, 

and Reynolds number of 50,000. The jet's dynamical structure 

up to five nozzle diameters is depicted in the accompanying 

figure for a dimensionless time of 170 based on the nozzle 

radius and acoustic speed. The contours of static pressure 

referenced to the free-stream value are given on the midplane of 

the jet as viewed from the side. The slightly subambient pressure 

regions highlight the low-pressure zones associated with the 

vortex cores. The structures are clearly not axisymmetric and 

appear to be helical based on an evaluation of sequential plots 

through the jet. This computation required approximately 

14 Cray Y-MP hours and 11 megawords of memory for a grid 

containing 59 points in each coordinate direction. It is important 

to remember that this is a preliminary result on a coarse grid in a 

small computational domain. Some noise reflection/generation 

at outflow boundaries and truncation errors certainly effect the 

computed solution. 

Significance 

A vertical takeoff and landing (VTOL) aircraft hovering on the 

thrust of jet-engine exhaust in close proximity to the ground can 

experience very high levels of acoustic and turbulent pressure 

fluctuations. The pressure fluctuations combined with elevated 

temperatures from the jet engine exhaust pose structural fatigue 

problems. The physical processes occurring in hot impinging jets 

that produce these loads are very complex. Interactions among 

turbulence, the shock structure of the exhaust jets, and the 

ground give rise to acoustic levels high enough to accentuate the 

jet turbulence. Phase locking and feedback occur, which may 

produce high sound-pressure levels at discrete frequencies. The 

present work uses simulations to provide more insight into 

processes relevant to VTOL aircraft ground effects.

Future Plans 

The simulations will be extended to three-dimensional round 

and non-round impinging jets, and comparisons of the predicted 

flow and acoustic fields will be compared with experimental 

data. 

Publication 

Wlezien, R. W.; Bower, W. W.; Childs, R. E.; Howe, M. S.; 

and Kibens, V. "Experimental and Computational Investigation 

of Supersonic STOVL Jet Flow and Acoustic Fields." NASA 

CR-i 89547, Jan. 1992. 

Contours of static pressure referenced to the free-stream value at 

the midplane of the jet as viewed from the side. Inside the 

highlighted regions bounded by the white boundary lines, the 

high value is white, 0. 98, and the low value is black, 0.90. For 

the field outside the white boundary lines, the scale changes and 

the dark blue corresponds to a normalized pressure of unity. 

12
Local Passive Scalar Dispersion in a Turbulent Boundary Layer 

James G. Brasseur, Principal Investigator 

Co-Investigators: P. K. Yeung, Samir Khanna, and Brian P. Moquin 

Pennsylvania State University/Georgia Institute of Technology 

Research Objective 

To analyze the structural interrelationships among passive scalar 

and hydrodynamic field structure during transport of passive 

scalar within wall-bounded flows and the interpretation of 

hydrodynamic field dynamics in terms of concentration field 

dynamics. 

Approach 

Spalart's "fringe" algorithm has been implemented for a direct 

simulation of a naturally growing, zero-pressure-gradient, flatplate boundary layer with sources of passive scalar. The 

algorithm makes use of fringe regions at opposite streamwise 

extremes of the computational domain in which sinks are placed 

to remove momentum and passive scalar. Additions to the code 

include a "heated" spanwise strip at the wall and a source term 

in the scalar equation for the introduction of passive scalar 

within the computational domain. Passive scalar sources with 

Schmidt number 1 are placed (1) along a heated spanwise strip 

at the wall, (2) at the outer edge of the boundary layer, and 

(3) within the narrow log layer. The scalar simulations begin 

from a stationary hydrodynamic field on a 720 x 70 x 256 grid 

with momentum-thickness Reynolds numbers from 350-780 in 

the streamwise direction. Full field-velocity, velocity-gradient, 

and passive-scalar data are collected every 4 viscous time units. 

Accomplishment Description 

Initial graphical analysis has confirmed most of the structural 

features of the vorticity and velocity fields, including high 

intensity sheet-like structures in the near-wall vorticity field, 

lower magnitude hook-shaped concentrations of vorticity rising

from the sublayer, and streaks of concentrated turbulent kinetic 

energy rising from the wall. We have begun analysis of the 

passive-scalar evolution from the heated wall strip. Regions of 

highest concentration, near the passive-scalar source, appear to 

be strongly influenced by the near-wall streaks, whereas regions 

of low scalar concentration (away from the wall) appear to be 

under the influence of more intense vorticity fluctuations. A 

typical job requires 4 Cray Y-MP hours and 10 megawords of 

memory with 60 megawords solid-state storage-device memory. 

Significance 

Our goal is to answer several questions. From a fundamental 

point of view, what information does dye, smoke, or temperature 

concentration in a wall-bounded turbulent-shear flow provide 

about the underlying structure and dynamics of the hydrodynamic field? What local dynamical processes contribute to net 

scalar and momentum transport toward and away from the wall? 

From a modeling point of view, how are mean-flow statistics 

established? What local dynamical processes contribute to 

global statistics? From an applications point of view, in what 

ways can local heat transfer be modified through manipulation 

of local dynamics in the turbulent boundary layer? 

Future Plans 

Careful analysis of the large amounts of data being collected is 

just beginning. Analysis will proceed at two levels, the visual 

and the quantitative. Much use will be made of the methodology 

in which intermittent regions are extracted and conditioned with 

different characteristics of interest in the study. 

*. 

4b\ 

Two isocontours of passive scalar dispersed from a "heated" spanwise strip at the wall. Streamwise, 

normal, and spanwise extent of the subdomain are 525, 238, and 675 wall units. High intensity scalar 

(red) concentration evolves to low intensity (blue) concentration over 18 viscous time units.

13
Global-Venus/Solar-Wind Interaction 

Stephen H. Brecht, Principal Investigator 

Berkeley Research Associates 

Research Objective 

To obtain a better understanding of how the solar wind interacts 

with the unmagnetized planets, Venus and Mars. Specifically, 

we want to understand the dynamics of the shock and magnetic 

barrier location as a function of solar-wind parameters and 

ionospheric-mass loading rate. 

Approach 

We have developed a three-dimensional hybrid particle code 

that simulates the electrons as a fluid and the ions as individual 

particles. Multiple species of ions are included, as is their 

complete kinetic behavior. 

Accomplishment Description 

The three-dimensional hybrid code HALFSHEL is running with 

both Navier and Reynolds codes. During the past year we 

simulated solar-wind flow past obstacles that are representative 

of Mars and Venus. These simulations included the nose region 

of the magnetosphere of these planets down to the terminator 

line. Many of the simulations were modified to include multiplespecies ionospheres. We demonstrated that the presence of the 

ionospheric ions affects the position of the Venusian bow shock. 

In addition, the HALFSHEL code has shown that Mars does not 

form a bow shock in the subsolar region, but only on the flanks. 

The results have been compared to and agree with actual orbit 

data from Phobos-2. The simulations using HALFSHEL require 

approximately 40 megawords and about 32 Cray Y-MP hours. 

Smaller simulations are performed which only require 

25-30 megawords and 10 Cray Y-MP hours. 

Significance 

The results of the simulations and the visualization of the results 

are changing the way scientists view the interaction of the solar 

wind with planets such as Venus and Mars. This is particularly 

true of the Mars simulation results, which indicate the lack of a 

subsolar bow shock and are being compared with the orbital 

data from the Phobos-2. 

Future Plans 

The HALFSHEL simulations of Mars and Venus will continue. 

We will investigate the role of Martian ionosphere and hot 

hydrogen corona on the flank shocks. Further investigation on 

understanding Venusian dynamics will be made this year. 

Publications 

1. "Global Hybrid Simulations of Unmagnetized Planets: 

Comparison of Venus and Mars."]. Geophys. Res. 96 (1991): 

11209. 

2. "Three Dimensional Simulation of the Solar Wind Interaction 

with Mars." Submitted to]. Geophys. Res., 1992. 

3. "Mass Loading Effects of the Venusian Bow-Shock Location." 

Submitted to Geophys. Res. Lett., 1992.

"Modeling Studies Related to Solar Interaction with Mars and 

Venus: A Review." Invited paper presented at General 

Assembly IUGG, Plasma Environments of Mars and Venus, 

GAM 4.7, Vienna, Austria, Aug. 1991. 

"Hybrid Particle Simulations for Large-Scale Phenomena." 

Invited paper presented at General Assembly IUGG, MHD 

Modeling and Kinetic Simulations as Tools of Heliospheric 

Physics, GAM 4.3, Vienna, Austria, Aug. 1991. 

HALFSHEL simulation of solar wind interaction with Mars. 

The solar wind ion trajectories are plotted. Note the structure in 

the magnetic field surface and the finite gyroradius behavior of 

the ions. 

14
Hot-Gas-Manifold Flow Simulation 

Richard C. Buggeln, Principal Investigator 

Co-Investigator: Sang-Keun Choi 

Scientific Research Associates, Inc. 

Research Objective 

To develop a numerical simulation of the flow in a hot gas 

manifold pilot model for a staged combustion-cycle manifold. 

This problem is a complex three-dimensional flow situation 

requiring a full three-dimensional Navier—Stokes solution. A 

relatively fine grid, 1.5 x 106 grid points, and a relatively coarse 

grid solution, 0.707 x 106 grid points, are to be obtained, with 

the latter solution to be obtained with and without wall functions. 

Approach 

The research objective is being pursued using the Scientific 

Research Associates, Inc., Navier—Stokes code, which solves the 

full Navier—Stokes equations using the Briley—McDonald 

alternating direction implicit approach, with the grid generated 

by EAGLE. With this approach, the entire domain is treated as 

one computational zone, thus avoiding internal zone boundaries. 

Accomplishment Description 

To date, the fine grid simulation has converged. Typical runs for 

the fine grid simulations required 6 megawords of memory and

5,000 CPU seconds. The computational region consists of a 

preburner, a turbine passage, a turnaround duct, an annular 

duct, a transfer duct, and a main injector bowl. A velocity vector 

plot in the symmetry plane is shown in the accompanying figure. 

The asymmetry in the turnaround duct and the annular duct are 

evident, as is the highly three-dimensional nature of the injector 

bowl, which has multiple recirculation zones. The second case 

(the grid containing 0.707 x 106 grid points) is well under way 

and approaching convergence and the third case (the wall 

function case) has been initiated. 

Significance 

The results demonstrate the ability of Navier—Stokes solvers to 

simulate the highly complex viscous flows in practical flow 

configurations. The ongoing calculations will serve to assess grid 

refinement and wall function issues. 

Future Plans 

The calculations will be continued and an assessment of results 

made. 

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2.40000 

2.50000 

2.60000 1 / 

'1)1100 

2.80000 

2.90000 

3.00000 l'' 3.10000 

TL20000 

Velocity vector plot for a hot gas manifold in symmetry plane.

15
Simulation and Modeling of Turbulence and Flow Acoustics 

Alan B. Cain, Principal Investigator 

Co-Investigators: Dennis F. Fuglsang and Nagi N. Mansour 

McDonnell Aircraft Company/NASA Ames Research Center 

Research Objective 

To compute, predict, and control complex shear-layer flow 

phenomena (particularly flows with acoustic feedback loops). 

Approach 

Finite-difference schemes are used in direct simulations of 

unsteady-cavity flows. To assure cost effectiveness, several 

schemes are evaluated analytically. It is advantageous to use grid 

sequencing with a one algorithm calculation. A more sophisticated scheme can be used for the final details. This work made 

use of the NASA Langley CFL2D code and several options in the 

three-dimensional MCAIR NASTD code. The final threedimensional oblique-mode results are based on second-order 

upwind spatial differencing and second-order time integration. 

Accomplishment Description 

Transonic cavity-flow results are in good agreement with 

experimental data for the surface sound pressure levels. More 

detailed evaluation indicates good prediction of resonant 

frequencies, but there is some scatter in the amplitude of the 

individual peaks. The results for a natural 110w produce a strong 

harmonic response; however, a harmonic forcing of the flow at 

nonresonant frequencies appears to produce a very energetic 

and chaotic response. In addition, a natural feedback loop is 

identified that can result in a "self-sustaining" effect of any 

temporary forcing. Analysis of the evolution of the flow from an 

oblique-mode initial condition is in progress. A typical threedimensional unsteady solution requires from 50-500 CPU hours 

and 16 megawords of memory for 662,000 cells. 

Significance 

This work shows that direct simulation is a meaningful way to 

analyze and predict weapons-bay-cavity behavior. In a more 

general context, this work suggests that a priori control schemes 

may easily fail to achieve their desired result due to the multiplicity of solutions to nonlinear problems. The initial state of the 

system may determine the ultimate response. 

Future Plans 

Further work is planned, including the use of higher-order 

schemes, the role of oblique modes (results are currently being 

analyzed), and the application of this research to the supersonic 

jet screech problem. 

Publication 

Fuglsang, D. F. and Cain, A. B. "Evaluation of Shear-Layer 

Cavity-Resonance Mechanisms by Numerical Simulation." 

AIAA Paper 92-0555, Aerospace Sciences Meeting, Reno, NV, 

Jan. 1992.

I..

Solution for a Mach 0.85 free-stream flow over an open cavity. 

Vorticity and pressure contours for an oblique-mode initial 

condition. 

16
CP ç

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

.8 

1.2 

Automated Transonic Wing Design 

Richard L. Campbell, Principal Investigator 

Co-Investigator: Leigh Ann Smith 

NASA Langley Research Center 

Research Objective 

To create a design approach for geometrically complex configurations by combining the direct iterative surface curvature (DISC) 

design method with the unstructured-method three-dimensional 

(USM3D) Euler code. 

Approach - 

The DISC design method modifies the surface curvatures and 

slopes of an initial geometry to match target pressure distribution. This method has been successfully coupled with a number 

of codes, varying in complexity from small disturbance to 

Navier—Stokes, and has been applied to the design of aircraft 

components such as airfoils, wings, winglets, and nacelles. The 

choice of USM3D and the unstructured gridding technique, 

VGRID3D, allows for the study of full configurations and thus 

the potential to design any component in the presence of other 

components. Critical to this technology was the development of 

a dynamic mesh routine for grid perturbation based on the 

spring analogy technique. This allows the modifications to the 

surface definition computed by the design method to be easily 

assimilated into the surrounding grid. The spring analogy 

technique was modified to reduce grid distortion. 

Accomplishment Description 

The DISC method was successfully coupled with the USM3D 

unstructured Euler flow solver to create a design tool (DUSM3D) 

that utilizes the advantages of unstructured grid technology. The 

coupling was verified by designing to known wing cross-sections 

for an existing transport. While this technique is currently 

implemented only for wing design, it has the potential for 

extension to all surfaces of an aircraft. For the verification

exercise, an existing generic transport configuration known as 

the low-wing transport was analyzed in USM3D. The resulting 

wing pressure distributions were used as target pressure distributions for the design. An initial configuration was created by 

replacing the transport's supercritical airfoils with NACA 0006 

airfoil sections. The DUSM3D method was applied to the 

starting configuration and allowed to iterate until convergence. 

This particular configuration requires about 10 megawords of 

memory and 1.5 Cray Y-MP hours to converge (only a slight 

increase in CPU time over a regular analysis run). The accompanying figure shows the results of this design effort for two span 

stations. Pressure distributions and airfoils are shown for the 

original wing (target), the starting configuration (initial), and the 

resulting design (final). The plots show that the design method 

successfully reproduced the target pressures and airfoils, in spite 

of the large differences between the starting and target shapes. 

Significance 

This exercise demonstrates the ability of the DISC design method 

to interact with an unstructured method and to produce verifiable wing designs. The ability to design using an unstructured 

method and the advantages in ease of grid development should 

reduce the time required to obtain advanced solutions for 

geometrically complex configurations. 

Future Plans 

We will expand the implementation of DISC in USM31D so that 

wings, nacelles, winglets, and other configuration components 

can be easily designed. USM31D can be used to readily obtain 

solutions for complex geometries, and the DISC method can be 

used to design any surface of a given geometry. 

Wing-design results for low wing transport, M = 0.77.

17
Simulation of a Time-Developing Incompressible Plane Wake 

Brian J. Cantwell, Principal Investigator 

Co-Investigators: R. Sondergaard, J. Soria, and J. H. Chen 

Stanford University/Sandia National Laboratory 

Research Objective 

To study the physics of fine-scale motions of turbulent-shear 

flows using topological methods applied to a three-dimensional 

direct numerical simulation of a time-developing, incompressible plane wake. 

Approach 

The project consisted of two components. The first was to 

generate sets of moderately high Reynolds number (Re) direct 

numerical simulations of a time-developing incompressible 

plane wake. These data sets were then studied in the space of 

invariants of the velocity gradient tensor using recently developed topological analysis tools. The direct numerical simulation 

employed a version of an existing free-shear-layer code developed by Rogers and Moser. Linear stability theory was applied to 

determine the unstable modes used to initiate the computations. 

A set of low-Reynolds-number (Re = 500), low-resolution test 

runs was performed to verify the code and to pick an appropriate 

set of initial disturbances for the higher Reynolds number (up to 

Re = 4,000) runs. The initial disturbances chosen were combinations of two-dimensional waves at the most unstable wavelength 

and pairs of oblique disturbances at the two-dimensional 

fundamental and subharmonic wavelengths. Topological 

analysis was performed on the data using post-processors and 

graphics routines we developed. 

Accomplishment Description 

Four low-Reynolds-number and four higher-Reynolds-number 

simulations were run. A typical simulation required between 

30-40 restart files. For the higher-Reynolds-number simulations, 

each restart file required approximately 4 Cray Y-MP hours and 

20 megawords of memory. Topological analysis of the fine 

scales, which is still in progress, has revealed similarities with 

flows previously studied. For regions of high dissipation rates, 

the rate of strain has an unstable-node/saddle/saddle-type 

topology (two positive and one negative principal strain rate), the 

entropy has a magnitude comparable to that of the dissipation 

rate, and the vorticity vector tends to align with the intermediate 

principal rate-of-strain direction. 

Significance 

This project enhances our understanding of the basic properties 

of turbulent motion. Better understanding of the physics of the 

fine-scale structures is needed to develop improved subgrid scale 

turbulence models that will enhance our ability to solve technical problems involving turbulent flow. 

Future Plans 

We will (1) modify the simulation code to run on one of the new 

massively parallel computers now available, allowing us to run 

simulations at a higher Reynolds number than on the Cray Y-MP

and (2) extend the topological analysis to a wider variety of 

turbulent flows and include studies of the pressure-gradient field 

to determine features of the fine-scale turbulent motions that are 

universal and are flow dependent. The effect of Reynolds 

number and how the fine scales evolve will also be explored. 

y

oil 

I 

4i

ix 

(a)

(b)

Simulation of a three-dimensional incompressible plane wake at 

Re = 500. (a) Physical space (contours of constant entropy). 

(b) Strain-rate invariant space. 

18
Prediction of Advanced Rotor Performance 

Frank Caradonna, Principal Investigator 

Co-Investigators: John Bridgeman and K. Ramachandran 

U.S. Army Aeroflightdynamics Directorate, AVSCOMVNASA Ames Research Center 

Research Objective 

To develop computational fluid dynamics (CFD) methods for the 

prediction of the aerodynamic loads and acoustics of helicopter 

rotor blades. Principal areas of investigation include steady and 

unsteady transonic flow, viscous contributions to loads, and the 

influence of the rotor wake. Research emphasis is placed on the 

development of codes for immediate use as rotor design tools. 

Approach 

Three-dimensional potential equations are solved for an inviscidrotor flow simulation. The rotor wake is modeled with a combined 

Eulerian—Lagrangian vorticity-embedding scheme to simultaneously compute rotor aerodynamics and free-wake solutions. 

An entropy correction term is added to the potential equations to 

determine viscous effects using either integral- or finite-difference 

boundary layer methods. 

Accomplishment Description 

HELIX-11, the free-wake CFD code, has been applied to analyze 

a two-bladed rotor in low-advance-ratio forward flight. At low 

advance ratios, the influence of the rotor wake is important in 

hover. Unlike hover solutions, the flow is unsteady and it is 

necessary to grid multiple blades. A grid containing approximately 500,000 points was used for this study with typical run 

times of 3 hours and 16 megawords of memory. A new unsteady 

full-potential code, FPX, was developed for application to rotors 

and propellers. Multiple topologies are treated, entropy and 

viscosity effects are included, and axial flow capability allows 

for propeller treatment. The code has been used to predict 

fuselage wall pressures in the cabin of the XV-1 5 standard 

aircraft, advanced technology blades, and has been compared 

with experimental results from the Army Aeroflightdynamics 

Directorate hover chamber for a NACA 0012 rectangular blade. 

Typical runs take 0.5 hour and 8 megawords of memory. 

Significance 

Aerodynamic flow phenomena are principally responsible for all 

crucial rotorcraft performance limitations. Accurate coupledrotor aerodynamics and free-wake calculations are mandatory 

for hover performance prediction and forward-flight analysis. 

Tilt-rotor configurations may play a significant role in both 

commercial and military markets in the near future. Prediction 

methods capable of analyzing the performance and acoustics of 

these vehicles will be vital to providing quality designs. 

Future Plans 

The HELIX-11 free-wake methodology will be applied to forwardflight investigations for several rotors for which experimental or 

flight-test data are available. FPX will be used to predict V22 

tilt-rotor aerodynamics and compared with experimental data. 

In a new effort, the vorticity confinement method will be 

incorporated into an existing Euler/Navier—Stokes code and 

demonstrated.

Publications 

1. Bridgeman, J. 0.; Caradonna, F. X.; and Prichard, D. "The 

Development of a CFD Potential Method for the Analysis of 

Tilt-Rotors." Presented at the AHS Technical Specialists 

Meeting on Rotorcraft Acoustics and Fluid Dynamics, 

Philadelphia, PA, Oct. 1991. 

2. NsiMba, M.; Ramachandran, K.; and Caradonna, F. X. 

"Computation and Validation of Hovering Rotor Performance." Paper No. 91-41, presented at the 17th European 

Rotorcraft Forum, Berlin, Germany, Sept. 1991. 

HELIX-I1-computed wake geometry for a two-bladed rotor in 

forward flight (advance ratio = 0.1).

19
Experiment 

13 JaIcuiati>/ 

10 

5 a) 

V 

w 

0

—5 

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—i

Aerothermal Analysis of Hypersonic Defense Interceptors 

R. R. Chamberlain, Principal Investigator 

Co-Investigators: Lawrence W. Spradley and Kenneth E. Xiques 

Adaptive Research Corporation 

Research Objective 

To develop and improve the methodology for predicting 

complex hypersonic flows associated with defense interceptors 

concepts. The computational effort is related to the time accurate 

prediction of the protective shroud removal from the high endoatmospheric defense interceptor (HEDI) configuration. 

Approach 

The three-dimensional, unstructured, finite-element flow code, 

FELFLO, is used to simulate the removal of a single shroud petal 

at Mach 8. The code uses an arbitrary Eulerian-Lagrangian 

approach to track the body motion and adaptive remeshing to 

resolve the shock—shock interactions. The time history of the 

petal motion is computed and compared to wind tunnel data. 

Accomplishment Description 

Using the symmetry of the four shroud petals that protect the 

HEDI forebody, the unstructured grid is constructed for a single 

petal in one quadrant. Initial conditions for the external flow are 

assumed to be free stream everywhere except for the shroud 

cavity (the region between the petal and the forebody), which 

has no flow. The free-stream flow enters the cavity and raises the 

pressure so that the shroud petal begins to move. Rotation about 

a rearward hinge creates physical gaps between adjacent petals, 

which causes a subsequent reduction in the cavity pressure. This 

filling and venting phenomenon is characteristic of the shroudseparation event and is well predicted by the present calculation. 

The petal is eventually released into free flight so that it has both 

translational and rotational degrees of freedom. The first figure 

shows the pressure contours along a cut through the centerline 

of the petal just after it is released. The forebody and petal 

shocks interact while the high cavity-pressure vents to the rear 

like an exhaust jet. The petal rotation angle versus time is 

compared to this experiment and excellent agreement is 

obtained (second figure). A typical run requires 40 megawords 

of memory and 50 Cray-2 hours. 

Significance 

Many of the latest concepts for theater defense missiles have 

some kind of shroud separation problem. A significant portion 

of the design effort requires detailed knowledge of the flow 

environment inside the shroud cavity, particularly with regard to 

structural loading on the forebody, shroud petals, and hinges. 

This work demonstrates that accurate prediction of the highly 

transient phenomena associated with the separation event is 

trackable. 

Future Plans 

An effort to improve the spatial resolution of the present 

calculation is under way. Although the time history of the petal 

motion is significant, certain wave interactions inside the shroud 

cavity require further definition. Enhanced grid resolution and 

alternative remeshing strategies are being explored.

Pressure contours along a cut through the symmetry plane of a 

single shroud petal. 

—15

10 12 

Time (msec) 

Comparison of the time histories of the petal motion. 

20
20 

15 

10 

5 

0 

20 

15 

10 

5 

0 

2000 3000 4000 5000 7000 8000 

[A] 

U) 

c',l 

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Future Plans 

Further additions and enhancements to the radiative transition 

and kinetic models of the NEQAIR code are planned as data 

from experiments and computational chemistry simulations 

become available. Work will continue on enhancing the flow 

solvers developed in this study and on the study of low-density 

phenomena in expanding flows. 

Publications 

1. Moreau, S. and Laux, C. "A More Accurate Nonequilibrium 

Air Radiation Code: NEQAIR Second Generation." AIAA 

Paper 92-2968, AIAA 23rd Plasmadynamics and Lasers 

Conference, Nashville, TN, July 1992. 

2. Laux, C. and Moreau, S. "Experimental Spectrum of an LTE 

Air Plasma, and Comparison with NEQAIR Theoretical 

Predictions." AIAA Paper 92-2969, AIAA 23rd Plasmadynamics and Lasers Conference, Nashville, TN, July 1992.

Boost-Phase Detection Study 

Dean R. Chapman, Principal Investigator 

Co-Investigators: Forrest E. Lumpkin Ill, Robert W. MacCormack, and Stephane Moreau 

Stanford University/NASA Ames Research Center 

Research Objective 

To develop the capability to predict the radiation signature of 

the heated shock layer ahead of self-propelled bodies and the 

expanded flow behind them. 

Approach 

The radiation signatures are affected by thermochemical 

nonequilibrium and possibly by low-density phenomena 

leading to the failure of the linear constitutive relations in the 

Navier—Stokes equations. To capture the significant flow 

features, advanced thermochemical nonequilibrium models are 

refined and incorporated into two-dimensional flow solvers. The 

Burnett additions to the the linear Navier—Stokes constitutive 

relations are being employed to quantify low-density effects. For 

a computed flow field, the desired radiation emissivity and 

spectra can be computed using a recently modified version of 

NEQAIR, a nonequilibrium radiation code. 

Accomplishment Description 

The performance and range of applicability of the NEQAIR code 

have been greatly enhanced. Vectorization and algorithm 

restructuring yielded a significant performance improvement. 

Removal of very low intensity rotational and vibrational lines 

from the computation resulted in nearly two orders of magnitude 

in total improvement. Data more applicable to the conditions of 

interest for radiative transition probabilities were derived from 

computational chemistry simulations and experiments. A 

measured radiation signature compares well to one computed by 

NEQAIR for a plasma torch experiment designed to represent 

desired physical conditions, thus validating the new data. A 

recent NEQAIR computation, which includes an improvement in 

the kinetic model used to compute nitric oxide electronic 

excitation distributions, agrees favorably with signatures for 

high-altitude flights from the first SDIO bow shock ultraviolet test 

flight. Previous computations under-predicted the emissivity by 

several orders of magnitude. Reacting-gas simplified Burnett 

computations were performed for the heated shock layer. Only 

at high speeds and low densities were these additions found to 

affect radiation signatures. 

Significance 

This work enhances the ability to predict the radiation signatures 

of ballistic missiles during boost phase. This capability is 

indispensable when developing a system for targeting hostile 

missiles. The technologies developed in this study will enable 

more accurate predictions of heating loads and aerodynamic 

performance for future transatmospheric vehicles.

Comparison of measured and computed radiation signatures for 

a plasma-torch experiment.
Transonic and Supersonic Flow Past Aircraft Configurations 

Denny Chaussee, Principal Investigator 

Co-Investigators: Jolen Flores and Eugene Tu 

NASA Ames Research Center 

Research Objective 

To develop computational technology to assess and aid the 

design of future aircraft, to develop and validate the capability to 

simulate the viscous flow about wing—body—flap configurations, 

and to investigate the accompanying flow physics. This was 

done using a subsonic experiment of Frink. 

Approach 

A three-dimensional thin-layer finite-volume Navier—Stokes 

code was used to compute the subsonic flow about a slender 

wing—body configuration with trailing-edge flaps. A virtual-zone 

technique was used in the region of the flap—wing interface. The 

Baldwin—Lomax turbulence model was used. 

Accomplishment Description 

A simulation of turbulent subsonic flow about a slender 

wing—body with trailing-edge flap deflected 20 degrees at 

0 degrees angle of attack and Mach 0.4 was completed. The 

accompanying figure includes the computed surface pressure 

and the chordwise surface pressure on the inboard flap. The 

oil flow shows the complexity of the turbulent flow over a 

20 degree deflected flap. Comparisons with the experiment are 

qualitative for the pressures on the wing and flap system. A 

typical run required 30 megawords of memory and 50 Cray Y-MP 

hours. 

Significance 

The contribution to controls from flaps is important throughout 

the Mach and angle-of-attack range. Also, the capability to 

predict flows for various flap settings can decrease the amount 

of wind tunnel downtime by reducing the number of runs 

needed for flap parametrics. 

Future Plans 

The finite-volume Navier—Stokes code will be applied to 

different flaps, including split flaps and vortex flaps. Eventually, 

the problem will be run in a time-accurate mode with control 

feedback.

(a)

(b)

Subsonic wing—body—flap computation. M00 = 0.4, 

a = 0 degrees, Re = 5.41 x 106. (a) Chordwise surface pressure 

comparison center of inboard flap. (b) Pressures on the surface. 

22
Numerical Analysis of Three-Dimensional Separated Juncture Flows 

C. L. Chen, Principal Investigator 

Co-Investigator: C. M. Hung 

Rockwell International Science Center/NASA Ames Research Center 

Research Objective 

To understand the flow structure of various horseshoe vortex 

systems generated by a cylindrical protuberance mounted on a 

flat plate. 

Approach 

The Reynolds-averaged Navier—Stokes equations are solved 

to study this generic juncture flow with the unified solution 

algorithms (USA) series code. The USA code, which is a 

multizonal finite-volume solver, utilizes a high-resolution 

total-variation-diminishing scheme. A modified Baldwin—Lomax 

model incorporating Goldberg's backflow treatment is utilized in 

this study. The flow topology is investigated for various Reynolds 

numbers and Mach numbers. Results are compared with 

experimental results. 

Accomplishment Description 

The various complex laminar- and turbulent-horseshoe vortex 

structures were computed and analyzed. The upstream, outermost singular point was either a saddle point of separation or a 

saddle point of attachment. It moved upstream when the freestream Mach number increased. The size of the vortex structure 

increased dramatically due to shock-wave and boundary-layer 

interaction. The flow structures of multiple vortices (up to six) 

were computed. The computation of supersonic turbulent flow 

predicted the same features as those indicated by experimental 

results, such as upstream shock-wave/boundary-layer interaction 

and a classical horseshoe vortex system. The calculations also 

provided downstream wake/shock-wave interaction and the 

near-wake tornado-like vortex structure. The calculations took 

about 60 Cray-2 hours and 20 megawords of memory for 

each run. 

Significance 

Juncture flows are important in external aerodynamics (including 

multibody space shuttle junctures, wing—body junctures, and 

wing—pylon junctures), turbomachinery, submarine conningtower flow, and applications in meteorology and geology. The 

results of this effort lead to a better understanding of such 

complex flow fields and will assist in future design efforts. 

Future Plans 

This research will be continued and expanded.

Publications 

1. Chen, C. L. and Hung, C. M. "Numerical Study of Juncture 

Flows? AIAA 1 30, no. 7 (July 1992): 1800-1807. 

2. Hung, C. M.; Sung C. H.; and Chen, C. L. "Computation of 

Saddle Point of Attachment? AIAA J. 30, no.6 (June 1992): 

1561-1569. 

Overall structure of the juncture flow.

23
30 34 38 

x (streamwise) 

32 36 

x (streamwise) 

Outboard fan cowl 

- Computation 

—1.0 o Experiment 

C 

1.0 L 

28 

Euler Analysis of Turboprop and Turbofan Integration 

H. C. Chen, Principal Investigator 

Co-Investigators: T. Y. Su, T. J. Kao, and D. A. Naik 

The Boeing CompanyNiGYAN, Inc ./NASA Langley Research Center 

Research Objective 

To develop an effective computational capability for the analysis 

of turboprop and turbofan engine—airframe integration. 

Approach 

A generar, multiblock, multigrid Euler code (GMBE) has been 

developed for the analysis of complete airplane configurations 

with under-wing turbofans or superfans. The volume grid is 

generated independently. GMBE is a generalized version of a 

code that was successfully used to address the problem of wingmounted and aft-mounted turboprop engine—airframe integration. Powered nacelles may be simulated by prescribing inflow 

and outflow conditions at the fan inlet face and fan/core exit 

faces respectively. Flow-through nacelles are analyzed by 

prescribing solid-surface boundary conditions along the inner 

walls of the nacelle. 

Accomplishment Description 

GMBE was successfully executed for a generic low-wing 

transport with an advanced turbofan nacelle (bypass ratio 6).

The volume grid contains approximately 1.2 million grid 

points and is composed of 32 blocks. A typical solution takes 

4 Cray Y-MP hours. The accompanying figure shows that GMBEcomputed pressure coefficients compare well with the experimental data on the pylon and fan cowl. Additionally, the GMBE 

solution indicates that a change in toe angle and pylon trailingedge closure geometry will improve the propulsion integration. 

Significance 

The computational fluid dynamics code that was developed 

in this project is an effective tool for turbofan/superfan 

engine—airframe integration. The computed pressure distributions 

are used to identify, in terms of pressure peaks and gradients, 

undesirable flow regions in the vicinity of the pylon and nacelle. 

Future Plans 

GMBE has been coupled with a NASA Langley direct-iteration 

surface-curvature algorithm for the inverse design of installed 

turbofan and superfan nacelles and testing is under way. 

Euler solution of a low-wing transport airplane with turbofan nacelle. Coefficient of pressure range: —1.2 (red) to 1.0 (blue). 

= 0.77, CL = 0.55. 

24
Advanced Transonic Wing Concepts 

Lee T. Chen, Principal Investigator 

Co-Investigators: K. C. Chang, R. Pelkman, and A. Shmilovich 

Douglas Aircraft Company 

Research Objective 

To calibrate and validate three-dimensional Navier—Stokes codes 

for the analysis of advanced transonic-wing concepts for 

subsonic transports. 

Approach 

The thin-layer Navier—Stokes codes being studied, CFL3D and 

TLNS3D, were developed at NASA Langley. The CFL3D code is 

based on an upwind-difference scheme and the TLNS3D code is 

based on a central-difference algorithm. Wind tunnel experiments and flight-test data are used to validate the numerical 

solutions obtained for wing—body configurations. 

Accomplishment Description 

Several wing—body configurations representing contemporary 

wing designs and advanced wing concepts were used to exercise 

the Navier—Stokes codes. Comparisons of the numerical predictions with wind tunnel and flight-test data were made at cruise 

and near-buffet conditions. For a grid of about 300,000 points, a 

typical TLNS3D run requires about 2 Cray-2 hours and a typical 

CFL3D run requires 3-5 hours. The codes require less than 

21 megawords of memory.

Significance 

CFL3D and TLNS3D predict the transonic flow characteristics for 

a variety of wing—body arrangements. More specifically, the 

codes simulate the progression from a double-shock flow pattern 

at low lift levels up to the single-shock structure at transonic high 

lift. This flow phenomenon is typical of advanced wing designs. 

TLNS3D also provides good correlation of separation progression predictions with flight-test measured buffet progression 

using the Johnson—King nonequilibrium turbulence model. In 

particular, the critical spanwise location for the first appearance 

of separated flow at the trailing edge and the chordwise progression of separation closely agree with the observed characteristics. 

Future Plans 

Further comparison of TLNS3D and CFL3D predictions with test 

data will be made for more wing—body configurations. The 

multiblock capability of both CFL3D and TLNS3D will be 

exercised to study the flow for more complex configurations. 

Emphasis will be placed on the interference flow phenomenon 

for wing-mounted engine installations. 

Mo = 0.83

RNC50x 106 

a=4.5deg a=5.Odeg a=5.5deg 

CL = 0.77 CL = 0.81 C1 = 0.85

= 

flPredicted streamwlse velocity at 1/2 cell off wing surface 

TLNS3D-predicted flow-separation progression for a contemporary wing illustrating the area immersed in reversed flow.

25
Turbulence-Model Development for Impinging Jet Flows 

Robert E. Childs, Principal Investigator 

Co-Investigator: Laura C. Rodman 

Nielsen Engineering and Research, Inc. 

Research Objective 

To use large-eddy simulation to gain knowledge about the 

physics of impinging jet flows, and to develop an improved twoequation model for these flows based on this knowledge and 

experimental data. Complex flow fields are generated by the 

impinging jets associated with thrust reversers and vertical takeoff or landing jet aircraft. Accurate turbulence modeling is 

needed for these types of flows. However, existing models, even 

advanced models such as the Reynolds-stress transport models, 

generally perform poorly in complex impinging jet flows. 

Approach 

Large-eddy simulations were performed on an impinging jet 

inclined 45 degrees toward a cross flow with a velocity of 22% 

of the jet centerline velocity. A compressible Navier-Stokes 

solver with fourth order spatial accuracy and second order 

temporal accuracy was employed to perform the simulations. 

Turbulent inflow boundary conditions were generated in 

auxiliary simulations of spatially periodic turbulence. Both the 

dynamical and Reynolds-averaged simulation results were 

evaluated in the study. Turbulence model development was 

performed with a different version of the same flow solver. 

Modifications to the k-c model that could account for the 

observed physics were sought. 

Accomplishment Description 

The large-eddy simulation requires 20 megawords of memory, 

and takes 60 single-processor Cray Y-MP hours. The simulation 

revealed that vortex stretching is an important phenomenon in 

impinging jet flows. Modifications for the effects of streamline 

curvature and vortex stretching, consisting of additional source 

terms in the c-equation, were incorporated in the k-c model. 

The accuracy of the modified model has proven to be very good. 

The predicted location of the ground (horseshoe) vortex, which 

generates a region of low pressure, is a good test of the model's 

accuracy. The error in the ground vortex core location predicted 

with the modified model is reduced by 70% or more in comparison to calculations with the standard k-c model. The model has 

given good results for three different impinging jet configurations. Grid dependence in the Reynolds-averaged results was 

shown to be negligible.

Significance 

Previously, the errors in the predicted turbulence in impinging 

jet flows were large-50% or more. Since these flows significantly depend upon the turbulence, numerical prediction 

methods were unreliable for flows that involved impinging jet 

phenomena. Present results indicate that a great improvement in 

accuracy has been achieved and that this improvement may 

extend to a range of impinging jet flows. 

Future Plans 

To determine its range of applicability, the modified k-c model 

will be tested in a wider variety of impinging jet flow calculations. 

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-6 -4 -2

X 

Surface centerline pressure coefficient for a normal impinging jet 

in cross flow; experimental data and calculations with standard 

(STD), vortex stretching (VS), and curvature/vortex stretching 

(CVS) k-c models. 

26
Prediction of Turbine Endwall Heat Transfer 

Rodrick V. Chima, Principal Investigator 

NASA Lewis Research Center 

Research Objective 

A three-dimensional Navier-Stokes code was used to compute 

endwall heat transfer in a linear turbine cascade at two Reynolds 

numbers. The cascade was tested at NASA Lewis using a liquid 

crystal technique to map the endwall heat transfer. Two distinctly different heat transfer patterns were seen at low- and highReynolds numbers. The objective of the work was to verify the 

code by computing the two heat transfer patterns, and to use the 

computations to explain the physics behind the two patterns. 

Approach 

The Rotor Viscous Code three-dimensional turbomachinery 

analysis code was used. It is a three-dimensional thin-layer 

Navier-Stokes code with an algebraic turbulence model. A 

finite-difference formulation and a Runge-Kutta scheme with 

local time stepping and implicit residual smoothing were used. 

Accomplishment Description 

A new algebraic turbulence model for the heat transfer analyses 

was developed. The model is based on the Cebeci-Smith 

boundary-layer model, but uses unique integral relations to 

determine turbulent length and velocity scales. Much of the 

work was aimed at developing and verifying this model. 

Computations were made on the Cray Y-MP using a grid with 

about 300,000 points. Approximately 8 megawords of memory 

were required. The code is highly vectorized and autotasked, 

with a measured performance of 710 MFLOPS on a nondedicated system. The cases run here had low inlet Mach 

numbers (0.05-0.2) and required 2,500-4,000 iterations 

(2.5-4 CPU hours) to converge. The accompanying figure shows 

computed blade and endwall heat-transfer contours for a lowspeed case (Re = 78,000) expressed as a dimensionless Stanton 

number. The contours show the heat transfer augmentation 

around the leading edge of the blade and the increasing heat 

transfer through the passage. The high-speed results (not shown) 

exhibit a distinctly different heat transfer pattern. Both patterns 

agree fairly well with the experimental data. 

Significance 

Gas-turbine engine efficiency can be improved by increasing the 

turbine inlet temperature or by decreasing the amount of coolant 

air. To use either of these approaches, it is necessary to understand the physics of the heat-transfer process in turbine blade 

passages and to be able to predict the heat transfer accurately. 

The code and turbulence model support several turbine experiments at NASA Lewis, including the low-speed cascade, a 

transonic cascade, and a radial inflow turbine. It is also being 

used at NASA Lewis for analysis of fan blades and wave rotors, 

and is available to United States industries and universities.

Future Plans 

Blocked-grid capability is being added to the code. Initial 

applications will be for detailed analysis of the flow in the tipclearance region above the rotor blades. 

Publications 

1. Chima, R. V. "Viscous Three-Dimensional Calculations of 

Transonic Fan Performance." AGARD Conference Proceedings. Neuilly-Sur-Seine, France, Feb. 1992. 

2. Chima, R. V.; Giel, P. W.; and Boyle, R. J. "Algebraic 

Turbulence Model for Three-Dimensional Viscous Flows." 

To be presented at the AIAA Aerospace Sciences Meeting, 

Reno, NV, Jan. 1993. 

Linear turbine cascade: Stanton number x 1,000; Re = 78,000.

27
Compressible Taylor—Couette Flow 

Chuen-Yen Chow, Principal Investigator 

Co-Investigators: Meng-Sing Liou and Kai-Hsiung Kao 

University of Colorado, Boulder/NASA Lewis Research Center 

Research Objective 

To investigate and understand the mechanisms of the vortical 

structure of the compressible flow between two concentric 

cylinders. The transition behavior of the boundary layer on an 

axisymmetric body is given a spin motion. This may occur on a 

rocket or along the shaft of a turbomachine and may drastically 

affect the performance of the original design. 

Approach 

The governing equations are numerically solved employing the 

implicit approximate-factorization Beam-Warming algorithm. 

The scheme is formulated using Euler-implicit time-differencing 

and second-order finite-differencing approximations for all 

spatial derivatives. It is simulated using a fully vectorized, threedimensional Navier-Stokes solver, FDL3D1, developed by 

Miguel R. Visbal at Wright-Patterson Air Force Base. 

Accomplishment Description 

The study is to numerically investigate the compressible 

Taylor-Couette flow with a radius ratio R 1 /R2 = 0.5 and infinite 

length. The wavelength of a complete Taylor vortex above 

critical Reynolds number is first evaluated by utilizing a large 

computational domain for which simple zero-axial derivative 

boundary conditions are imposed at the two end plans. Using a 

finer grid system, computation is restarted to improve the 

solutions so that the pre-calculated wavelength is used as the 

integration domain and, therefore, the periodic boundary 

condition can be applied at the end plans. For the case of steady 

flow, convergence was assured by monitoring the dissipation 

rates (torques) on both inner and outer cylinders so that altering 

grid resolution produced no changes in flow topology and 

introduced only negligible variations between the inner and 

outer torques. The computational grids for the Taylor-Couette 

flow are clustered in the radial direction near the surfaces of the 

inner and outer cylinders and equally spaced in the axial and 

circumferential directions. Using stretching grids near the solid 

boundaries would allow a more accurate computation not only 

to extract energy from the moving wall, but also to resolve rapid 

variations of flow structures in the boundary layer. For a single 

wavelength, the grid resolution employed in the flow is 

62 x 35 x 31 in axial, azimuthal, and radial directions, respectively, at Mach 0.1 and a Reynolds number of 100. As Mach or 

the Reynolds number increases, the required grid size must 

increase. The computational requirement of the code was found 

to be approximately 35 words per grid and 2.0 x 10-5 CPU 

seconds per grid point for each iteration on the Cray Y-MP.

Significance 

The study concludes some preliminary investigations on the 

formations of the Taylor vortices. The compressible Taylor vortex 

flows appear to develop single period waves as the wavelength 

sustains a constant value. Analysis of flow variables strongly 

supports the concept of an evolving jet-like structure when the 

Mach number and Reynolds number are increased. 

Future Plans 

The evolution of the vortical structure at Reynolds numbers 

higher than the critical value will be studied and we will provide 

a detailed account of (1) vortex-loop development at various 

spike stages during transition and (2) the path the vortical flow 

takes that leads to a final chaotic state. Numerical methods will 

be used to study the control of transition with mechanisms such 

as wall heating or cooling, acoustic excitation, and blowing or 

suction at the body surfaces. 

Particle traces and contour plots of axial and radial velocity 

components. 

28
Bleed with shock incidence 

4acIt cw.,a,,, 

Airframe and Inlet Aerodynamics 

Wei J. Chyu, Principal Investigator 

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

NASA Ames Research Center/Cornell University/Carnegie Mellon University 

Research Objective Future Plans 

To develop an analytical capability to predict integrated Future efforts will focus on code validation cases for complex 

performance of forebody and inlet systems for highly maneuver- inlet geometries, including the integration of internal and 

able aircraft. forebody solutions with various inlet performance control 

techniques. 

Approach 

Adapt and develop Navier—Stokes codes with two numerical 

approaches, one with a finite difference method combined with 

a multizonal, two-grid topology and Chimera grid-embedding 

technique (F3D) to treat flow problems in complex geometries, 

and the other with a finite-volume method implemented with 

multigrid diagonal implicit schemes on block structured grids to 

allow flexibility in grid generation and to improve computational 

efficiency. 

Accomplishment Description 

The F3D Navier—Stokes code was used to study the inlet 

boundary-layer control by simulating a transonic flow through a 

circular bleed hole on a flat plate with a plenum. The results of 

this computation reveal inlet-bleed/shock-wave/boundary-layer 

interactions that provide valuable information to ongoing inlet 

experiments. The development of a finite-volume Navier—Stokes 

code resulted in a unique parallel-block multigrid solution with 

appreciably improved convergence. 

Significance 

The increased maneuvering requirements of advanced aircraft 

dictate more complicated inlet geometries and operations. 

Advanced computational techniques permit analysis of more 

complex inlets, making possible the investigation of a broader 

range of design variables associated with integration of the 

forebody and inlet systems and the inlet performance with 

boundary-layer control.

Shock-wave/bleed interaction

M.., 15 

*

iIIUAF ggq .u4 ^ljtd We 

Simulations of a transonic flow through a circular bleed hole on a flat plate with a plenum.

29
Generic National Aero-Space Plane Forebody—Inlet Integration 

Charles E. Cockrell, Jr., Principal Investigator 

Co-Investigator: Lawrence D. Huebner 

NASA Langley Research Center 

Research Objective 

To study the inlet module for a generic hypersonic 

forebody—inlet model and determine the feasibility of obtaining 

internal inlet drag force predictions based on computational 

fluid dynamics (CFD) solutions simulating exhaust flows of 

body-mounted scramjet propulsion systems. 

Approach 

Three-dimensional parabolized N4avier—Stokes (PNS) solutions 

were obtained for the inlet flow field using the General Aerodynamic Simulation Program. Data from CFD solutions were 

compared to experimental pressure data from a wind tunnel test 

and were used to obtain predictions for the internal drag force of 

the inlet module. 

Accomplishment Description 

Computational solutions were obtained for the inlet flow field for 

free-stream conditions of M,,,, = 6.0 and Re,, = 2.0 x 106 per foot, 

and are identical to those in previous wind tunnel tests of a 

generic forebody—inlet model. The three-dimensional PNS 

solutions were compared with experimental Pitot pressure data 

at the inlet exit plane. Analysis of the CFD solutions generally 

showed good agreement with exit-plane Pitot pressure data. 

Some discrepancies exist due to an inaccurate prescription of the 

inflow boundary conditions at the inlet entrance plane. The top 

figure shows a comparison of CFD Pitot pressure values at the 

exit plane with the vertical surveys of Pitot pressures taken in 

wind tunnel tests superimposed. The bottom figure shows a side 

view of the inlet (flow is from left to right), near the inlet centerline, with computational Pitot pressure values. This view shows 

qualitative features of the inlet flow field such as the forebody 

boundary layer swallowed by the inlet on the upper surface and 

the formation of a shock due to boundary-layer growth on the 

lower surface. The three-dimensional momentum equation was 

used to obtain an expression for the internal inlet drag force at 

the entrance and exit planes of the inlet. The CFD solution 

pressure data were used to evaluate this force. A typical PNS 

solution required approximately 38 megawords of memory and 

2.9 Cray-2 hours.

Significance 

This analysis provides a method for obtaining the internal inlet 

drag force of the inlet module under investigation. An accurate 

value of this force is needed to determine the forces and 

moments of individual components of wind tunnel models used 

in powered testing. The internal inlet drag force is difficult to 

resolve using experimental methods, so an accurate computational procedure is needed. 

Future Plans 

Differences exist in comparing the CFD solutions with experimental Pitot pressure values at the inlet exit plane. A CFD 

solution of the forebody flow field may be required to initialize 

the inlet solution. Future work will focus on obtaining solutions 

of the forebody flow field to initialize inlet flow-field solutions. 

Publication 

Cockrell, C. E. and Huebner, L. D. "Generic Hypersonic Inlet 

Module Analysis." AIAA Paper 91-3209, 9th AIAA Applied 

Aerodynamics Conference, Baltimore, MD, Sept. 1990. 

Exit plane (top) and inlet center line (bottom) Pitot pressure 

values from a three-dimensional parabolized Navier—Stokes 

solution of a hypersonic inlet module. The experimental Pitot 

pressure values are superimposed on the exit plane values and 

are shown as vertical and horizontal bars. The Pitot pressure 

values range from 0.0 (dark blue) to 8.0 (dark red) pounds per 

square inch. 

30
Three-Dimensional Afterbody Flow with Jet Exhaust 

William B. Compton, Ill, Principal Investigator 

Co-Investigator: Khaled S. Abdol-Hamid 

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

Research Objective 

To numerically investigate the propulsion integration effects on 

a fighter airplane at subsonic, transonic, and supersonic speeds. 

An intermediate goal is to develop and validate a threedimensional Navier—Stokes numerical method to investigate 

the problem. Steps include obtaining accurate solutions for 

afterbodies with attached and massively separated flows, 

including the internal nozzle flow and jet exhaust in the 

calculations, and including tails and multiple jets. 

Approach 

The multiple-block versions of the CFL3D and PAB3D threedimensional Navier—Stokes codes are used as a basis for thi 

method. The codes employ the finite-volume principle and 

feature upwind-biased flux-difference splitting combined with 

gradient-limiting procedure to ensure monotonicity across 

discontinuities. Areas of research include the evaluation of both 

algebraic and two-equation turbulence models for predicting 

attached and massively separated flows over afterbodies typical 

of those for advanced fighters. For validating the numerical 

techniques, the computational effort is correlated with highly 

detailed experiments in the Langley 1 6-Foot Transonic Tunnel, 

the Basic Aerodynamic Research Tunnel, and the Jet Noise 

Laboratory. 

Accomplishment Description 

Simulations have been made for the external flow, the jet 

exhaust plume, and the internal nozzle flow at free-stream Mach 

numbers from 0.80-1.20. Jet total pressure ratios ranged from 

1.5-9.8 (from below to above the design value). The computations, made with algebraic turbulence models, successfully 

predicted the qualitative jet-exhaust effects on the afterbody 

flow. However, when massive shock-induced separation on the 

afterbody exists, the ability to predict the shock locations and 

pressure rises downstream of the shocks vary from good to fair. A 

typical solution takes approximately 20 megawords of memory 

and 40 Cray Y-MP hours. In addition to these investigations, a 

two-equation turbulence model has been added to the code 

PA133D and tested in the jet plume. 

Significance 

In a high performance fighter, 40-50% of the drag is associated 

with its afterbody. In this region, flows over the fuselage and 

empennage merge and further interact with the propulsive 

exhaust to create a flow field that is complex, rotational, and 

strongly interactive. A numerical procedure to analyze this flow 

for advanced configurations would be useful in designing new 

airplanes.

Future Plans 

The two-equation model will be tested for wall-bounded shear 

flows on the afterbody. In addition, the computations will be 

compared to detailed laser Doppler velocimeter, pressure, and 

hot-film data taken inside the nozzle and in the plume. 

Velocity vectors colored by Mach number. Vertical plane of 

symmetry: M,,, = 1.20, nozzle pressure ratio = 9.8, R = 20.5 x

31
Analyses of F/A-i 8E/F Upgrades 

Raymond R. Cosner, Principal Investigator 

Co-Investigators: F. Creasman, R. S. Dyer, T. D. Gatzke, J. A. Johnson, P. J. Malloy, W. W. Romer, and P. G. Wilihite 

McDonnell Aircraft Company 

Research Objective 

Computational fluid dynamics (CFD) analysis methods are being 

developed to complement wind tunnel testing in two key areas. 

First, CFD codes combined with NASA structural analysis 

models are being developed to compute aeroelastic load 

increments to wind tunnel data. This will allow engineers to 

compensate for the flexibility of flight vehicles, compared with 

rigid wind tunnel models. These aeroelastic modifications to 

flight loads can be significant in setting structural design 

requirements. The second area of investigation is to develop 

methods to correct wind tunnel data for support-system interference. Test data are modified by the presence of a support sting 

and by distortion of the aerodynamic geometry for physical 

mounting of the model on the sting. These "sting and distortion" 

increments can have a significant impact on estimates of flight 

performance. 

Approach 

Multiblock grids were generated for 10 near-complete geometries from the F/A-i 8 family. All grids included the fuselage, 

strake, wings, and tails. Some also included flowing inlets and 

nozzles, a wingtip missile, a wind tunnel support system, and/or 

wind tunnel walls. Subsonic and transonic solutions were 

executed to develop engineering procedures for obtaining the 

best possible accuracy in predicting aeroelastic flight loads and

sting/distortion increments. Accuracy was assessed using wind 

tunnel and flight data from the YF-i 7, F/A-i 8A, and F/A-i 8C 

aircraft. Predictions were executed for the F/A-i 8E aircraft in 

advance of confirming wind tunnel tests. 

Accomplishment Description 

Accuracy of 2-3% was demonstrated for wing root bending 

moment at Mach 0.9. The flow over the wing at this condition is 

dominated by viscous interactions, and is substantially modified 

by elastic deformations of the wing in flight. Accurate sting and 

distortion corrections were obtained at Mach 0.85 for F/A-i 8 

models tested in the polysonic wind tunnel. 

Significance 

Both aeroelastic increments and sting/distortion increments 

demonstrated new CFD capabilities that complement wind 

tunnel testing. Each showed that CFD analysis can increase the 

value of wind tunnel data by correcting for limitations that 

otherwise cannot be assessed adequately prior to flight testing. 

Future Plans 

The elastic loads prediction capabilities will be tested and 

improved for additional limiting conditions, including asymmetric states. The sting/distortion correction methodology will be 

extended across the transonic range, from Mach 0.6-1.2. 

Sting-mounted F/A- 18C with a distorted aft end. M = 0.85, a = 0.0. 

32
Endwall and Casing Treatment Flow in a Transonic Fan Rotor 

Andrew). Crook, Principal Investigator 

General Motors Corporation, Allison Gas Turbine Division 

Research Objective 

To determine the stall inhibiting features or mechanisms of 

casing treatment (particularly circumferential grooves) and to 

identify the important geometric parameters of the treatment. An 

important preliminary objective is to understand qualitatively the 

roll of the passage shock and the tip-clearance flow in a stalling 

or unstarting transonic rotor since this is the undesirable flow 

condition that casing treatment often delays. 

Approach 

Three-dimensional Euler and Navier—Stokes solutions are run to 

a steady state for the rotor geometry. Exit pressure is raised and 

more steady solutions are found to simulate the throttling of the 

rotor. 

Accomplishment Description 

Euler and Navier—Stokes solutions were obtained for the flow 

field about a turbomachinery fan section consisting of a transonic rotor, two downstream ducts, and guide vanes in those 

ducts. The solutions were generated with a finite-volume explicit 

procedure that contained an average passage formulation for 

neighboring blade-row effects. The accompanying figure shows 

a Navier—Stokes solution about the rotor of the GMA 3007 fan 

section. Fluid streamlines originating in the tip-clearance region 

are seen as black lines. The streamlines show the influence of 

the clearance flow and the trajectory of the vortex in the 

passage. Contours of relative total pressure are shown on a 

plane just downstream from the rotor trailing edge. 

Significance 

Casing treatment can improve surge margin in fans and will be 

implemented in the fan development of the integrated highperformance turbine-engine technology program. While the 

casing treatment's potential for performance improvement is well 

known, a complete explanation of the stall suppressing mechanism has not been established. Analysis of fan-rotor stalling or 

unstarting provides a better understanding of the general endwall 

flow problem.

Future Plans 

Circumferential grooves on the case over the rotor will be 

modeled. Performance of the treated geometry will be simulated 

up to the stall or numerical divergence point and solutions will 

be compared and analyzed. 

Navier—Stokes solution about the rotor of the GMA 3007 fan 

section.

1i 

33
Fighter Acoustic-Load Predictions 

Richard D. Crouse, Principal Investigator 

Northrop Corporation 

Research Objective 

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 is calculated using 

existing three-dimensional Navier—Stokes flow codes. The timevarying pressure field is transformed with a Fourier transform to a 

frequency domain. This allows the identification of the frequencies and the amplitude of the varying pressure. 

Accomplishment Description 

A three-dimensional grid consisting of 899,475 grid points was 

generated within and above a cavity and flat plate. The cavity 

model matched the weapon's internal carriage and separation 

program test fixture with a length/depth of 4.5 and a length/width 

of 4.5. The Mach number studied was 0.6. ARC3D was run in a 

time-accurate mode for 232 milliseconds (23,200 iterations). The 

pressure versus time was saved for various wall locations along 

with full-flow solutions at every 1,000 iterations. The timevarying pressure was converted with a fast Fourier transform to 

sound pressure level versus frequency. These three-dimensional 

results will be compared to previous two-dimensional computations and to experimental results. The accompanying figure 

shows the pressure versus time at one location within the cavity 

after the startup transients have dissipated. This three-dimensional 

computation required 400 Cray-2 hours and 32 megawords of 

memory. 

Significance 

Acoustic load prediction is needed when considering the 

increased performance levels of advanced tactical air vehicles. 

This study will provide an extremely useful tool to aid in the 

design of new vehicles. The computational means will enable 

solutions to be calculated on a variety of topologies and over a 

vast flight regime. It will not present case-dependent empirical 

results, nor simply recalculate old data. Rather, it will have 

applicability 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 methodology will continue to be applied to three-dimensional 

internal weapon cavity flow. This will include analysis of 

innovative designs to reduce the aeroacoustic loading within the 

cavity. The effect of boundary layer control, spoilers, and sloped 

bulkheads will be considered. The methodology for predicting 

aeroacoustic pressure fluctuations in three-dimensional cavities 

will be applied to various geometries to show the effect on the 

forces expected within the cavity. 

40 80 120 160 200 

Time (ms) 

Pressure versus time at one location within the cavity after the 

startup transients have dissipated. 

.33 

. .17 

w U 

C 

0 

0 = .01 

0 

C) 

Q.

—.31 

34
Flexible Aerobrake Aerothermodynamic Study 

M. I. Cruz, Principal Investigator 

Co-Investigators: D. E. Ressler, T. P. Shivananda, and E. F. Zabrensky 

TRW, Federal Systems Division/TRW, Ballistic Missiles Division 

Research Objective 

To validate and utilize computation fluid dynamic simulations 

to determine the aerodynamic and thermodynamic environments 

an aerobraking vehicle would experience during approach and 

landing on the Martian surface. 

Approach 

Perfect gas, equilibrium, and nonequi I ibrium Navier—Stokes 

solvers combined with current grid generation methods are used 

in vehicle parametric studies. The sensitivity of drag coefficient 

and surface heat transfer rates to variations in flexible and rigid 

aerobrake geometry was derived. 

Accomplishment Description 

Two axisymmetric Navier—Stokes solvers developed at NASA 

Ames have been used to obtain flow field solutions for candidate 

aerobrake vehicles. The solvers incorporate both equilibrium 

and nonequilibrium Martian chemistry (CO2—N 2). The nonequilibrium thermochemical solver incorporates surface catalysis 

enhancing the predicted heat transfer. Results from the simulations have been compared with available ground-test data for 

similar configurations and agree quite well. The initial conditions for the simulations have been obtained from 7 km/sec 

direct entry trajectories of the Mars Environmental Survey 

mission. The results of the simulations have shown the vehicle 

drag is maximized with a 70 degree cone angle. The vehicle 

thermal loads are manageable using available rigid heat-shield 

materials with the nonequilibrium thermochemical results 

displaying 30% higher surface heat flux than the equilibrium 

results. Each solution required approximately 10 Cray-2 hours. 

Significance 

The current studies have shown that a 7 km/sec direct entry 

using aerobraking on the approach and descent to the Martian 

surface is attainable with current technology. Reliance on 

current fluid simulation is necessary for the design and analysis 

of the aerobraking vehicles. 

Future Plans 

Analysis will continue using the codes for flow-field simulations 

at other points on the trajectories, including perturbed trajectories. 

Publications 

1. Ressler, D. E.; Shivananda, T. P.; Zabrensky, E. F.; and Cruz, 

M. I. "Drag and Heat Transfer Calculations about a MESUR 

Aerobrake Vehicle." AIAA Paper 92-2684, AIAA 10th 

Applied Aerodynamics Conference, Palo Alto, CA, June 1992.

2. Post, D. L. and Cruz, M. I. "CFD Application to Flexible 

Aerobrake Design for Mars Entry." AAS Paper 91-424, AAS/ 

AIAA Astrodynamics Specialist Conference, Aug. 1991. 

(a)

(b)

Mars Environmental Survey flow-field Mach number contours 

(a), and fully catalytic wall heat flux (b) are shown at = 31.0, 

Re= 1.1051meter, a=0.0.
Particle-Gas Dynamics in the Protoplanetary Nebula 

Jeffrey N. Cuzzi, Principal Investigator 

Co-Investigators: Joelle Champney and Anthony Dobrovolskis 

NASA Ames Research Center/A.T.M., Inc./University of California, Santa Cruz 

Research Objective 

To model the primordial disk-shaped nebula from which the 

planets formed. The stage studied is when a particulate phase 

has condensed out of the cooling gas and the particles have 

grown to sufficient size to settle into a layer near the midplane 

of the nebula. Here, the particle-mass density becomes much 

greater than that of the local gas. Because the gas is pressure 

supported and the particles are not, a small, but significant, 

orbital velocity difference results between the gas and particle 

phases. The resulting wind shears and turbulence in the gas 

phase cause both advective and turbulent transport of the 

particles and gas relative to their unperturbed states. It was 

during this phase that the primitive meteorites formed. 

Approach 

We developed a fully viscous, compressible, two-phase flow 

model of the nebula that solves the gas and particle conservation 

and momentum equations in the rotating system using a perturbation technique. The turbulent boundary layer surrounding the 

particles is modeled with a Prandtl model and a k-c model. We 

developed a model for the particle Schmidt number to couple 

our mean flows and turbulence into upward diffusion of the 

particle phase, which balances their downward settling tendency. Thus, the dynamics of the gas and particle phases are 

fully coupled in a system with significant feedback.

Accomplishment Description 

Using a fully Reynolds-averaged formulation, with correlation 

terms modeled by a generalized gradient-diffusion hypothesis, 

we have have determined the mean gas and particle flows and 

the vertical diffusion of particles at a variety of locations in the 

protoplanetary nebula and in a possible circumplanetary nebula. 

At each location we have run particle size cases, including 

particles of between 1-1,000 cm radius with internal density of 

0.1-1.0. A typical run requires 1 Cray-2 hour and 1 megaword 

of memory to reach a non-oscillatory state for each set of model 

parameters. 

Significance 

Particles must grow to at least 1,000 kg mass before being able 

to accumulate under their own gravity. This contradicts a 

widespread belief in a 'gravitational instability" that would lead 

immediately to the formation of planetesimals. Our result implies 

that growing particles probably undergo significant spatial, 

physical, and chemical evolution as they grow. 

Future Plans 

We will extend the code to two dimensions and to cover a range 

of particle sizes simultaneously. We will also study the threedimensional clumping of particles in turbulence. 

Solutions for the mean radial, axial, and azimuthal velocities of gas (smooth colors) and particles (dots), as a (unction of vertical 

distance 0-20,000 km from the nebula midplane—a very small scale compared to the thickness of the nebula gas disk at this location 

(about 10 million km). Velocities (cm/sec) are referenced to an observer moving with the unperturbed gas and particles respectively; 

(a) looking inward, (b) looking backward (away from direction of orbital velocity). The spatial density of the dots is proportional to the 

logarithm of particle density. A layer is concentrated around the nebula midplane. Within this layer the gas is driven to a higher 

rotation rate than outside the layer. Outside the layer, the particulates drift inward toward the forming sun, while near its upper 

boundary the gas is wafted gently outward across the face of the particle layer. Near the midplane there is a slow outward gas flow 

and a slow inward particle drift. Image created by Robert Hogan of Synernet, Inc. 

36
Dynamical Modeling of the Solar Atmosphere 

Russell B. Dahlburg, Principal Investigator 

Co-Investigators: S. K. Antiochos, Jill P. Dahlburg, and J. T. Karpen 

Naval Research Laboratory 

Research Objective 

To investigate turbulent magnetohydrodynamic (MHD) processes 

in the solar atmosphere. 

Approach 

Direct numerical simulation, using spectral methods, of the 

dissipative compressible and incompressible MHD equations in 

two- and three-dimensional geometries. 

Accomplishment Description 

We used our spectral-method codes to investigate transition to 

MHD turbulence in three-dimensional magnetic reconnection 

via secondary instability (a typical run requires 8 megawords 

of memory and 10 Cray-2 hours). The accompanying figures 

illustrate the evolution of the secondary mode. Contours 

enclosing the magnetic neutral lines in the plane of the threedimensional electric-current sheet are shown. Two periods 

of the mode are shown in x, with a magnetic 0-line in the 

center.

Contour levels 0.000 Mach 

0.00000 0.00 dog Alpha 

0.00200 50.00 Re 

0.00400 4400 Time 

0.00600 33x33x65 Grid 

0.00800 

0.01000

Significance 

Magnetic reconnection is an important energy release mechanism 

for solar activities such as flares and coronal heating. Classical 

analyses generally show that magnetic energy is transformed into 

heat and kinetic energy on the relatively slow dissipative time 

scale. Our work has identified a process that enables the magnetic 

field to release energy on the ideal time scale. 

Future Plans 

To extend the present study to the case of electric-current sheets 

and study the response of finite-length magnetic arcades to 

photospheric footpoint shearing. 

Publications 

1. Dahlburg, R. B.; Antiochos, S. K.; and Zang, T. A. "Dynamics 

of Solar Coronal Magnetic Fields." Astrophys. J. 383 (1991): 420. 

2. Dahlburg, R. B.; Antiochos, S. K.; and Zang, T. A. "Secondary Instability in 3D Magnetic Reconnection." NRL Memo-- 

randum Rept. 6966, 1992. 

Contour levels 0.000 Mach 

0.00000 0.00 dog Alpha 

0.00200 50.00 Re 

0.00400 1.33x102 lime 

0.00600 33x33x65 Grid 

0.00800 

0.01000

(a) (b) 

The evolution of the secondary mode (a) at an early time and (b) during impulse phase.

37 
Simulation of Turbulent Jets for Aeroacoustic Applications 

Sanford M. Dash, Principal Investigator 

Co-Investigators: Neeraj Sinha, Brian J. York, and Robert A. Lee 

Science Applications International Corporation 

Research Objective 

To develop a computational methodology for simulating the 

structure of turbulent, imperfectly expanded jets. This research 

is in support of the High-Speed Civil Transport program and 

focuses on the interpretation of jet data and jet noise suppression 

concepts. 

Approach 

Initial work has focused on the utilization of Reynolds-averaged 

Navier—Stokes (RNS) methodology with advanced turbulence 

models to simulate the mean flow structure of axisymmetric jets 

with shocks, jets with plug centerbodies, nonaxisymmetric 

(rectangular/elliptical), and multiple jets. The approach has 

involved the exploration of new turbulence models 

(compressible-dissipation model with pressure dilatation 

terms) to deal with high Mach number compressibility effects 

in RNS codes such as PARCH and CRAFT. 

Accomplishment Description 

The compressible-dissipation models of Sarkar and Zeman were 

reduced to k-c form and assessed using a high-speed shear-layer 

data base. A new version of k-c was formulated (k-c CD) unifying 

the models by adding pressure dilatation terms and additional 

terms to enforce agreement with data at high Mach numbers. 

The new model has improved behavior over that of k-c for highspeed jets and shear layers. Each run requires 10-40 megawords 

of memory and 10-20 Cray-2 hours depending on the grid size. 

Significance 

The ability to simulate high-speed jet structure has been greatly 

enhanced by this research. A new turbulence model is under 

development and is extremely promising. A building-block 

approach to extend capabilities to more complex flows is being 

guided by newly obtained data.

Future Plans 

The jet research codes will be applied to flows of increasing 

complexity. Unsteady jet flows will be investigated using largeeddy simulation methodology. Support to infrared and acousticsoriented jet simulation will occur. 

—4 —2 0 2 4

X 

Improvements provided by the new k-c CD model over the k-c 

turbulence model. 

CL 2 it 

38
Three-Dimensional High-Speed Plume/Propulsive Flow-Field Analysis 

Sanford M. Dash, Principal Investigator 

Co-Investigators: Neeraj Sinha, Brian J. York, Robert A. Lee, Ashvin Hosagadi, and Donald C. Kenzakowski 

Science Applications International Corporation 

Research Objective 

To establish a computational methodology for the simulation of 

steady and transient chemically reacting, multiphase, high-speed 

plume/propulsive flow fields. Advanced turbulence models are 

utilized and new techniques for simulating nonequilibrium 

thermochemical and multiphase processes are being developed. 

Approach 

The PARCH time-asymptotic Navier—Stokes code using matrixsplit/loosely coupled thermochemistry and the CRAFT timeaccurate Navier—Stokes code utilizing large-matrix/strongly 

coupled thermochemistry are implemented. Specialized versions 

have been developed for different applications. Complex threedimensional configurations have been analyzed simulating 

scramjet propulsive flow fields, rocket propulsive flow fields, 

and missile—plume interactions (both steady and transient). 

Accomplishment Description 

PARCH was utilized to simulate missile flow fields from nose to 

plume farfield for single- and multi-engine boosters with 

afterburning chemistry and multiphase flow: A new upwind/ 

implicit finite-volume particle solver was developed and 

incorporated into CRAFT for simulation of transient plume 

effects. The particle solution included variable particle properties, different particle types (AL203 and Zr02), and phase 

change. Detailed studies of transient lateral multiphase plumes 

have been performed. New compressibility upgraded turbulence 

models have been incorporated into these codes. Each run 

requires 20-100 megawords of memory and 10-50 Cray-2 

hours depending on the grid dimensions and number of chemical species included. 

Significance 

PARCH and CRAFT are the only codes to analyze threedimensional combusting multiphase flows using advanced 

implicit numerics and improved two-equation turbulence 

models with compressible-dissipation and pressure-dilation 

terms. 

Future Plans 

Further work on multiphase flow will include upgrading the 

Roe/total-variation-diminishing schemes methodology to include 

liquid equations of state. Applications to hypervelocity launcher 

problems with dynamic adaptive gridding will be performed.

Publications 

1. York, J. J.; Sinha, N.; Kenzakowski, D. C.; and Dash, S. M. 

"PARCH Code Simulation of Tactical Missile Plumes/ 

Airframe/Launcher Interactions." CPIA Pub. 568, 19th 

JAN NAF Exhaust Plume Technology Meeting, May 1991. 

2. Hosangadi, A.; Sinha, N.; and Dash, S. M. "Solution of TwoPhase Diluted Gas Particle Mixtures Using Upwind FiniteVolume Methodology." AIAA Paper 92-0244, AIA.A 30th 

Aerospace Sciences Meeting, Reno, NV, Jan. 1992. 

3. Sinha, N.; Dash, S. M.; and Hosangadi, A. "Applications of 

an Implicit, Upwind NS Code, CRAFT, to Steady/Unsteady 

Reacting, Multi-Phase Jet/Plume Flow Fields." AIAA 

Paper 92-0837, AIAA 30th Aerospace Sciences Meeting, 

Reno, NV, Jan. 1992. 

Gas-phase and particle-cloud temperature contours at several 

times during jet injection.

39
Hydrodynamic Performance Evaluation 

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 cross-flow separation over unappended bodies of 

revolution at nonzero yaw angles using Reynolds-averaged 

Navier—Stokes (RANS) techniques. 

Approach 

A three-dimensional steady-state incompressible RANS solver 

based on the method of pseudo-compressibility, and utilizing 

centered finite differences together with an alternating-direction 

implicit/approximate factorization algorithm, was used to 

perform the first set of computations. An algebraic Boussinesq 

turbulence model (Briley—MacDonald—Fish) was used. The flow 

domain was modeled using a multiblock hybrid grid topology 

that eliminated the appearance of pole-type boundaries. 

Computations at the lower Reynolds numbers are to be compared with available experimental data. Computations at the 

higher Reynolds numbers are to be compared with predictions 

from semi-empirical and singularity based techniques in an effort 

to validate and enhance these less costly methods. 

Accomplishment Description 

The low-Reynolds-number laminar flow around a sphere was 

used to test the hybrid grid topology. Turbulent-flow computations for two Reynolds numbers (4,900,000 and 13,000,000) 

were completed for a lift/drag of 9.5 body of revolution at an

angle of yaw of 15 degrees. These computations were performed 

on an eight block computational grid (totaling 814,200 grid 

points) which modeled half of the symmetric physical domain. 

A typical solution required 5.5 Cray Y-MP hours and approximately 7 megawords of memory. Comparisons with experimental results have shown good correlation with body-separation 

patterns and with the size and extent of the leeside cross-flow 

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 of geometry forces and moments are required. 

Full-scale predictions of complex flow structures will help to 

improve the understanding of high-Reynolds-number incompressible flows, contributing to advanced design concepts, 

improved performance, and increased operating envelopes. 

Future Plans 

The analyses for two additional Reynolds numbers (1.0 x 108 

and 1.1 x 10) will be completed. Future objectives are focused 

on extending this effort by including more sophisticated turbulence models and by investigating the viability of and techniques 

for including the effects of laminar-turbulent transition for the 

model-scale Reynolds number calculations. 

Reynolds-averaged Navier—Stokes predictions for lift/drag of 9.5, body at 15 degrees yaw angle, and Reynolds number 1.3 x 107. 

40
Turbine "Hot Spot" Alleviation using Film Cooling 

Roger L. Davis, Principal Investigator 

Co-Investigator: Daniel J. Dorney 

United Technologies Research Center 

Research Objective 

To simulate the migration of a three-dimensional combustor 

hot streak in the flow of an axial turbine stage and to predict its 

impact, as well as that of rotor surface film cooling and heat 

transfer, on the rotor time-averaged surface temperature 

distribution. 

Approach 

An extended version of the unsteady three-dimensional 

Navier—Stokes code, ROTOR3, is used to simulate threedimensional viscous flows in axial turbomachinery. This 

approach is a third-order accurate, upwind, approximately 

factored, implicit finite-difference procedure. Boundary condition modifications have been made to allow for film cooling 

and heat transfer. 

Accomplishment Description 

A three-dimensional unsteady simulation of hot-streak migration 

in a turbine stage with turbine rotor-surface heat transfer and film 

cooling has been completed. The film cooling parameters were 

chosen to be representative of those used in modern gas turbine 

engines. The placement of the film cooling holes was determined by performing numerous two- and three-dimensional 

simulations. The addition of film cooling effectively alleviates

the rotor pressure surface temperature increase associated with 

the combustor hot streak. An average run in this study required 

50 Cray Y-MP hours and 16 megawords of memory for one 

rotor-blade-passing cycle. 

Significance 

The research resulted in the development of a numerical 

procedure that can accurately predict the impact of combustor 

hot streaks on turbine-blade heat transfer and the film-cooling 

requirements necessary to keep surface temperatures below 

critical levels, thus leading to turbine blade designs with 

increased durability. 

Future Plans 

The three-dimensional Navier—Stokes analysis will be used to 

optimize the film-cooling distribution for a high-pressure turbine 

blade representative of those used in advanced military aircraft 

engines. 

Publication 

Dorney, D. J. and Davis, R. L. "Numerical Simulation of Turbine 

'Hot Spot' Alleviation using Film Cooling." AIAA Paper 92-3309, 

28th AIAAJASME/ASEE Joint Propulsion Conference, Nashville, 

TN, July 1992. 

Trailing edge Trailing edge 

Hub Hui

Film cooling ports 

Leading edge Leading edge 

Rotor pressure surface time-averaged temperature contours (a) without film cooling and (b) with film cooling.

41
Transition over a Rough Surface 

Russell G. DeAnna, Principal Investigator 

Co-Investigator: Eli Reshotko 

U.S. Army Propulsion Directorate, AVSCOM/Case Western Reserve University 

Research Objective 

To predict surface stress and heat transfer coefficients and 

estimate transition location in order to create improved turbine 

airfoil designs. Turbine blade surfaces become rough after 

extensive operation due to high temperature and combustion 

products. This increases skin friction and heat transfer and 

promotes transition between laminar and turbulent boundarylayer flow. Delaying or estimating transition is important since 

we expect higher friction and heat transfer in transitional and 

turbulent boundary layers. 

Approach 

To gain an understanding of the complex features associated 

with transition over a rough surface by performing a direct 

numerical simulation using a spectral-element method. A 

smooth wall is distorted in a random manner and periodic 

boundary conditions are applied in the horizontal directions, 

effectively simulating an infinite boundary layer in the 

streamwise and spanwise directions. Thus, the boundary layer 

does not grow downstream and parallel flow is assumed. 

Accomplishment Description 

At time equal to zero, a Blasius velocity profile is applied to the 

entire domain. Since the periodic boundary conditions preclude 

a pressure gradient, the flow is sustained by a body force 

designed to yield a Blasius profile in the absence of surface 

roughness. Hence, any vertical or spanwise motion is due to the 

roughness. The streamwise vorticity indicated in the accompanying figure is due to the rough surface since a smooth boundary 

layer contains only spanwise vorticity. The code requires 

6 megawords of memory when a 9 x 5 x 5 mesh is used with 

each element. The time required depends on the information 

desired. In unsteady flows where turbulent quantities such as 

Reynolds stresses are required, the code must run for tens of 

"convection" cycles for a particle to cross the domain. One 

cycle requires approximately 100 Cray Y-MP hours. 

Significance 

This random-roughness calculation represents a first step toward 

understanding the transitioning boundary layer over a random 

surface shape. Hopefully, these ideas will be extended by other 

researchers so that we will gain a better understanding of the 

complex mechanisms associated with transitional, and ultimately 

turbulent, flow over surface roughness.

Future Plans 

A grid-resolution study will be done to ensure that numerical 

turbulence is not being generated. Further data will be obtained 

so that unsteady frequency response curves and turbulent 

quantities may be accurately obtained. 

Publication 

DeAnna, R. G. and Reshotko, E. "Laminar Flows over Distributed 

Surface Roughness." Forth International Symposium on Computational Fluid Dynamics, Davis, CA, Sept. 1991. 

2L • 

(a) Grid for random surface roughness. (b) Instantaneous 

streamwise vorticity contours on a horizontal plane at a 

displacement-thickness Reynolds number of 150 or a 

roughness height-based Reynolds number of 30. 

42
Flow about Almost Complete Aircraft and Hypersonic Configurations 

Jerry E. Deese, Principal Investigator 

Co-Investigators: Ramesh K. Agarwal, Thomas P. Gielda, Jerry G. Johnson, and Mark Axe 

McDonnell Douglas Research Laboratories/NASA Ames Research Center 

Research Objective 

To develop computational methods capable of predicting the 

flow field over a complete aircraft at subsonic to hypersonic 

speeds, including chemical kinetics effects. 

Approach 

The Reynolds-averaged Navier—Stokes equations are solved 

on body-conforming curvilinear grids using a finite-volume 

Runge—Kutta time-stepping scheme. Thin-layer, slender-layer, 

and full Navier—Stokes options are available. A multiblock zonal 

implementation allows complex configurations to be easily 

modeled. Algebraic, k-c, and Johnson—King turbulence models 

are available. Equilibrium air, hydrogen—oxygen, and hydrocarbon chemical kinetics models are included. 

Accomplishment Description 

Navier—Stokes solutions have been computed for wing—body and 

wing—body—winglet configurations. Inviscid calculations for a 

model in a wind tunnel have been performed with porous and 

solid wall-boundary conditions and compared with experimental 

data and free-flight predictions. The flow fields over typical 

launch vehicle configurations have been studied with 

Baldwin—Lomax, k-c, and Johnson—King turbulence models to 

determine the accuracy of these models for flows including 

separation. Inviscid solutions for a launch vehicle with nine 

boosters have been computed; the pressure distribution over this 

configuration is shown in the accompanying figure. Predictions 

for a Delta/Thor launch vehicle with plumes, a complete singlestage rocket configuration, and the Aeroassist Flight Experiment 

vehicle required analysis using the nonequilibrium chemistry 

capabilities of the code. Calculations for the more complex 

configurations required as much as 64 megawords of memory 

and 10 CPU hours. 

Significance 

Development of a multizone nonequilibrium chemistry 

Navier—Stokes code allows the prediction of the flow about 

complete configurations. Component integration effects can be 

analyzed, thus enabling configuration designers to optimize the 

aerodynamic configuration of complete vehicles. 

Future Plans 

Multigrid capability is being incorporated into the code. The 

resulting code will be capable of rapid and efficient analysis of 

complex configurations.

Surface pressure distribution on a launch vehicle with nine 

boosters.

43
Sensitivity of Turbulence to Initial Conditions 

Robert G. Deissler, Principal Investigator 

Co-Investigator: Frank B. Molls 

NASA Lewis Research Center 

Research Objective 

To determine whether sensitive dependence on initial conditions 

(chaos) is a characteristic of decaying turbulence rather than a 

characteristic of the numerical method used. 

Approach 

To rule Out spurious chaos, we investigated the effect of numerical resolution on the chaos. 

Accomplishment Description 

To better characterize turbulence, numerical solutions of the 

unaveraged Navier—Stokes equations were obtained using the 

Cray-2. A small change in initial conditions completely changes 

the values of the instantaneous velocity components a short time 

later. Initial conditions are a characteristic of decaying and 

forced turbulence. The research results are shown in the 

accompanying figure, where u 1 is a component of the velocity 

fluctuation at the numerical grid center, and, at time (t), u 0 is 

the initial velocity fluctuation. The overbar indicates a space 

average, v is the kinematic viscosity, and x0 is the initial scale 

of the fluctuations. The effect of spatial resolution on the flow 

sensitivity to small changes in initial conditions can be seen by 

comparing the initially perturbed (dashed) curves to the unperturbed (solid) curves. We take, as a measure of that sensitivity, 

the value of x, where a perturbed solution first shows a definite 

break with the unperturbed solution (incipient breakaway). It is 

clear that improved resolution increases the sensitivity of the 

solution to small initial-condition changes. The perturbed 

solution breaks away from the unperturbed one sooner for the 

more highly resolved cases with more grid points. Similar results 

were obtained for other velocity components and at other grid 

points. 

Significance 

The results indicate that the chaos observed in our numerical 

solutions is not spurious. 

Future Plans 

We will continue to investigate the physics of turbulence 

by obtaining and interpreting numerical solutions of the 

Navier—Stokes equations for homogeneous turbulence with 

and without mean gradients. 

Publications 

1. Deissler, R. G. "Is Navier—Stokes Turbulence Chaotic?" Phys. 

Fluids 29, no. 5 (1986): 1453-1457. 

2. Deissler, R. G. "Effect of Spatial Resolution on Apparent 

Sensitivity to Initial Conditions of a Decaying Flow as it 

Becomes Turbulent." J. Computational Phys. 100, no. 2 

(1992): 430-432.

2

(a)

'It 

Initial conditions 

Unperturbed 

Perturbed 0.1% 

- - - Incipient 

breakaway 

—3 

0 

—2 

—3 

0 

N 0

—2 

—3

.002 .004 .006 .008 .010 

(L)/x 2

)t 

Calculated evolution of velocity fluctuations at grid center 

(normalized by initial conditions) for initial Reynolds number 

(u )112 

x0 1v = 1108. Root mean-square fluctuations are spatially 

averaged; (a) 323 grid points, (b) 64 3 grid points, and 

(c) 7283 grid points. 

44
Self-Adaptive Grid Code Applied to Complex Three-Dimensional Flows 

G. S. Deiwert, Principal Investigator 

Co-Investigators: Carol B. Davies and Ethiraj Venkatapathy 

NASA Ames Research Center/Sterling Software/Eloret Institute 

Research Objective 

To develop a three-dimensional adaptive grid capability that will 

improve the accuracy and efficiency of computational fluid 

dynamics (CFD) flow-field calculations. 

Approach 

The self-adaptive grid code (SAGE) is applied to the initial grid 

after a CFD flow solver has computed an interim flow solution. 

SAGE adapts the grid to the solution by redistributing the grid 

points into high gradient regions, thus reducing the solution 

error. The interim flow solution is transferred onto the adapted 

grid and new files are created for input to the flow solver. The 

adaption procedure is based on a line by line grid-point-error 

equi-distribution scheme. Grid smoothness is maintained by 

applying torsional constraints between adjacent lines in the 

appropriate computational coordinate directions. 

Accomplishment Description 

The successful two-dimensional version of the adaptive grid 

code was extended to explore the complexities of threedimensional flow problems. Since three-dimensional flow 

structures need not be aligned to the initial grid, the ability to 

choose one-, two-, or three-directional adaptions in any order, 

make the code a powerful tool in obtaining the most appropriate 

grid adaption. Extensive use in a wide variety of three-dimensional 

problems proved the code to be reliable, robust, and effective. 

An example of a grid adaption for a generic hypersonic aerospace plane is shown in the accompanying figure. The initial 

grid (with equally distributed grid points) has been adapted with 

respect to the Mach gradients computed from the flow solution. 

The figure shows two views of the adapted grid: the cross plane 

and the symmetry plane. The clustering of points in the shock 

region can be clearly seen. A typical adaption takes up to 

12 megawords of memory and 5 Cray Y-MP minutes. 

Significance 

The SAGE code improves the accuracy of flow solutions. The 

line-by-line marching approach uses little computer time in 

comparison to the flow solver. 

Future Plans 

To fully analyze flow around complex structures, grids are 

frequently generated in multiple grid format. The SAGE code 

adapts single grids only; any matching of grids at common 

interfaces must be handled independently. The next upgrade to 

SAGE will include an internal grid transfer capability.

Publications 

1. Davies, C. B. and Venkatapathy, E. "Application of 

a Solution-Adaptive Grid Scheme, SAGE, to Complex ThreeDimensional Flows." AIAA Paper 91-1594, June 1991. 

2. Davies, C. B. and Venkatapathy, E. "The Multidimensional 

Self-Adaptive Grid Code, SAGE." NASA TM-1 03905, 1992. 

Adaptive grid for a generic aerospace plane; (a) computational 

domain, (b) adaptive grid in symmetry plane, and (c) adapted 

grid in cross plane.

45
Complex Three-Dimensional Turbulent Flows 

A. 0. Demuren, Principal investigator 

Old Dominion University 

Research Objective 

To develop accurate computational procedures for the analysis 

of complex three-dimensional turbulent flows with applications 

for ducts, jets, and turbomachinery. 

Approach 

The technical approach is to solve Reynolds-averaged 

Navier—Stokes equations and the continuity and energy equations with a computer code based on the finite-volume method. 

Turbulent stresses and fluxes that result from the averaging 

process are approximated with second-moment closure models. 

Conventional numerical methods usually suffer from deterioration in the convergence rates as computational grids are refined 

for improved accuracy. To overcome this shortcoming in the 

study, multigrid techniques are utilized to achieve gridindependent convergence rates. 

Accomplishment Description 

It has been possible to make computations of three-dimensional 

flows on very fine grids up to 2 million points using multigrid 

methods for convergence acceleration and Reynolds stress 

turbulence models for improved realism. Quite promising results 

have been obtained in a number of practical, useful flows such 

as jets in cross flow, impinging jets, and transition duct flows. 

An average job required about 20 megawords of memory and 

1 single processor Cray-2 hour. 

-

® Symmetry plane 

iu Wail 

IJVJ 

Pair of opposed jets in cross flow.

Significance 

We are close to having a general purpose three-dimensional 

flow solver with state of the art turbulence models that enable 

computations of important flows on refined grids within reasonable CPU times. The code can also be used as a test bed for the 

development of turbulence models. 

Future Plans 

To complete the development and testing of the code. At a later 

date, the code will be parallel ized. 

0

a) si mom. 

—1 

0 -J

Grid 

—3 --.3 

2 

—4

1 •

• 

0

(b) Residual mass 

—1 

E 

0 -J

—3 

—4' 

0 100 200 300 400 500 

Iteration number 

Multigrid convergence acceleration for a laminar jet in cross 

flow. 

46
Rarefied Hypersonic Condition Wake Structures 

Virendra K. Dogra, Principal Investigator 

Co-Investigators: Richard G. Wilmoth, James N. Moss, and Joseph M. Price 

ViGYAN, Inc./NASA Langley Research Center 

Research Objective 

To examine the basic feature of wake flows under rarefied 

hypersonic conditions and to provide a basis for establishing 

guidance as to when the continuum description becomes 

inappropriate for such flows. Precise determination of wake 

closure is a critical issue for aerobrake design. At typical perigees 

of 70-80 km, the local Knudsen numbers near separation are of 

the order of 0.1 or larger. For such conditions, the continuum 

description becomes suspect, and at higher altitudes the 

continuum description is invalid. 

Approach 

The direct simulation Monte Carlo (DSMC) method was used to 

simulate the flow about spheres for flow conditions that had 

been previously obtained in low-density wind tunnels. The 

DSMC approach is the only practical approach for simulating 

the complex physics of rarefied flows. 

Accomplishment Description 

DSMC simulations were made for six wind tunnel conditions at 

Mach 12 where the free-stream Knudsen number (length based 

on sphere diameter, d) ranged from 0.8 to 0.009. For these 

conditions, the calculated and measured drag coefficients (CD) 

agree to within 4%, which is well within the experimental error

As for the wake structure, the DSMC calculations show no 

evidence of separation in the wake, whereas a continuum 

solution has predicted incipient separation with a well defined 

recirculation zone. The wind tunnel case with a Knudsen 

number of 0.009 was simulated using 170,000 modeled molecules and 7,225 cells. The computation required approximately 

40 Cray Y-MP hours and 3 megawords of memory. 

Significance 

This study demonstrates the ability of the DSMC method to 

simulate hypersonic rarefied flows where continuum conditions 

may dominate the forebody flow (the overall drag) and yet 

the rarefaction effects are clearly dominant in the wake. This 

study provides a data base for assessing the bounds for which 

Navier-Stokes solutions are applicable for assessing the nature of 

wake flows for Aeroassisted Space Transfer Vehicles (ASTV). 

Future Plans 

The study will be expanded to examine the wake flows resulting 

from hypersonic flow about blunted cone shapes that are more 

representative of ASTV configurations. Also, the effect of stings 

and payloads on the wake flow will be examined. Hopefully, 

comparisons with wind tunnel measurements of wake flow 

structure can be made. 

Drag Density 

1.5 Kn = 0.009 

1.75 1.0 ptp1.25 0.60 

1.00 d2

I 010 

0.000 0.025 0.050 0.075 0.100 0.0 0.5 1.0 1.5 2.0

oo 

Knudaen Number, Kn_ r/d 

1.51 Mach Contours Streamlines 

• Kn = 0.009 Kn = 0.009 

1.0 1.0 

MACH

::: 

10.0

0.0 f 

y/d

:::

Vid 3.00 

4.0 __-. 0.60 

-l.a ..--, 3.0 -1.0 0.40 

• 2 0 030 

1.0 -1.5 ...L: 0.20 

05 0.10 

0.0 0.5 1:0 . . 2:0 0.0 0.5 1.0 1.5 2.0 0.05 

rid rid 

Direct simulation Monte Carlo calculations for a 2 cm diameter sphere at Mach 12.9 nitrogen flow and a wall temperature of 600 K.

47
Fuel—Air Mixing Enhancement by Jet—Shock Interactions 

I. Philip Drummond, Principal Investigator 

Co-Investigators: Peyman Givi, Cyrus K. Madnia, and Craig J. Steinberger 

NASA Langley Research Center 

Research Objective 

To study phenomena that control high-speed fuel—air mixing, 

and to use information gained from these studies to improve 

mixing and combustion efficiency in high-speed propulsion 

devices. 

Approach 

To solve, using an accurate numerical algorithm, the equations 

governing the mixing and chemical reaction of a multicomponent mixture in combination with physically realistic chemical 

kinetics models to describe combustion processes. 

Accomplishment Description 

A computer program was developed that solves the equations 

governing a multicomponent mixing and reacting flow. The code 

was validated against experiments involving nonreacting and 

reacting flow fields, and then applied in studies of combusting 

high-speed engine flows. Recent research has been directed 

toward the optimization of a scramjet combustor and the 

efficiency of fuel—air mixing and reaction in the engine. Mixing 

is significantly reduced in the engine as the combustor Mach 

number increases with flight Mach number, and mixing 

enhancement is required to achieve a sufficient degree of 

combustion efficiency. Because of this difficulty, alternate 

combustor fuel-injector configurations were studied to evaluate 

their potential for producing an improved degree of mixing and 

reaction. A fuel injector configuration was parametrically studied 

using the code to improve fuel—air mixing. A typical run required 

16 Cray Y-MP hours and 64 megawords of memory. A gaseous 

hydrogen-fuel jet, co-flowing with air, was processed through a 

10 degree oblique shock produced by a wedge. When the low 

density hydrogen jet passed through the shock, the pressure field 

and density field became misaligned, bifurcating the jet and 

producing a vortex pair. Mixing was significantly enhanced and 

large amounts of water were produced through reaction of the 

hydrogen and air. The accompanying figure shows the resulting 

water mass fraction, without and with the shock, in the central 

streamwise plane and at the final cross plane considered in the 

study. The shocked jet shows a large cross section of water with 

a peak mass fraction of 25%. The resulting vortex pair, at the 

same location, is also shown on the figure. When the hydrogen 

jet is not processed by the shock, little mixing of fuel and air 

results, and only a narrow annular ring of water, shown in the 

figure, is produced with a water peak of 0.6%. So, as can be 

seen, the fuel injector design results in a significantly larger 

amount of mixing and a chemical reaction that produces a 

marked increase in combustion efficiency.

Significance 

Numerical simulations provide an improved insight of the 

mechanisms controlling high-speed mixing and combustion in 

supersonic reacting flows. Techniques can be developed for 

enhancing mixing and reaction in supersonic combustors, thus 

improving the overall level of combustion efficiency. 

Future Plans 

Work is under way to optimize the degree of enhancement that 

can be achieved from the fuel-injector configuration. Other 

injector configurations that utilize swirl or alternate fuel-nozzle 

designs for enhancement will also be examined. 

Unshocked hydrogen-fuel jet. 

Shocked hydrogen-fuel jet. 

48
Aerobraking Studies of Three-Dimensional Nonequilibrium Viscous Flow 

C. T. Edquist, Principal Investigator 

Co-Investigator: T. J. Galambos 

Martin Marietta Astronautics Group 

Research Objective 

To use a computational fluid dynamics code to examine the 

aerothermochemical aspects of aerobraking during lunar return 

missions; validate the selection of aerobrake forebody shapes 

and sizes and payload configurations; calculate the heating 

environment experienced by the payload in the base region; and 

establish a matrix of torebody solutions. 

Approach 

We used the Langley Aerothermodynamics Upwind Relaxation 

Algorithm (LAURA) code which considers all of the aerothermochemical phenomena necessary for detailed consideration of lunar return aerobraking to study a number of possible 

aerobrake configurations and verify the conclusions arrived at by 

using less elaborate engineering procedures. 

Accomplishment Description 

A series of blunt cone forebody calculations of a lower corridor 

(high heating) trajectory near peak heating have been made. A 

70 degree sphere cone with base and nose radii of 22.5 and 

11.25 feet respectively and an aft edge radius of 1 foot was 

chosen for the reference geometry. An additional forebody 

configuration was formed by increasing the nose radius to its 

largest value corresponding to a spherical segment. Three points 

on a trajectory with ballistic coefficient (13) of 14 psI were 

considered and the reference forebody flow field at a 10 degree 

angle of attack was calculated at each trajectory point. The 

spherical segment forebody flow field was also calculated at the 

peak heating altitude. A converged solution required about 

20 Cray-2 hours and 40 megawords of memory. Solutions at 

peak heating for 13 values 10 and 20 psf were also obtained for 

the reference body. The effect of edge radius was studied by 

doubling that radius to 2 feet for both the sphere cone and 

spherical segment. Results showed general agreement with 

experiment. The calculated distribution of pressure and convective heating on the Viking forebody compared well with 

experimental wind tunnel data. Initial wake and payload flowfield calculations were made using a perfect gas version of 

LAURA to develop an understanding of the flow characteristics. 

Significance 

Aerobraking is a cost effective method of capturing a manned or 

unmanned vehicle on return from the moon or Mars or during a 

Mars encounter. The benefits of aerobraking will be maximized 

when the aerothermodynamics of high velocity, high altitude 

flight in the Earth's or other planetary atmospheres are well 

understood. Using a code such as LAURA will provide valuable 

insight for the design of these vehicles. 

Future Plans 

We will use the forebody solutions to examine various 

afterbodies and payload configurations to determine the most

efficient means of protecting the aerobrake cargo. Additionally, 

Mars entry and return missions will be studied because they 

require consideration of higher energies and different chemical 

species and reactions. 

Comparison otforebody convective heating to representative 

aerobrakes at peak heating during lunar return at a 10 degree 

angle of attack. 

Mach number contours in the plane of symmetry of the Viking 

aeroshell at Mach 14.9 and an 11.2 degree angle of attack.

49
Optimized

with

camber line

modificatior:

(3.8%

improvemem

Original 

M=1.7 ii4deg 

LD=18.75 L 0 = 18.032 

Lower surface of the wing-body configuration 

0 .05 .10 .15 .20 .25

Time (sec) 

1.5 

1.0 

.5 

0. 

< 0 

—.5 

—1.0 

—1.5

Aerodynamic Optimization of Supersonic Transport Aircraft 

Thomas A. Edwards, Principal Investigator 

Co-Investigator: Samson H. Cheung 

NASA Ames Research Center/MCAT Institute 

Research Objective 

To develop, validate, and apply computational fluid dynamics 

(CFD) optimization methods to design supersonic aircraft, 

focusing on the dual design objectives of low sonic boom and 

high aerodynamic performance. This research helps define the 

economic viability of low-sonic-boom aircraft. 

Approach 

A numerical optimization routine based on a sequential quadratic 

programming algorithm was coupled with an implicit Euler! 

Navier—Stokes solver and a parabolized Euler/Navier—Stokes 

code. For aerodynamic performance optimization, design 

parameters such as the wing-twist angle and airfoil camber are 

used to maximize the lift/drag VD) ratio. Sonic-boom minimization uses the equivalent area distribution as the objective 

function. The optimization procedure uses CFD to compute the 

equivalent area of the configuration, while the optimizer adjusts 

the cross-sectional area to achieve more desirable sonic-boom 

characteristics. 

Accomplishment Description 

The sonic-boom minimization and aerodynamic optimization 

procedures have been developed and validated. The design 

capability has been applied to a High-Speed Civil Transport 

(HSCT) concept. The first figure shows that the redesigned 

aircraft has a 3.75% improvement in L/D. These calculations 

typically require 10 Cray-2 hours and 16 megawords of memory. 

To exercise the sonic-boom minimization procedure, a target 

"flat top" pressure distribution for the sonic boom was specified 

as the objective function. The original configuration exhibited an 

N-wave pressure distribution with a small intermediate shock 

(shown in the second figure). The configuration gives a sonic 

boom that is much closer to the objective function while still 

satisfying fuselage size constraints.

Supersonic wing—body optimization. 

Significance 

The methodology and computational tools used in this project 

provide the capability to design a high-performance, low-boom 

HSCT configuration prior to fabricating wind tunnel models. 

This study will help define the economic viability of low-boom 

airplanes compared to alternative mission profiles. Other cruise 

conditions, such as takeoff and landing, can also be used as 

constraints or variables to design an HSCT with complete 

performance consideration. 

Future Plans 

The procedure is being applied to advanced low-boom designs. 

When the performance of these designs reaches goals identified 

for LID and sonic boom, wind tunnel models will be built and 

tested to validate the results.

Comparison of sonic booms; H/L = 133, M = 1.7, and a = 4.0. 

50
Rotary-Wing Ai rload Performance Prediction 

T. Alan Egoif, Principal Investigator 

Co-Investigator: Brian E. Wake 

United Technologies Research Center 

Research Objective 

To develop rotary-wing blade airload and performance prediction capability without reliance on empirical airfoil characteristics to provide a numerical rotor test facility for application to 

helicopters and high-speed propellers. 

Approach 

To refine and enhance advanced helicopter and high-speed 

propeller computational fluid dynamics (CFD) and wake 

modeling codes to include more rigorous representations of the 

rotor wake and provide viscous drag prediction capability, we 

coupled the methods and incorporated the rotor-wake influence 

in the CFD codes using velocity embedding techniques and 

transpiration boundary conditions. The wake influence is based 

on prescribed wake geometries from analytical, empirical, or 

numerical predictions. We developed improved wake geometry 

prediction capabilities based on free-wake vortex-lattice 

methods and performed validation studies of the component 

and coupled codes. 

Accomplishment Description 

Activity focused on validation of the rotary-wing viscous flow 

solver (NSR3D) and code refinements for robustness and 

improved convergence. NSR3D was applied to an initial 

oscillating-wing geometry for an assessment of the ability of 

the code to predict unsteady three-dimensional viscous-flow 

behavior. Initial unsteady predictions show reasonable agreement with test results. The study used 325,000 grid points, 

requiring approximately 100 Cray Y-MP hours for two cycles of 

oscillation and about 30 megawords of memory for an average 

case. In addition to this validation effort, the research was used 

to support a NASA contract for an acoustic investigation of 

rotor-tip designs using NSR3D. Over 400,000 grid points and 

130 Cray-2 hours (including trimming to the desired rotor thrust 

level) and 40 megawords of memory were required. 

Significance 

Application of NSR3D to geometries and unsteady flow conditions for test data that are available provides an assessment of 

thecode's prediction capabilities. 

Future Plans 

Refinements to the solver and the incorporation of the hover 

wake influence into NSR3D using a refined hover free-wake 

model are planned.

Publication 

Wake, B. E. and [golf, T. A. "Initial Validation of arr Unsteady 

Euler/Navier—Stokes Flow Solver for Helicopter-Rotor Airloads in 

Forward Flight." American Helicopter Society International 

Technical Specialists' Meeting on Rotorcraft Research. Atlanta, 

GA, Mar. 1991. 

Particle trace illustrating post-stall separated flow behavior of an 

unsteady wing as predicted by NSR3D.

51
Ii 

Transpiration Cooling for Scramjet Combustor Flow Fields 

Dean R. Eklund, Principal Investigator 

National Research Council/NASA Langley Research Center 

Research Objective 

To develop the capability to predict the surface heat-transfer 

rates generated by the turbulent, three-dimensional reacting flow 

fields occurring within supersonic scram jet combustors. Two 

intermediate steps are to develop the capability to efficiently 

model turbulent mixing and finite-rate kinetics within scramjet 

combustors. 

Approach 

A three-dimensional Navier—Stokes code with finite-rate 

chemistry models has been used to model the internal reacting 

flow fields. 

Accomplishment Description 

The SPARK three-dimensional Navier—Stokes code was validated 

for nonreacting internal flow fields with transverse injection. A 

series of calculations of the reacting flow from the transverse 

injection of hydrogen behind a rearward-facing step into a Mach 

2 flow were performed. The effects on the efficiency of the 

combustor of two geometric parameters, the angle of the expanding portion of the duct and the length of the constant area 

section, were examined. The calculations revealed that, even at 

the relatively low flight Mach numbers (5-7) associated with the 

conditions investigated, the chemical constituents are far from 

equilibrium and kinetics effects are important. The calculations 

will be compared to nitrogen temperature and density measurements obtained with a coherent anti-Raman spectroscopy system 

during a series of future experiments. The accompanying figure 

shows the calculated temperature contours along the centerline 

plane of the injector, the bottom wall of the combustor, and four 

cross-flow planes. Reaction is observed (red contours) in the

recirculating region upstream from the jet and in the mixing 

region between the injectant and the free stream downstream 

from the injector. A typical calculation required approximately 

30 Cray Y-MP hours and 10 megawords of memory. 

Significance 

Prediction of the surface heat-transfer rate and the design of 

methods to relieve these heating loads is one of the critical 

technical obstacles in the development of proposed air-breathing 

propulsion systems for hypersonic flight. The complexity of the 

three-dimensional, turbulent reacting flow fields within the 

combustors makes calculations of heat-transfer rates a challenging task. Successful modeling of these flow fields will provide a 

powerful design tool for these systems and assist in the analysis 

of their performance. 

Future Plans 

Simplified finite-rate chemistry models are being tested in the 

SPARK Navier—Stokes code. A zonal gridding capability will be 

incorporated into the SPARK code. Calculations with the 

General Aerodynamic Simulation Program Navier—Stokes code 

are also planned to compare the efficiency of the two codes. 

Finally, the codes will be used to model an experiment to test a 

transpiration cooling system for perpendicular injection of 

hydrogen fuel into Mach 2 flow. 

Publication 

Eklund, D. R. and Northam, G. B. "A Numerical Study of the 

Effects of Geometry on the Performance of a Supersonic 

Combustor." AIAA Paper 92-0624, Jan. 1992. 

Iinperature contours red = 2,200 t; blue = 20c) i ;or the normal injection vi /idroeii iuei behind a reatoatd-ldcalo step into a 

heated Mach 2 airstream. The direction of the airstream is from left to right. 

52
(a)

)lane - 

8 CL 

0.

Flow Solver for Euler Equations on Unstructured Tetrahedral Meshes 

Larry L. Erickson, Principal Investigator 

Co-Investigator: M. Jahed Djomehri 

NASA Ames Research Center/Eloret Institute 

Research Objective 

To assess and improve unstructured three-dimensional flow 

solvers and mesh generators for complex aerodynamic shapes at 

transonic/supersonic speeds (specifically the FELISA code). 

Approach 

The method is based on an explicit Taylor—Galerkin weighted 

residual scheme with local time-stepping for steady computations. The flow field is an assembly of tetrahedral elements 

generated by an advancing front technique. Variable grid 

spacing is achieved by user-specified control parameters. The 

grid can be adapted to flow features, such as density, by the 

equi-distribution of error concept. 

Accomplishment Description 

FELISA was applied to two standard wind tunnel configurations 

(a low-aspect-ratio wing, and a cone cylinder) used to assess 

sonic boom signatures. The cone cylinder and its adapted grid 

are shown in the accompanying figure. The mesh has been 

adapted to the flow features and consists of 61,299 grid points. 

The sectional view of the grid on the xy- and xz-symmetry 

planes and an enlarged view of the body are shown. The main 

feature of the flow field is an attached weak bow shock at the 

cone apex and an expansion wave at the intersection of the cone 

and cylinder. The grid is concentrated in these regions. The 

residual of the computed solution was dropped three orders of 

magnitude within 2,000 time iterations. This calculation 

required 1.5 Cray Y-MP hours and 10 megawords of memory.

More complicated geometries require about 30-40 megawords. 

Although the geometry is simple, the formation of a smooth 

grid around the nose of the cone, whose half angle was only 

3.6 degrees, was complicated. The surface and volume grid 

generators were improved to handle these configurations. 

Computed relative pressure data, LP/P O , along a line half a cone 

length away from the body in the free stream flow direction, has 

been extrapolated to 10 cone lengths away from the body and 

compared against wind tunnel measurements. The figure 

contains computed solutions obtained on a multiblock grid by 

the structured grid solver TEAM. The plot shows pressure 

signatures due to bow shock and expansion waves. Results 

appear in agreement near the shock, but differ slightly near the 

expansion wave. The inconsistency is believed to be due to the 

coarseness of the grid to the right of the expansion wave. 

Significance 

Automated solution-adaptive grid generation for realistic aircraft 

appears to be possible with this approach and could be a 

significant step toward making computational fluid dynamics 

useful to the practicing aerodynamicist. 

Future Plans 

Calculations are planned to assess the versatility of the mesh 

generator and flow solver on further complex three-dimensional 

problems such as supersonic wing—body and complete Lear-jet 

configurations. 

M=1.68,cz=0,H/L=10 

Swans.., 310k elem., 

extrap. from H/L = 1/2 

Team, extrap. from •\ 

H/L1.1 

ExperImental date 

(b) •: 

.5 1.0 1.5 2.0 

Enlarged view X/L 

Cone cylinder at M,,. = 1.68; a =0 degrees; (a) surface and volume grid and (b) off-body pressure signatures.

53
Discrete Particle Simulation of Compressible Flow 

William J. Feiereisen, Principal Investigator 

Co-Investigators: Jeffrey McDonald, Brian Haas, lain Boyd, Donald Baganoff, Michael Fallavollita, 

Terry Denery, and Avijit Goswami 

NASA Ames Research Center/Stanford University 

Research Objective 

To develop and validate a new particle-simulation method for 

rarefied hypersonic flow and its application to realistic threedimensional geometries, modeling of nonequilibrium thermochemistry, and efficient implementation on supercomputer 

architectures. 

Approach 

Direct particle-simulation methods model a rarefied flow as a 

large collection of discrete particles which travel and interact 

through collisions. To improve the direct simulation Monte Carlo 

technique, new collision selection rules and other algorithmic 

changes permit efficient implementation of the new methods on 

vector supercomputers. Unlike previous models, collision 

models are enhanced to capture coupled vibration-dissociation 

behavior in high-temperature gases. 

Accomplishment Description 

New collision models were developed, accounting for multilevel 

vibrational quantum transitions and vibrationally favored 

dissociation rules, that capture essential nonequilibrium thermochemistry. Results agree favorably with experimental data. The 

code was employed to simulate proposed aerobraking maneuvers of the Magellan spacecraft orbiting Venus. Temperatures 

and flow-field particles are plotted in the accompanying figure 

for Mach 35 flow at Knudsen numbers exceeding 0.10 about the 

complex multibody geometry. New models for surface heat

transfer permit assessment of surface temperatures. The significance of flow-field molecular collisions was evaluated and 

compared to free molecule flow. Employing 400,000 particles 

and 44,000 domain cells, each simulation required less than 

2 Cray-2 hours and 12 megawords of memory. 

Significance 

Computational predictive capability is essential to assess the 

aerodynamic and thermal environment encountered by 

atmospheric-entry spacecraft and high-altitude maneuvering 

vehicles. Direct particle-simulation methods are most appropriate for modeling rarefied flows since computational fluid 

dynamics based upon continuum Navier—Stokes equations are 

not applicable. The computational burden and model simplicity 

of previous methods are improved upon in this work, thus 

permitting large-scale simulation of flows about geometries of 

engineering interest. 

Future Plans 

Applications will include studies of rarefied turbulence and 

Magellan aerobraking at various angles of attack. Models will 

be enhanced for vibrational kinetics and for particle—surface 

interaction. Methodologies to extend the simulation application 

into the near-continuum flow regime with minimal computational penalty will be investigated. 

Temperatures and particle distribution during proposed aerobra king of the Magellan spacecraft through the atmosphere of Venus; 

altitude = 125 km, M = 35, and Kn = 0.103. 

54
.2 .3 .4 .5 .6 .7 .8 

1011 

1010 

i09 

108 

107 

106 

105

104

1

100 

.01 

U) 

0 

U 

a) 

a) 

C) 

C.)

10

U) 

E 

U 

a) 

0 

CL 

a) 

0 

C) 

Aeroassist Flight Experiment Flow Simulation 

William J. Feiereisen, Principal Investigator 

Co-Investigators: G. E. Palmer, E. Venkatapathy, D. S. Babikian, D. K. Prabhu, and E. E. Whiting 

NASA Ames Research Center 

Research Objective 

To use the numerically computed flow fields in the pre-flight 

design of the high-resolution spectrometers and total-radiation 

detectors to be carried on board the Aeroassist Flight Experiment 

(AFE) vehicle. 

Approach 

The three-dimensional aerothermodynamic environment 

around the AFE vehicle is computed using thermochemical 

nonequilibrium Navier—Stokes codes. The computed flow fields 

are then used in conjunction with high-resolution radiation/ 

spectral codes to estimate the radiative flux at the various points 

on the vehicle surface where the radiometers and spectrometers 

are located. 

Accomplishment Description 

The computed three-dimensional real-gas flow field was used 

with a high-resolution spectral code to determine the spectrum 

incident on the vehicle surface. This was done for various lines 

of sight from the proposed instrument locations on the vehicle. 

The computed spectra were used with the instrument response 

functions to determine the sensitivities and data-integration time 

intervals. A sample spectrum and the computed detector current 

corresponding to this spectrum are shown in the accompanying 

figure. This particular calculation was for a line of sight in the 

forebody stagnation region of the flow. The wake radiation was 

studied with both the nonequilibrium flow solutions and idealgas solutions for the extended wake. The calculations were used 

to determine the orientations and view angles of the instruments. 

The sensitivity of the radiation calculations was determined for 

different entry velocities and altitudes and for various vehicle 

masses. 

Significance 

The computations helped in the preflight design of some of the 

radiometers to be carried on board the AFE vehicle. The 

computations highlighted the need for better thermochemical 

models in the expanding flow regime and in the wake.

Future Plans 

There are no future plans for this project because it has been 

canceled by NASA. 

Publications 

1. Venkatapathy, E.; Palmer, G. E.; and Prabhu, D. K. "AFE Base 

Flow Calculations." AIAA Paper 91-1372, June 1991. 

2. Strawa, A. W.; Park, C.; Davy, W. C.; Craig, R. A.; Babikian, 

D. S.; Prabhu, D. K.; and Venkatapathy, E. "Radiometric 

Investigation of the Wake Flow of the Forthcoming Aeroassist 

Flight Experiment." AIAA Paper 91-1408, June 1991. 

Wavelength (.t) 

Computed spectrum and detector current in the forebody 

stagnation region of the Aeroassist Flight Experiment vehicle; 

detector angle (ad) = 0.0 129 sr, area (Ad) = 6.25 x 10 cm2, 

resolution tA,) = 6 x 10 i.

55
El 6xL Supersonic Laminar Low-Control Experiment 

Michael C. Fischer, Principal Investigator 

Co-Investigator: Chandra S. Vemu ru 

NASA Langley Research Center 

Research Objective 

To design and develop passive and suction laminar flow gloves 

for the Fl 6xL supersonic laminar flow control experiment. 

Approach 

Three-dimensional Euler and Navier—Stokes codes are used to 

predict the flow over the Fl 6xL configuration at supersonic 

Mach numbers. 

Accomplishment Description 

The Fl 6xL SHIP] and SHIP2 configurations were analyzed by 

using the EMTAC code at Mac 1.7 and a = 3.0 degrees. The 

pressure coefficient contours over the planform indicated that 

the shock location for SHIP2 moved further aft compared to the 

SHIP1 configuration. However, the strength of the shock is 

stronger compared to the canopy shock of SHIM. The canopy 

shock sweeps across the wing in a location that may limit the 

fullest extent of laminar flow from being realized, especially at 

lower Mach numbers. In order to achieve 50-60% laminar flow 

on the wing, a canopy fairing may be necessary to move the 

shock further aft. The SHIP2 canopy was faired in the rear part to 

study the effect of canopy shape. An analysis of the new 

configuration indicated that the fairing weakened the shock, but 

the shock location did not change much compared to the SHIP2 

geometry. Analysis of the boundary layer and compressible

stability (with and without curvature effects) over the inboard 

section of the El 6xL SHIM glove configuration was conducted. 

The results were compared to the flight-test transition data to 

determine a critical N-factor value. The boundary layer calculations showed a wide difference in Re values whereas the 

N-factor values at 4% chord did not indicate any significant 

difference. The F16xL SHIP2 configuration was analyzed at 

Mac 1.9 and a 2.965 degrees using a C-H grid and the 

TLNS3D code. 

Significance 

Laminar flow control has been identified as one of the key 

technologies necessary for ensuring the economic viability of 

High Speed Civil Transport. Laminar flow glove shapes designed 

in this research program will be flight tested and the wing 

pressures measured in flight will allow validation of computational design/analysis methods at supersonic speeds. 

Future Plans 

The Euler and Navier—Stokes codes will be used to analyze 

laminar flow glove shapes for the El 6xL flight experiment. Grids 

with different topologies and distributions will be generated 

using the GRIDGEN program and the theoretical pressure 

distributions will be compared with the flight data. 

N,, = 1.7 

a 3.0 deg 

- Single place Canopy 

Dual place canopy 

Z= 70.00 Inches 

20

Dual place Single place 

canopy canopy 

How field 

Laminar 

glove 

outline

Pressure coefficient contours over an F16xL aircraft; M,,, = 1.7, a = 3.0 degrees. 

56
Analysis of Slender-Wing Geometries using Unstructured Grids 

Neal T. Frink, Principal Investigator 

Co-Investigators: Paresh Parikh and Shahyar Pirzadeh 

NASA Langley Research Center 

Research Objective 

To assess the accuracy and efficiency of a new three-dimensional 

unstructured-grid software package to compute transonic flow 

around a high-speed transport configuration. 

Approach 

To construct an unstructured tetrahedral grid around a generic 

High-Speed Civil Transport (HSCT) configuration using the 

advancing front-grid generation code VGRID3D and compute 

the inviscid transonic flow field surrounding the configuration at 

a moderate angle of attack with the unstructured upwind flow 

solver USM3D. The results will be displayed with the graphic 

postprocessing code VPLOT3D and quantitative comparisons 

with experimental data will be provided. 

Accomplishment Description 

Two unstructured tetrahedral grids were generated around an 

HSCT configuration consisting of 184,997 cells and 33,499 nodes 

for the coarse grid, and 379,561 cells and 67,848 nodes for the 

fine grid. The total number of cells was kept down by longitudinally stretching the cells near the configuration by a 4:1 ratio. 

Grid resolution was maintained in the spanwise direction to 

resolve vortex-induced pressure gradients. Transonic inviscid 

solutions were obtained for the conditions M,,,, = 0.901 at 

a = 6.47 degrees and M = 1.194 at a = 3.34 degrees. The

coarse grid residual was reduced 2.5 orders of magnitude using 

approximately 1 Cray Y-MP hour and 12 megawords of memory. 

The fine grid residual was reduced 4.1 orders of magnitude and 

required approximately 7 Cray Y-MP hours and 24 megawords 

of memory. The accompanying figure depicts the coarse grid 

solution. The surface velocity vectors and field Mach contours 

indicate the presence of leading-edge vortex flow. A second 

vortex is formed on the outboard panel, while the inboard vortex 

passes over the aft wing region. Comparisons with experimental 

data show generally good agreement of the coefficient of 

pressure inboard of the suction peak and the expected overprediction of the outboard suction peak by the inviscid solution. 

Significance 

The USM3D flow solver provides efficient solutions to the Euler 

equations within the accuracy of the inviscid assumptions. 

These solutions were the first results produced by USM3D with 

leading-edge vortex flow, and with flow in the low-supersonic 

speed range. 

Future Plans 

Work to reduce the turnaround time to produce a satisfactory 

grid and flow solution is continuing. The range of applications 

for the software is expanding. Work is under way to develop an 

unstructured viscous-grid generator and flow solver. 

Unstructured-grid solution for a generic High-Speed Ca ii Trdnsport contiu ration; 'l = 0.907, a = 6.47 degrees.

57
Turbulent Flow Past a Complete Hypersonic Reentry Configuration 

Datta Gaitonde, Principal Investigator 

Co-Investigator: Joseph Shang 

WL/FIMM, Wright Patterson AFB 

Research Objective 

To examine the turbulent flow field about complete lifting-body 

configurations at hypersonic speeds utilizing an accurate and 

efficient numerical algorithm. 

Approach 

The full three-dimensional mean compressible Navier—Stokes 

equations in mass-averaged variables are solved in a finite 

volume formulation utilizing Roe's flux-difference split algorithm 

for the inviscid fluxes and centered evaluation of viscous fluxes. 

The implicit numerical algorithm is based on Gauss—Seidel line 

relaxation. Turbulence closure is presently achieved with the 

Baldwin—Lomax algebraic eddy-viscosity model. 

Accomplishment Description 

The choice of algorithm was based upon a critical examination 

of MUSCL-based higher-order upwind methods with a range of 

limiters. Specific emphasis was placed on the prediction of heat 

transfer rates in viscous flows. The flow past the X24C-1 OD 

configuration was then solved with a mesh of roughly 

500,000 grid points. The so-called "carbuncle" anomaly in 

the nose region was eliminated by a combination of different 

splittings without affecting the overall accuracy. Analysis of the 

computed turbulent flow field, including a comparison with 

experimental heat transfer data, is ongoing. The accompanying 

figure shows surface pressures and Mach contours with an Euler 

calculation. 

Significance 

The ability to accurately and efficiently simulate flow fields 

past complete reentry configurations is critical in the design 

process, especially at points in the flight path where the stringent 

conditions encountered are extremely difficult, if not impossible, 

to simulate in wind tunnels. This effort should provide a costeffective alternative in vehicle design. 

Future Plans 

The theoretical model will be extended to incorporate hightemperature effects such as thermochemical nonequilibrium and 

ionization. A two-dimensional version of the code, employing 

nonequilibrium five-species chemistry and a simple harmonic 

oscillator model for vibrational energy excitation, has already 

been developed. A two-equation turbulence closure model will 

also be examined.

Publications 

1. Gaitonde, D. and Shang, J. "The Performance of FluxSplit Algorithms in High-Speed Viscous Flows." AIAA 

Paper 92-0186, Aerospace Sciences Meeting Exhibit, 

Reno, NV, 1992. 

2. Kroll, N.; Aftosmis, M.; and Gaitonde, D. "An Examination 

of Several High-Resolution Schemes Applied to Complex 

Problems in High-Speed Flows." WL-TR-91 -3089, WrightPatterson AFB, 1991. 

Surface pressures and Mach contours with an Euler calculation; 

M = 6, a = 6 degrees. 

58
Integrated Hypersonic-Propulsion Flow Paths 

Joseph L. Garrett, Principal investigator 

Co-Investigators: Kevin Van Dyke and Balu Sukar 

Pratt & Whitney/General Electric Aircraft Engines 

Research Objective 

To calibrate and use the General Aerodynamic Simulation 

Program (GASP) to design hypersonic-propulsion flow paths. 

Approach 

Calibration of the GASP code was done in two parts. First, the 

GASP chemistry algorithm was validated against the predictions 

of the two-dimensional kinetics code. The GASP code was then 

used to predict the film-coolant distribution for the subscale 

National Launch System (NLS) calorimeter nozzle. Pressure, 

temperature, and heat transfer data from the calorimeter-nozzle 

test will be used to validate the pretest predictions. In addition to 

the NLS calorimeter nozzle predictions, the GASP code was 

used to compute the Arnold Engineering Development Center 

(AEDC) 50 megawatt arc-tunnel nozzle flow for the National 

Aero-Space Plane (NASP) Mach 11.5 combustor rig. 

Accomplishment Description 

An average run required 8 Cray Y-MP hours and 8 megawords 

of memory for convergence. Pretest predictions for the NLS 

calorimeter nozzle indicated a potential problem in the filmcoolant delivery design that was substantiated and corrected 

through the use of a model water-flow rig. Pretest predictions of 

the AEDC 50 megawatt arc-tunnel nozzle flow were used in 

designing the fuel injection system for the NASP Mach 11.5 

combustor-rig test. 

Significance 

A potential for burning up the NLS calorimeter injection ring due 

to a maldistribution of coolant flow was averted. Further, pretest 

predictions for the entire system will provide the opportunity to 

validate the GASP code for film injection and heat transfer 

analyses.

Non dlms,.sonaI

I

( ' I velocity contours

0.04000 • 

0.10000 

0.16000 

0.22000

0.281103 

0.3 111(11 i.

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t 'L U. 460 lit ' .' \\%iil\\\ 

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Preliminary National Launch System subscale secondary flowdistribution geometry. 

Future Plans 

Upon completion of the NLS and NASP rig tests, the pretest 

predictions will be compared with test results, and appropriate 

modifications to the GASP algorithm will be made, if necessary. 

The GASP code will then be used for design studies in both the 

NASP and NLS propulsion systems.

Annulus

36 Flow metering holes 

L72 Vanes

300 

0.055 High

L EDM 

material 

View A

removal 

Computational fluid dynamics analysis indicates a 

ma/distribution of fl/rn coolant that impacts the design of the 

secondary flow-distribution system.

59 
Compressible Turbulent Flows 

Thomas B. Gatski, Principal Investigator 

Co-Investigator: Joseph H. Morrison 

NASA Langley Research Center 

Research Objective 

To develop a compressible Navier–Stokes algorithm that can 

use both two-equation and Reynolds stress turbulence closure 

models to solve complex turbulence flows. 

Approach 

A multiblock, cell-centered, finite-volume scheme with good 

shock-capturing capabilities, accuracy, and general geometric 

capabilities is developed. A Roe flux-difference splitting technique coupled with a MLJSCL scheme is applied to the mean 

conservation equations for the mass, momentum and energy, 

and the transport equations for the turbulent Reynolds stresses. 

Viscous fluxes are discretized using a finite-volume representation of a central-difference operator and the source terms are 

treated as an integral over the control volume. 

Accomplishment Description 

A code capable of handling complex high-speed compressible 

flows has been developed. Both two-equation and Reynolds 

stress-transport equations are included in the numerical algorithm. In addition, near-wall models are used in the turbulence 

closure scheme in order to integrate directly to solid boundaries

where needed. An average run takes about 4 Cray Y-MP hours 

and uses 10 megawords of memory. The accompanying figures 

show the Reynolds stress-component profiles. The flat plate 

section precedes the ramp which starts at x = 0.5. 

Significance 

Complex turbulent flow fields require more sophisticated 

turbulence closure models than the simple eddy-viscosity flow 

fields. Unfortunately, incorporating the Reynolds stress models 

into Navier–Stokes codes is cumbersome, especially for compressible flows where mass-flux and heat-flux correlations are 

also present. With the development of the present code, it is 

possible to validate Reynolds stress-turbulence closure models in 

complex compressible flows and evaluate their performance 

relative to simpler turbulence closure models. 

Future Plans 

We will apply the Navier–Stokes code to a variety of high-speed 

compressible flows—both wall-bounded and free-shear. The 

results will be used to compare the predictive capability of both 

two-equation and Reynolds stress models with available 

experimental data. 

x=O.396 x 0.499 x=O.598 x=O8.80 xO.396 x =O.499 x 0.598 x= O.808 

L L H 6 

0 6 L L LL 

Normal stress component, t (X 102) Shear stress component, ty (X 102) 

Normal Reynolds stress component profile for a 10 degree 

compression ramp.

Reynolds shear stress component profile for a 10 degree 

compression ramp. 

60
Hypersonic Flows in Chemical and Thermal Nonequilibrium 

Peter A. Gnoffo, Principal Investigator 

NASA Langley Research Center 

Research Objective 

To simulate hypersonic flow fields in chemical and thermal 

nonequilibrium. Particular emphasis is placed on the Aeroassist 

Flight Experiment and the convective and radiative heating 

distributions on lunar transfer vehicles (LTV) configurations. 

Approach 

The three-dimensional Navier—Stokes equations, 11 speciescontinuity equations, a total-energy equation, and a vibrationalenergy equation are solved in a fully coupled, point-implicit, 

symmetric, total-variation-diminishing algorithm. Uncoupled 

radiative heating levels can be calculated from converged profile 

solutions. 

Accomplishment Description 

The hypersonic flow over 70-degree spherically capped cones 

with 50-foot-diameter bases and spherical-segment brakes that 

are near peak heating conditions for an LTV earth return have 

been studied. Three-dimensional flows require 5-20 Cray Y-MP 

hours and less than 40 megawords of memory, depending on the 

quality of the initial solution and grid density. Significant 

nonequilibrium effects are confined to a region less than 1 inch 

thick in a shock layer that is approximately 2 feet thick. As 

shown in the accompanying figure, grid topologies and adaption 

algorithms were developed to resolve nonequilibrium phenomena behind the bow shock and eliminate singularity-induced 

irregularities in heating contours in the stagnation region. 

Significance 

Total heating loads on LTV configurations border the upper 

limits of present thermal protection system materials and are 

dominated by radiative heating. Accurate predictions of the 

heating load are required to assess the cost of aerobraking 

options for LTV applications.

Future Plans 

We will study some of the difficulties with the convergence of 

nonequilibrium radiative heating levels with equilibrium 

radiative levels for thick shock layers dominated by equilibrium 

processes. 

Publications 

1. Gnoffo, P. A.; Price, J. M.; and Braun, R. D. "On the Computation of Near-Wake, Aerobrake Flow Fields." J. Spacecraft 

and Rockets 29, no. 2 (Mar./Apr. 1992): 182-189. 

2. Gnoffo, P. A. "Point-Implicit Relaxation Strategies for Viscous 

Hypersonic Flows." In Computational Methods in Hypersonic Aerodynamics, ed. T. K. S. Murthy. Southampton, UK: 

Computational Mechanics Publications, 1991. 

Convective heating contours and a surface grid for a lunar 

transfer vehicle.

61
Unsteady Delta-Wing Flow 

Raymond Gordnier, Principal Investigator 

WL/FIMM, Wright Patterson AFB 

Research Objective 

To calculate the flow field over a delta wing undergoing an 

oscillatory rolling motion, thus simulating a wing-rock behavior. 

This requires the ability to simulate both the steady flow over a 

delta wing at a fixed angle of attack and roll angle, and the 

unsteady flow over a delta wing undergoing a rolling motion. 

Enhancing the maneuverability and controllability of current and 

future fighter aircraft requires the ability to accurately predict the 

vortex-dominated flow environment delta wings encounter 

during rapid maneuvers. 

Approach 

The unsteady three-dimensional full Navier—Stokes equations 

are solved using the Beam—Warming implicit, approximately 

factored algorithm. A subiteration procedure improves the 

accuracy and robustness of the algorithm. The governing 

equations accommodate moving grids, providing the means 

to incorporate the unsteady motion of the body. The 

Baldwin—Lomax turbulence model is used for computing 

turbulent flows. 

Accomplishment Description 

Computations were performed for an 80 degree swept delta wing 

at a 30 degree angle of attack and roll angles of 0 and 45 degrees. 

Both fully laminar and fully turbulent solutions were considered 

since the experimental Reynolds number was 400,000. Comparisons with experimental surface pressure measurements showed 

good agreement for both the laminar and turbulent computations. The laminar flow solution more generally represented the 

character of the experimental surface pressure. However, the 

turbulent solution predicted the minimum pressure under the 

primary vortex more closely. The effects of grid resolution, 

boundary conditions, and grid topology on the computational 

results were also investigated. A typical run on the finest grid 

required approximately 75 Cray-2 hours and 71 megawords of 

memory.

Significance 

The successful computation of a delta wing undergoing a static 

roll is an important first step in developing the ability to simulate 

the unsteady aerodynamics of a dynamic roll maneuver. These 

computations provide a significant validation of the present 

numerical technique before undertaking the more costly and 

complicated wing-rock computations. 

Future Plans 

The current numerical procedure will be applied to an 80 degree 

swept delta wing undergoing a forced wing-rock motion and the 

unsteady aerodynamics of the maneuver will be investigated. 

Delta Wing at Steady Roll Angle 

a = 30° 

0 = _0

Re = 400,000 

M. = 0.2 

X/L = 04 

Vorticit y Contours 

Computations for an 80 degree swept delta wing at 30 degrees 

angle of attack. 

62
Numerical Calculation of a Three-Dimensional Separated Flow 

Isaac Greber, Principal Investigator 

Case Western Reserve University 

Research Objective 

To elucidate the vortex structures that develop in the threedimensional flow pattern induced by a nominally two-dimensional 

shock-wave boundary-layer interaction in a wind tunnel or inlet 

configuration. Special effort was taken to investigate the role of 

boundary-layer thickness and nonuniformities of the approach 

boundary layer on the three-dimensionality of the interaction 

pattern, especially on the vortex-like pattern that develops. 

Approach 

Three-dimensional Navier—Stokes codes were used to compute 

the flow field in a rectangular geometry. The shock wave was 

modeled as an inflow condition over a section of the computational boundary. The effects of turbulence models were investigated along with the ability to describe the corner flows that 

occur. Graphical depictions of the vortex structures were 

developed. 

Accomplishment Description 

A three-dimensional Navier—Stokes code using a LUSSOR 

scheme, the NASA RPLUS code, was modified for this application. Computations of two-dimensional shock-wave impingement in laminar and turbulent flow were made and showed 

reasonable agreement with published computational results. 

The results of duct flows show similar vortex structure in both 

laminar and turbulent flows and the structure agrees with 

experimental observation. The calculated results demonstrate 

that the confinement of the flow is the essential cause of the 

three-dimensionality of the separation. The thickness of the

approach boundary-layer relative to the duct width is a critical 

parameter in determining the three-dimensional shape of the 

separation structure. Each three-dimensional computation 

required about 8 Cray Y-MP hours and 5.5 megawords of 

memory. 

Significance 

The three-dimensional interaction with vortical structures is 

fundamental not only to shock-induced separation in a confined 

geometry, but to any separated flow field that has a spanwise 

length scale. The results are directly applicable to inlet design for 

jet engines operating in supersonic flight. 

Future Plans 

In order to have an accurate quantitative comparison with 

experimental results for duct flow, further work on developing a 

turbulence model that reflects the physical characteristics of the 

flow field is required. A finer grid is needed and a numerical 

scheme with less artificial damping will be extremely useful. 

Publications 

1. Yang, W. and Greber, Isaac. "Numerical Simulation of a 

Three-Dimensional Shock Wave/Boundary Layer Interaction." Presented at the AIAA Mini-Symposium on Aerospace 

Technology in Northern Ohio, Cleveland, OH, Sept. 1990. 

2. Yang, W. "Numerical Simulation of a Shock Wave/Turbulent 

Boundary Layer Interaction in a Duct." Ph.D. Dissertation, 

Case Western Reserve University, Cleveland, OH, May 1992. 

Shock/turbulent duct flow interaction; M = 2.9, Re = 736 x 106, a = 12.99 degrees, ö/w = 0.0712. (a) Streamwise velocity field of the 

separation structure near the center plane. (b) Spanwise velocity field of the corner vortex structure.

63
Spatially Evolving Reactive Jets 

Fernando F. Grinstein, Principal Investigator 

Naval Research Laboratory 

Research Objective 

To understand the dynamics and topology of the large-scale 

coherent structures controlling the development of reactive 

free jets. 

Approach 

To develop a numerical solution of the three-dimensional timedependent compressible flow equations on structured grids. The 

model uses a fourth order flux-corrected transport (FCT) algorithm and appropriate inflow and outflow boundary conditions 

for convective transport, one-step, irreversible, finite-rate 

(Arrhenius) chemistry, and realistic (temperature- and speciesdependent) diffusive transport. No subgrid modeling other than 

that provided by the FCT high-frequency filter is included in the 

numerical model. 

Accomplishment Description 

The investigation of the effects of exothermicity, diffusive 

transport, and three-dimensional spanwise excitation on reactive 

mixing layers was continued. We performed simulations of 

spatially evolving round and square free-jets. The accompanying 

figure shows a typical instantaneous visualization of the outstanding topological features of a developed, nonreactive square 

jet. Shown are the rolling-up of the square vortex sheet near the 

inflow into vortex-ring structures; deformation and axis switching 

of the rings due to self-induced velocity in the corners; longitudinal (streamwise) structures caused by the stretching of the 

streamwise vorticity introduced by the corners; and a disorganized region of the flow depicting the later stage of vortex 

merging. Streamwise vorticity of both signs introduced as a 

consequence of the self-induced distortion of the loops is crucial 

in determining the larger entrainment and mixing properties of 

the square jet relative to its circular counterpart. This simulation 

used 120 x 80 x 80 computational cells, and about 29 Cray Y-MP 

hours. 

Significance 

The simulations provide insight into the physics of large-scale 

coherent structures in free-shear flows and the mechanisms 

affecting the growth of mixing layers and the transition to 

turbulence. This effort advanced state-of-the-art transitional 

shear-flow simulations.

Future Plans 

The studies will be extended to rectangular jets and the investigation of flame-extinction phenomena in jets. 

Publications 

1. Grinstein , F. F. and Kailasanath, K. "Compressibility, 

Exothermicity, and Three-Dimensionality in SpatiallyEvolving Reactive Shear Flows." 13th ICDERS, Nagoya, 

Japan, July 1991. 

2. Grinstein, F. F. "Coherent-Structure Dynamics in SpatiallyDeveloping Square Jets." Bull. Amer. Phys. Soc. 36 (1991): 

2699. 

higher vorticiry rnagrItuo - .

merging 

,1J' '$•. 

).

%. 

• ku,giodinaI vortices 

.'. .

med vortex ring 

axis switching 

Visualization of a square jet in terms of vorticity magnitude. The 

jet is subsonic (M = 0.6) and is axially forced (fD/U = 0.3). 

64
Three-Dimensional Atmospheric Simulation Model 

W. L. Grose, Principal Investigator 

Co-Investigators: W. T. Blackshear, R. S. Eckman, and R. E. Turner 

NASA Langley Research Center 

Research Objective 

To conduct a three-dimensional simulation of the evolving 

atmospheric circulation and trace constituent distributions over 

a multiyear cycle in order to better understand stratospheric 

dynamics, transport, and photochemistry and to apply this 

knowledge to support the Upper Atmospheric Research Satellite 

WARS) program. 

Approach 

A three-dimensional, spectral, primitive-equation, atmospheric 

model incorporation chemistry has been developed. It consists 

of two components: a model that calculates the winds and 

temperatures to describe the global atmospheric dynamics, and 

a separate transport/chemistry model (driven by the simulated 

dynamics) to describe transport of photochemically active trace 

constituents. 

Accomplishment Description 

A model simulation provided a baseline tor assessing the 

temporal and spatial variation of stratospheric trace species. 

Initial efforts are in progress to compare model results of

stratospheric chlorine monoxide (CIO) with measurements 

made by an instrument on board the recently launched UARS. 

Calculations of enhanced levels of CIO resulting from heterogeneous chemical processes during the southern polar winter and 

spring are in satisfactory agreement with observation. Studies of 

the budget of the odd nitrogen family in the stratosphere are 

continuing. Modifications to the transport/chemistry module are 

made on a continuing basis in an effort to enhance model 

performance. Typically, the model requires 600 Cray Y-M[' 

seconds per model day and uses 4 megaworcls of memory. 

Significance 

This model provides increased understanding of stratospheric 

dynamics, transport, and photochemistry, and supports the 

UARS and Earth Observing System programs. 

Future Plans 

Improvements in the horizontal and vertical resolution of the 

model are in the planning stage. The increased resolution will 

better describe small-scale features that are presently not fully 

resolved in the model. An additional multiyear simulation is 

planned with an enhanced parameterization of heterogeneous 

chemistry. One focus of the simulation will be the appraisal of 

the importance of background aerosols on the chemistry of the 

mid-latitude and tropical stratosphere.

Publications 

1. Grose, W. L.; Waters, J.; Eckman, R. S.; Froideveaux, L.; and 

Elson, L. S. "The Distribution of Ozone and Chlorine 

Monoxide in the Stratosphere: Comparisons of Microwave 

Limb Sounder Data with a Three-Dimensional Chemistry! 

Transport Model Simulation." Presented at the International 

Symposium on Middle Atmosphere Science, Kyoto, Japan, 

March 1992. 

2. Eckman, R. S.; Turner, R. E.; Blackshear, W. T.; Fairlie, 

T. D. A.; and Grose, W. T. "Some Aspects of the Interaction 

Between Chemical and Dynamical Processes Relating to the 

Antarctic Ozone Hole." To be published in Advances in 

Space Research, 1992. 

NASA Langley three-dimensional model calculates temperatures 

to describe global atmospheric dynamics.

65
Rotor—Stator Interaction in Turbomachines 

Karen L. Gundy-Burlet, Principal Investigator 

Co-Investigator: Aki I Rangwal Ia 

NASA Ames Research Center 

Research Objective 

To investigate unsteady rotor—stator interaction in single-stage 

and multistage turbomachines. 

Approach 

Unsteady three-dimensional Navier—Stokes zonal codes have 

been developed to investigate unsteady flows in single-stage 

(ROTOR-4 code) and multistage (STAGE-3 code) turbomachines. 

Accomplishment Description 

The nonlinear equations are advanced in time using a third-order 

accurate, implicit, upwind-biased algorithm. Body conforming 

"0" grids are used to accurately resolve the viscous effects 

associated with each airfoil. These grids are overlaid on sheared 

cartesian grids which resolve the flow field between blades. The 

cartesian grids are allowed to slip past each other, simulating the 

relative motion between rotors and stators. The STAGE-3 code 

has been applied to a 2.5-stage compressor configuration for 

which a large body of experimental data exists. Preliminary timeaveraged surface pressures are in good agreement with experimental data. The computation requires 10 megawords of main 

memory and approximately 600 Cray Y-MP hours. The ROTOR-4

code has been used to perform a fine-grid three-dimensional 

calculation of a turbine stage. The numerical results have 

indicated some remarkable secondary flow features that had not 

been seen in previous rotor—stator simulations. These results are 

in the process of being validated and will improve our understanding of the physical nature of flows through turbomachines. 

This computation requires 35 megawords of memory and 

approximately 400 Cray Y-MP hours. 

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 in order to design turbomachines 

that are light, compact, reliable, and efficient. 

Future Plans 

On completion of the turbine-loss calculation and the 2.5 stage 

compressor computation, the results will be compared with 

experimental data from test rigs. Future efforts will include a 

fine-grid 2.5-stage compressor computation. 

Preliminary instantaneous surface-pressure contours in a 2.5-stage compressor. 

66
Fluid and Structure Integration for Aerospace Applications 

Guru P. Guruswamy, Principal Investigator 

Co-Investigator: Shigeru Obayashi 

NASA Ames Research Center 

Research Objective 

To develop the numerical capability to couple Euler and 

Navier—Stokes solutions with finite-element structural equations 

to conduct aeroelastic analysis of complete aircraft. This capability 

is required for aerospace vehicles of national importance. 

Approach 

The three-dimensional Euler/Navier—Stokes equations directly 

coupled with the finite-element structural equations of motion 

are solved by time-accurate numerical schemes. 

Accomplishment Description 

ENSAERO, a code based on the Euler/Navier—Stokes equations, 

is used to simulate unsteady flows over flexible wing—body 

configurations. The code uses implicit finite-difference methods 

for aerodynamic calculations based on both central and upwind 

schemes. A capability of directly coupling the aerodynamic 

solutions from the Euler or Navier—Stokes equations with the 

structural equations for computing aeroelastic responses is 

incorporated in the code for wing—body configurations. This is 

the first time that such a computational capability has been 

developed in the area of computational fluid dynamics using the 

Navier—Stokes equations for wing—body configurations. Using

ENSAERO, vortical and transonic flows are computed over rigid 

and flexible wing—body configurations in unsteady motions, 

including vortex breakdown. The effect of body flexibility on the 

aeroelastic responses is demonstrated. A typical aeroelastic 

computation on a wing—body model undergoing a ramp motion 

from 0 to 15 degrees requires about 25 Cray Y-MP hours with a 

500,000 point grid and needs about 32 megawords of memory. 

Significance 

This project demonstrates the capability of computational 

methods to simulate aeroelasticity associated with complex 

vortical and transonic flows for improving the performance of 

an aircraft. It will also serve as a major stepping stone toward 

developing a multidisciplinary version of ENSAERO for fullaircraft analysis. 

Future Plans 

Research will be continued to compute unsteady flows over 

flexible wing—body configurations with vertical tails and control 

surfaces. ENSAERO will be extended as a multidisciplinary code 

to study fluid—structural interactions of complete aerospace 

vehicles. 

Unsteady Navier—Stokes computations on a wing—body configuration in ramp motion.

67
Airfoil Lift and Thrust Generation in Hover Mode 

Karl E. Gustafson, Principal Investigator 

Co-Investigator: Robert R. Leben 

University of Colorado, Boulder 

Research Objective 

To complete a full-parameter study of the relevant input 

variables for effect on computed coefficients of lift. The results 

will be compared to experimental coefficient-of-thrust measurements and will be used to tune airfoils for maximum lift and 

thrust. 

Approach 

We use a two-dimensional stream-function vorticity scheme with 

a grid-generation scheme that avoids truncation approximation 

in the far field. Stream function computations are performed by 

a multigrid solver and an alternating direction-implicit scheme 

that advances the vorticity values in time. A multigrid generation 

scheme provides the boundary-fitted coordinates. Coefficients 

of lift are calculated by a surface pressure integration. 

Accomplishment Description 

Higher Reynolds number (1,700) runs were performed for a full 

range of plunge and pitch amplitudes. A typical computation on 

a 65 x 65 grid required 5 megawords of memory and 1 CPU 

hour for four periods in a typical hovering mode. The data were

used to calculate coefficients of lift. Excellent correlation was 

found between the surface calculated coefficient of lift and the 

laboratory downstream measurements of coefficient of thrust. 

The accompanying figure shows a typical vorticity plot from 

which the coefficients were calculated. In all modes, the 

qualitative relationships between different pitch and plunge 

parameters were preserved. In modes similar to water treading, 

the quantitative similarities were remarkable. Coefficients of lift 

in excess of 5 were easily obtained. 

Significance 

The use of dynamic stall vortices for production of very high 

thrust are well demonstrated by hovering-mode flight. This 

hovering-mode model, based on observations of the flight of 

small birds, dragonflies, and other small insects, should be useful 

in understanding the characteristics of unsteady aerodynamics. 

Future Plans 

An algorithmic upgrade is under way to allow larger Reynolds 

numbers to be tested. Other unsteady hovering modes are under 

investigation. 

Vorticity plot for an oblique-jet hovering mode. Coefficients of lift of approximately 3 are obtained. 

68
Unsteady Cou nterrotation of Ducted Propfans 

Edward J. Hall, Principal Investigator 

General Motors Corporation, Allison Gas Turbine Division 

Research Objective 

To investigate the aerodynamics of ultra-high-bypass ducted-fan 

aircraft propulsion systems. Large-scale aerodynamic predictions 

provide a detailed description of the flow through the fan and 

the external flow about the fan cowl and engine nacelle. 

Approach 

A three-dimensional time-marching finite-volume Euler! 

Navier—Stokes analysis is used to predict steady and timedependent flows about ducted- and unducted-fan aircraft 

propulsion systems employing multiple blade rows. The analysis 

is based on a multiple-blocked grid solution utilizing a cylindrical coordinate system. The multiple-blocked grid discretization 

is applied to simulate the flow about the internal fan components 

(the rotor and stator) and the fan cowl. Individual grid blocks are 

numerically coupled through the mutual specification of 

boundary conditions along the interfaces joining adjacent 

blocks. Blade passages are discretized by a series of H-type 

grids. Grid blocks for neighboring blade rows share a common 

surface of revolution boundary, across which either circumferentially averaged (steady) or spatially reconstructed (unsteady) 

flow data are transferred. Ducted fans utilize an additional 

C-type mesh about the leading edge of the cowl, which 

improves the resolution of the flow predictions in this critical 

region. The scheme employs a number of steady-state convergence acceleration techniques, such as local time-stepping, 

residual smoothing, and multigrids. The residual smoothing and 

artificial dissipation terms are also carefully constructed to 

permit their application for time-accurate flow predictions. 

Accomplishment Description 

The calculations demonstrate the aerodynamic interactions 

occurring between adjacent blade rows, and illustrate the need 

for design analysis tools capable of predicting such phenomena. 

A typical time-dependent calculation required 30-80 hours 

on the Cray Y-MP and Cray-2 and 20-80 megawords of 

memory, depending on the grid density and number of blades 

in each row. 

Significance 

There is a substantial improvement in propulsive efficiency over 

current high-bypass ratio engines. However, ultra-high-bypass 

ratio engines impose stringent demands on the aerodynamic and 

structural integration of large, variable-pitch airfoils in minimum 

length, low-drag cowls. This code will contribute to improved 

efficiency and reliability for ultra-high-bypass fans, ultimately 

leading to the incorporation of this technology in the civil and 

military transport industry.

Future Plans 

The multiple-blocked Euler!Navier—Stokes scheme will be used 

to study endwall casing treatments that enhance fan stall margin 

and reduce noise. The code will be used to predict the timedependent flow field in the NASA Stage 37 test geometry. 

Parallel computing concepts will be explored to reduce the 

computation time for these calculations. 

Publication 

Hall, E. J. and Delaney, R. Investigation of Advanced 

Counterrotation Blade Configuration Concepts for High-Speed 

Turboprop Systems: Task V—Unsteady Counterrotation Ducted 

Propfan Analysis." NASA CR-i 87126, Jan. 1993. 

Predicted static-pressure contours for the NASA 1. 15 pressureratio fan.

69
(c) Kapingamarangi (154.8E, 1.1N) 

S.D. = 6.6 5.8 6.3 Corr 0.70 0.74 0.51

(d) Santa Cruz (90.3W, 0.8S) 

S.D. = 8.1 2.8 6.6 Corr = 0.30 0.91 -0.03 

30 

20 

10 

0 

-10 

-20 

-30 

30 

20 

10 

0 

-10 

-20 

-30

Satellite Data Assimilation and Ocean General Circulation Models 

David Halpern, Principal Investigator 

Co-Investigators: Y. Chao and C. Roberto Mechoso 

Jet Propulsion Laboratory/University of California, Los Angeles 

Research Objective 

To determine a method to extrapolate satellite measurements of 

sea-surface height throughout the ocean interior by combining 

the geodetic satellite (GEOSAT) altimeter data with the National 

Oceanic and Atmospheric Administration Geophysical Fluid 

Dynamics Laboratory (GFDL) ocean general circulation model 

(OGCM) of the tropical Pacific Ocean. 

Approach 

The first phase was completed last year when the OGCM was 

transferred to NASA Ames. The second phase consists of a 

comparison of monthly mean sea-surface heights measured at 

island stations, computed from GEOSAT, and simulated with 

the OGCM and the Florida State University ship-based surfacewind field. 

Accomplishment Description 

The accompanying figure shows that the quality of the data sets 

at the selected sites was similar. The average standard deviations 

of the sea-level measurements, GEOSAT data, and OGCM 

simulations were 9.8, 8.9, and 7.1 cm, respectively. Had daily 

wind stresses been used instead of monthly mean values, the 

variability of the OGCM simulations would have been higher. 

At sites located far from the equator, such as Guam (a) and 

Pago Pago (b), which are located at about 14'N north and south 

of the equator, the annual cycle is identified. Sea surface heights

at Guam and Pago Pago are lower in northern hemisphere 

winter and spring. The absence of an out-of-phase relationship 

between Guam and Pago Pago requires additional study. In 

1987, when an El Niño occurred, the sea surface was lower near 

the equator in the western Pacific (c) and higher in the eastern 

equatorial Pacific (d), which was a priori expected. The pattern 

reversed in 1988 during the intense La Nina. The average 

correlation coefficients between GEOSAT and OGCM data and 

the sea-level measurements were 0.64 and 0.71, respectively. 

At some sites the correlation was poor. The low correlation at 

Santa Cruz (d) might be caused by inadequate accounting of the 

amount of atmospheric water vapor, which lengthens the radaraltimeter signal travel time. 

Significance 

The differences between OGCM simulations and sea-level data 

were comparable to the differences between GEOSAT and sealevel measurements. The amount of improvement in the 

description of oceanographic subsurface parameters when 

GEOSAT data are assimilated into the OGCM remains to be 

determined. 

Future Plans 

We will determine a method to use the difference field between 

OGCM simulation and GEOSAT data in a reinitialization of 

the OGCM. 

(a) Guam (144.6E,13.4N) (b) Pago Pago (170.7W,14.3S) 

S.D. = 11.4 6.8 5.6 Corr = 0.76 0.85 0.82 S.D. = 5.5 5.2 3.2 Corr = 0.69 0.75 0.65 

Jan87 Apr Jul Oct Jan88 Apr Jul Oct Jan89 Jan87 Apr Jul Oct Jan88 Apr Jul Oct Jan89 

Time series of sea surface height anomalies at 4 locations; solid lines = sea level measurements, dashed lines = GEOSAT data, and 

dotted lines = the ocean general circulation model simulation with Florida State University winds. 

Fill
Atmospheric General Circulation Model Sensitivity to Sea Surface Temperature Fields 

David Halpern, Principal Investigator 

Co-Investigators: C. Roberto Mechoso and Robert Haskins 

Jet Propulsion Laboratory/University of California, Los Angeles 

Research Objective 

To determine the sensitivity of the University of California, 

Los Angeles (UCLA) atmospheric general circulation model 

(AGCM) to global monthly mean sea-surface temperature (SST) 

distributions compiled from sparsely and unevenly sampled 

observations recorded by ships (COADS) and from satellite 

measurements with very little aliasing (HIRS2/MSU). 

Approach 

UCLA AGCM simulations with HIR52/MSU and COADS SST 

data sets prescribed as boundary conditions will be compared 

with a control run involving only SST climatology. Ensembles of 

five 30-day mean fields are obtained from integrations performed 

in the perpetual-January mode. Results are presented as anomalies (departures of each ensemble-mean from control simulation). 

Accomplishment Description 

Large differences are found between the anomalies obtained 

using COADS and HIRS2/MSU SSTs, even in the northern 

hemisphere where the data are most similar. Patterns of 

geopotential height anomalies obtained using COADS SSTs 

resembles the first empirical orthogonal function (EOF 1) in the 

control simulation, while those obtained with HIR52JMSU SSTs 

resembles second empirical orthogonal functions. To examine 

the reasons for the results, three additional simulations are made 

with SST anomalies confined to regions where COADS SSTs are 

substantially warmer than HIRS2/MSU SSTs. As shown in the 

accompanying figure, regions correspond to warm pools in the 

northwest and northeast Pacific and the northwest Atlantic. 

Warm pools tend to produce positive geopotential height 

anomalies in the northeastern oceans through the effects of 

transient-eddy anomalies, thus shifting northward the corresponding storm tracks. Both Pacific warm pools produce largescale circulation anomalies that resemble those obtained using 

COADS SSTs; the Atlantic warm pool does not. The EOF 1 

pattern of the control simulation has a closer association with 

positive geopotential height anomalies over the northwest 

Pacific. In the northern hemisphere and the tropics, the 500 mb 

geopotential height anomalies obtained using COADS SST 

are more similar to the observed than those obtained using 

HIRS2/MSU SST. In the southern hemisphere, i-IIRS2/MSU SST 

generated heights are positively correlated.

Significance 

The impact on atmospheric circulation of uncertainties in current 

SST distributions for the world ocean can be as important as the 

SST anomalies. Caution is warranted when linking anomalous 

circulation and SST patterns in long-range prediction. 

Future Plans 

We will determine whether the conclusions from this study are 

similar for a perpetual-July mode. 

Publication 

Ose, T.; Mechoso, C. R.; and Halpern, D. "A Comparison 

Between General Circulation Model Simulations using Two SeaSurface Temperature Data Sets for January 1979." Journal of 

Climate, in revision.

Regional SST 

90N

L 

00

0 00 

ATco

AD C

EPCOADS (WPC OADS 

I I I 

180 120W 60W 0 60E 120E 180 

Longitude 

The regional sea-surface anomalies used in the Atlantic 

(ATCOADS), northeast Pacific (EPCQADS), and northwest 

Pacific (WPCQADS). Contour interval is 0.5° C. 

60N 

.5

3ON 

-j

EQ

VA
Transonic Analysis on Unstructured Grids 

David W. Halt, Principal Investigator 

McDonnell Aircraft Company/McDonnell Douglas Research Laboratories 

Research Objective 

To demonstrate the accuracy and efficiency of a characteristicbased unstructured-grid Euler solver for complex flow fields by 

comparison of numerical solutions against test data. 

Approach 

Roe's scheme is implemented on an unstructured grid with 

second-order accuracy using Frink's methodology. Grids are 

generated with Lohner's advancing-front grid generator using a 

background mesh to control desired grid clustering. A three 

stage explicit Runge—Kutta scheme with implicit residual 

smoothing is used to advance the solution in time. 

Accomplishment Description 

An unstructured mesh with a quarter million tetrahedra is used 

for the ONERA M6 wing transonic case at Mach 0.84 and 

3.06 degrees angle of attack. A lambda shock results on the 

upper surface and pressures compare well with experimental 

Pitot measurements. A more complex F-i 5 wind tunnel model 

configuration is chosen for comparison including wing, fuselage, 

tails, and inlet detail. An unstructured mesh with a quarter 

million tetrahedra is used for this configuration. In the accompanying figure, numerical pressures compare relatively well with 

Pitot measurements at Mach 0.9 and 5 degrees angle of attack. 

About 4 Cray-2 hours and 17 megawords of memory were used 

for the ONERA M6 wing case. About 10 Cray-2 hours and 

17 megawords of memory were used for the F-i 5 case. 

Significance 

This method is efficient for analyzing complex configurations 

with relatively accurate comparisons made from existing 

test data.

F- 15 surface color Mach contours; Mach = 0.9 and angle of 

attack =5 degrees.

M.0 0.9. .ngl. 59 8tI.oS 5 dog 

CP -:t - Nlm..fl05l ,00oIIO 

.50 5 Wind h,nn59d.t8 

Future Plans 

Store calculations with an F-i 8 wind tunnel model are planned 

and will be compared to existing force and moment wind tunnel 

data. Grid adaptation using a Delaunay grid-generation method 

will be investigated. Future plans include compact high-order 

characteristic-based schemes and Navier—Stokes implementation.

8

77% 

-50 

C92 

.75(l 1 1 1 ", , 

0 .2 .4 .6 .9 5.0

s/c (traction chord)

.4 .6 .8 1.0 

c/c (11.51105 chord) 

Wing pressure-coefficient comparisons with test data. 

72
2.5 5.0 7.5 10.0 12.5 

15.0 

12.5 

E U

10.0 

C 0 

Iii 7.5 

S

5.0 S V 

U .E 2.5

Radiative Structure in Aerobrake Shock Layers 

H. Harris Hamilton, Principal Investigator 

Co-Investigator: Robert B. Greendyke 

NASA Langley Research CenterNiGYAN, Inc. 

Research Objective 

To investigate the influence of kinetic modeling assumptions 

from several flow-field solution methods on the noncoupled 

radiative structure of aerobrake shock layers. Analysis of the 

excited-state population distribution will provide insight into the 

effect of macroscopic-flow boundary conditions upon radiative 

heat-transfer studies. 

Approach 

Flow-field solutions calculated with LAURA code were obtained 

from a related study of electron-number densities. The LAURA 

code used was an axisymmetric Navier–Stokes solution method 

coupling 11 species-continuity equations and both a total-energy 

equation and a vibrational-electronic energy equation. The 

results were used in the NEQAIR line-by-line nonequilibrium 

radiation code and the flow-field solution stagnation streamline. 

The NEQAIR code calculates the emission and absorption 

coefficients for the nitrogen and oxygen species used to calculate the absorbing incident radiative flux in the shock layer. The 

individual radiative processes were separated from the total flux 

to provide process profiles from the shock to the wall. 

Accomplishment Description 

The NEQAIR code was applied to flow fields having two 

ionization-potential energy assumptions, three commonly used 

chemical rate models, and three different values for the limiting 

cross section for vibrational excitation. The radiative structure 

was analyzed for the LAURA-code baseline kinetic model 

at several points along a typical Aeroassist Flight Experiment 

trajectory. Two radiative processes were found to be significant 

in this study—radiation resulting from atomic nitrogen 

bound–bound transitions and ionized diatomic nitrogen's 

first negative band. Considering the high computational cost of 

the NEQAIR code, substantial computational savings could be 

realized by using only these two processes. Flow-field assumptions exerted a considerable effect on the total radiative flux at 

the wall. Occasionally, the modification of the chemical rate 

model used had a factor of 3 difference on the calculated 

radiative heat transfer to the wall (shown in the accompanying 

figure). The effect of temperature profiles, and the assumption 

that electron temperature could be assumed equal to vibrational 

temperature, was highly significant to the overall radiative 

structure. 

Significance 

Valuable insight into the the effects of flow-field solution 

methods upon the microscopic processes involved in hypersonic 

nonequilibrium radiation was provided. The results directed us 

to areas of further flow-field solution method adaptations for the 

more accurate calculation of radiative heat transfer.

Future Plans 

The accuracy of wall-boundary conditions will be examined 

related to boundary-layer radiative absorption and total flux to 

the wall. Similar methodology will be applied to the FIRE II 

vehicle data. 

Publications 

1. Greendyke, R. B. "A Parametric Analysis of Radiative 

Structure in Aerobrake Shock Layers." AIAA Paper 92-2970, 

23rd AIAA Plasmadynamics and Lasers Conference, 

Nashville, TN, July 1992. 

2. Greendyke, R. B.; Gnoffo, P. A.; and Lawrence, R. W. 

"Electron Number Density Profiles for the Aeroassist Flight 

Experiment." AIAA Paper 92-0804, 30th AIAA Aerospace 

Sciences Meeting, Reno, NV, Jan. 1992. 

15.0 [(a) 0 Atomic nitrogen • N2 (LBH) 

Atomic oxygen A NO () 

12.5 I- v N2+(-1) NO (y) 

E U 1 0 N2 (+1) • 02 (S-R) 

10.0 0N2(+2) •Total 

F 

5.0 

2.5

Distance from wall (cm) 

Incident radiative flux resulting from (a) the Kang and Dunn 

chemical-rate model and (b) Park's 1990 chemical-rate model.

73 
Incorporation of a Three-Dimensional Multistage Viscous Code into a Compressor 

Design System 

Jeff L. Hansen, Principal investigator 

Co-Investigator: John Adamczyk 

General Motors Corporation, Allison Gas Turbine Division/NASA Lewis Research Center 

Research Objective 

Emphasis on improved fuel efficiency, reduced weight, and 

lower life-cycle costs for gas turbine engines has produced 

aggressive compression system design, performance, and 

operability goals. To meet these goals, there is a critical need to 

incorporate a three-dimensional Navier—Stokes multistage flow 

model in the design process. The goal of this research is to use 

such a model to guide the design of a high-performance, highly 

loaded axial-flow multistage compressor. 

Approach 

The aerodynamic design of a two-stage compressor was 

completed with a NASA-developed three-dimensional multistage 

viscous code (V-STAGE) with a traditional design system. The 

multistage code solves the three-dimensional average-passage 

Navier—Stokes equations, which are obtained from the Reynoldsaveraged form of the equations. The resulting system of equations contain additional source terms to account for the presence 

of neighboring blade-rows. After parametric optimization with a 

through-flow analysis, the detailed airfoil design was achieved 

using an iterative loop between the through-flow analysis and a 

three-dimensional viscous isolated blade-row analysis and 

continued until a geometry that yielded acceptable aerodynamics was reached. An additional iteration loop involving the 

multistage code was used to make design modifications as a 

result of multistage effects on stage mismatch, blade-row 

interactions, mixing, and stall characteristics. 

Accomplishment Description 

The code was implemented in the design process for a multistage compressor. Many design changes that normally would not 

have been made until after rig testing were made as a result of 

using this code. Extensive design modifications were made to 

eliminate suction-surface boundary-layer separation and 

improper exit-flow angles caused by multistage influences. The 

multistage code revealed a stage mismatch with the second rotor 

overflowing. The second rotor would unstart as the compressor 

was loaded above the design pressure ratio. Modifications to the 

rotors were made and an acceptable design was reached. The 

final design for this compressor was the result of extensive use 

of advanced computational fluid dynamics codes requiring 

approximately 300 Cray-2 hours.

Significance 

In the design process, it is important to incorporate a threedimensional Navier—Stokes multistage flow model with the 

capability to model the complex interactive flow phenomena in 

multistage compressors. This allows the design of efficient, 

lightweight, and reliable compression systems with reduced 

development time and costs. 

Future Plans 

Near term plans for this project are to complete the assembly of 

and test the two-stage compressor. This experimental data will 

be used to further calibrate the multistage code. A follow-on 

project may involve using this code to match a centrifugal 

compressor to the two-stage compressor to obtain a compact 

high-pressure-ratio compression system. 

Three-dimensional multistage Navier—Stokes analysis fully 

incorporated in the design process. 

74
N-Facbrs 

(a) 8 (c)

(d)

Swept-Wing Leading-Edge Transition 

Julius E. Harris, Principal Investigator 

Co-Investigator: Venkit Iyer 

NASA Langley Research CenterNiGYAN Inc. 

Research Objective 

To determine the effects on boundary-layer stability and 

transition due to single-parameter variations of the angle of 

attack, boundary-layer suction, wall cooling, and free-stream 

Reynolds number on the leading edge of a highly swept supersonic wing. 

Approach 

A 77.1 degree swept-leading-edge model undergoing tests at 

the NASA Langley Mach 3.5 Supersonic Low-Turbulence Pilot 

Tunnel is the geometry considered. The mean flow on this wing 

is obtained by solving the Navier—Stokes equations. The solution 

profiles are interfaced to a three-dimensional compressible 

version of the linear stability analysis code COSAL. The N-factors 

calculated are used as guidelines for predicting transition trends 

when the parameters are varied. 

Accomplishment Description 

First, the N-factors and transition front corresponding to the 

nominal conditions of Mach 3.5, 0 degrees angle of attack, 

Re = 2.4 million/foot, adiabatic wall, and zero suction were 

obtained. Results were then obtained with changes in angle of 

attack (1 .5, 3.0, and 4.5 degrees). In spite of the larger cross 

flow, the transition front did not move upstream significantly. 

Results were obtained with two different suction distributions

that had a strong stabilizing influence. Two cases of wall cooling 

(TJT.. = 2.0 and 2.6) were run, verifying the strong stabilization 

potential of wall cooling. Finally, results were obtained with 

Re values of 1.2, 4.0 and 6.0 million/foot. The flow was completely stable at the lower value. Significant upstream movement 

of the transition front was seen for Re. = 6.0 million/foot. 

Typical runs required approximately 30 megawords of memory 

and 2 CPU hours. Some fine-grid runs tripled the requirements. 

Significance 

Results of this study are being used to determine the test 

envelope, instrumentation, and measurement requirements for 

the experimental study. The influence of these parameters on 

transition is important in the design optimization of natural or 

forced laminar-flow control on a swept supersonic wing. 

Future Plans 

We will extend the calculations with multiparameter variations 

corresponding to actual test conditions and evaluate the results 

with experimental data. 

Publication 

lyer, V.; Spall, R. E.; and Dagenhart, I. R. "Computational Study 

of Transition Front on a Swept-Wing Leading-Edge Model." AIAA 

Paper 92-2630, June 1992. 

(a) Swept leading edge N-factors and transition front, M, = 3.5, a = 0 deg. Re,, = 2.4 x 106, wall cooling and suction off. (b) Streamlines for a =0 deg case. (c) N-factors and transition front with suction (q5 = 0.0004). (d) N-factors with Re,,,, = 1.2 x 106.

75
40 Wall radiative flux 

E

20 

0

Tangent slab 

30 ------------ MDA 

10 

Radiation Transport around Axisymmetric Blunt-Body Vehicles 

Lin C. Hartung, Principal Investigator 

NASA Langley Research Center 

Research Objective 

To develop a practical method to compute radiation transport 

in complex flow fields without invoking the tangent slab 

approximation. 

Approach 

A modified differential approximation (MDA) employs the 

P-i approximation to model the radiation transport in a medium 

without accounting for wall emission and reflection. This 

reduces the governing integro-differential equations to coupled 

partial differential equations. A numerical finite-volume solution 

method is developed to obtain the radiative-flux incident on the 

wall and the radiative-flux divergence in the medium. 

Mars return condition:

V 12km/s

H = 80 km

R = 1 m 

60 sphere cone

Accomplishment Description 

The MDA method has been applied to two forebody aerobrake 

flow fields. Comparisons with tangent slab results suggest the 

MDA method is valid. The computer time required by the 

method is within reason. 

Significance 

Radiative heating is a concern for large-scale or large-velocity 

missions, particularly if unshielded payloads are placed behind 

a heat shield. This method offers the potential to predict the 

radiative heat load on such vehicles. 

Future Plans 

We will continue to improve the efficiency and robustness of the 

method and apply it to wake flows. 

5o Tangent slab flux divergence

qFI 

-100 

0 50 100 150 200 250

S (cm)

5O flux divergence

V.q 

_100L. - 

0 50 100 150 200

x (cm) 

Mars return conditions; V = 12 km/sec. H =80 km, R = 1 m, 60 degree sphere cone. 

76
.25 .50 

Y(in.)

.25 .50 

V (In.) 

.25 .50 

V (In.)

.25 .50 .75 

V(in.) 

.150 

.125 

.100 

.025 

.150 

.125 

.100 

H

.075 

.050 

.025 

.100 

.075 

H

.025 

.150 

.125 

.100 

.075 

.050 

.025

New Approach for Transitional-Flow Modeling 

H. A. Hassan, Principal Investigator 

Co-Investigators: T. Wayne Young and Eric S. Warren 

North Carolina State University 

Research Objective 

To develop models similar to available turbulence models for 

studying transitional flows. 

Approach 

If y represents the fraction of time the flow at a given point is 

turbulent, then transitional stresses can be represented as 

(U.J;) = y(uJ1) +( 1— y)(ucu;) + y(l - 

AU i = U — Ui, 

where subscripts r, t and £ indicate, respectively, transitional, 

turbulent, and laminar or non-turbulent. Reynolds-averaged 

equations using a one-equation model and a stress model were 

used to calculate the turbulent components. A number of 

procedures were used to calculate the contribution of the large 

structures (the term involving 

Accomplishment Description 

Emphasis is placed on predicting the measurements of 

Schubauer and Klebanoff, which consist of velocity and streamwise intensities. A procedure was developed for the incorporation of some of the linear stability theory results in order to 

estimate the contribution of the laminar or non-turbulent terms. 

The results showed that the non-turbulent and large-eddy 

components have a considerable influence on the streamwise 

intensities, but little influence on the shear stress. Moreover, 

predictions of the one-equation model were as good as those 

obtained for the stress model. The accompanying figure shows a 

comparison between theory and experiment. Each job requires 

about 10 megawords of memory and 2.5 CPU hours to converge. 

Significance 

Recent research on the National Aero-Space Plane identified the 

need to develop computational tools for the prediction of highspeed transitional flows. Understanding transitional flows has a 

major impact on the feasibility of this plane. 

Future Plans 

Work is focusing on pressure gradient and Mach number 

effects. 

Stream wise turbulent intensifies.

77
Hot-Gas Ingestion by a Short Takeoff and Vertical Landing Aircraft in Ground Proximity 

James D. Holdeman, Principal Investigator 

Co-Investigators: David M. Fricker and S. Pratap Vanka 

NASA Lewis Research Center/University of Illinois, Urbana/Champaign 

Research Objective 

To simulate the three-dimensional turbulent flow around a short 

takeoff and vertical landing (STOVL) aircraft in ground proximity. 

Studying the gas environment is part of the effort to understand 

hot-gas ingestion for advanced STOVL aircraft. 

Approach 

A state-of-the-art, multigrid, three-dimensional, Reynoldsaveraged, Navier—Stokes flow solver with a two-equation 

turbulence model is used to predict the hot-gas environment 

around a STOVL aircraft operating in ground proximity. 

Accomplishment Description 

Work included modifying the three-dimensional multigrid 

computational fluid dynamics (CFD) code and running more 

parametric calculations of test cases. These calculations included 

variations of ground proximity, head-wind speed, turbulence 

intensity, and thrust-splay angle. Over 30 cases, each involving 

over 200,000 grid points, were computed. A typical test case 

required about 1.5 Cray Y-MP hours and 16 megawords of 

memory. Qualitatively, the results compare well with experimental data. A number of interesting flow phenomena have been 

noticed in the numerical results. Other work included revising 

and debugging the CFD code from a cartesian geometry to a 

curvilinear geometry. 

Significance 

The problem of hot-gas ingestion for STOVL aircraft has typically 

been solved by a combination of experience and trial and error. 

With the advent of highly efficient flow solvers, CFD is becoming a viable engineering tool in the study of hot-gas ingestion 

problems and STOVL aircraft design. 

Future Plans 

Testing of the curvilinear version of the multigrid flow solver will 

be completed and the code will be used to study the hot-gas 

environment of an advanced STOVL design. This will be an 

improvement over the generic cartesian geometry used before. 

The studies will include parametric variations similar to those 

done with the cartesian code, but will involve a much more 

realistic aircraft geometry to investigate the effects of aircraft 

shape.

• :... .-.. 

Overhead view of ground temperature contours, select particle traces, and a three-dimensional isotherm. The lighter area reveals 

widespread influence of the hot exhaust gases, while the particle traces (red and blue lines) and the isotherm (light red surface) 

indicate the deflection of the ambient flow. 

78
Full Navier—Stokes Analysis of a Three-Dimensional Scramjet Inlet 

Yeu-Chuan Hsia, Principal Investigator 

Co-Investigators: Endwell Daso and Sukumar Chakravarthy 

Rockwell International, Rocketdyne Division/Rockwell International Science Center 

Research Objective 

To analyze the flow structures in practical high-speed inlets and 

to determine their effects on the inlet performance. The present 

work is to validate the computational fluid dynamics code on 

a three-dimensional scramjet inlet and to evaluate the applicability of the full Navier—Stokes solution on inlet performance 

predictions. 

Approach 

The unified solution algorithm code was used to compute the 

inlet flow field. The code solves full or Reynolds-averaged 

Navier—Stokes equations based on finite-volume discretization 

using the total-variation-diminishing formulation. Perfect gas 

relation was used as the thermodynamic property. The turbulence effect was simulated by the Baldwin—Lomax algebraic 

turbulence model. The code also possessed a multi-zone grid 

bookkeeping capability. 

Accomplishment Description 

A fixed-geometry sidewall compression inlet was computed with 

a Mach 5 inflow. For 0 angle of attack and yaw, double symmetry was assumed and only one quarter of the flow path was 

computed. In order to simulate the inlet spillage properly, the 

external flow region was also computed. A grid consisting of 

10 zones with 242,800 nodes was used. The solutions were 

obtained by a time marching method. A laminar-flow solution 

was computed first. The turbulent-flow solution was computed 

with a boundary layer transition location estimated by comparing the laminar solution and test data. Each solution took about 

60 Cray Y-MP hours and 16 megawords of memory. 

Significance 

Both the laminar- and turbulent-flow solutions closely predicted 

the overall inlet performance in terms of air-capture ratio and 

total-pressure recovery. However, comparing the flow-field data, 

only the turbulent-flow solution showed good agreement. The 

laminar-flow solution had boundary-layer separations that might 

lead to inlet unstart. Hence, the inlet functionality should be 

assessed by not only the performance parameters, but also the 

flow structures. 

Future Plans 

More validation computations will be done on inlets with heattransfer and skin-friction data. Time-accurate computations will 

be carried out to simulate the inlet unstart.

16

It 

12 16 20 

Mach 5 computations of the sidewall compression inlet. 

12 

CL 8

79
Ni! 

CP 0 

High Angle-of-Attack 

Carl T. Hsieh, Principal Investigator 

Naval Surface Warfare Center

Missile Aerodynamics 

Research Objective 

To evaluate numerical simulations of vortical flow over an ogive 

cylinder and to numerically investigate incompressible, threedimensional, laminar-flow structure about a hemisphere cylinder 

at large incidences. 

Approach 

The CFL3D code was used to simulate supersonic flow over an 

ogive cylinder at a = 18 degrees and Mach 3.5, and incompressible laminar flow over a hemisphere cylinder at a = 10, 30, and 

50 degrees. Solutions were compared to experimental data and 

observations. 

Accomplishment Description 

The CFL3D code has been adapted to simulate both laminar and 

turbulent flows over an ogive cylinder with a Reynolds number 

of 1.23 million based on diameter. A C-grid with points of 

91 x 68 x 75 was used. The turbulence model used was the 

Baldwin—Lomax model modified by Degani and Schiff for 

vortical flow. Tests of the effects of grid sensitivity were conducted and an optimum grid distribution was selected. Comparisons of cross-flow surface pressure at four axial stations, Pitot 

pressure along the radial lines, and Pitot pressure contours at the 

axial station ctD = 11.5 (see accompanying figure), where x is 

the distance from the nosetip and D is the cylinder diameter, 

were presented. Good qualitative prediction was found for flow

features such as the leeside vortex, the shear layer emanating 

from the separation point, the bow shock, the cross-flow shock, 

and the shock atop the vortex at far downstream. The laminar 

solution agrees better with the experiment; however, the point of 

separation predicted is farther leeward than experiment. For the 

incompressible laminar flow over the hemisphere cylinder at 

incidences, a C-grid with points of 101 x 68 x 75 was used. 

Converged solutions were obtained, but analysis and comparison with experiments have not yet been performed. Computation requires about 30 Cray-2 hours and 24 megawords of 

memory. 

Significance 

Above an incidence of 5-10 degrees, the flow in the leeside of a 

circular body separates to form vortices. The interaction of the 

leeside vortical flow with control surfaces contributes to the 

highly nonlinear aerodynamics exhibited by these missiles. 

Prediction of vortical flow is important in the design of a new 

generation of highly maneuverable missiles. 

Future Plans 

The CFL3D code will be used to compare two flux-splitting 

schemes of Roe and Van Leer for steady flow, and study unsteady/ 

transient flow about a pitching missile body. Computation of a 

missile with fins will also be performed. 

(a) (b) (c)

Pitot pressure contours at xJD = 11.5; (a) experiment, (b) laminar solution, (c) turbulent solution. 

80
Submarine-Appendage Design and Turbulence Modeling 

Thomas T. Huang, Principal Investigator 

Co-Investigators: Michael Griffin, Jeff Tsai, William Smith, and Dane Hendrix 

Naval Surface Warfare Center/Jason Associates/Scientific Research Associates, Inc./MCAT Institute 

Research Objective 

To design a small appendage on the underside of a submarine 

bow using an incompressible Reynolds-averaged Navier–Stokes 

(RANS) flow solver for the flow analyses. The appendage is 

required to produce a minimum variation in pressure and to 

induce a minimum longitudinal vorticity in high-Reynoldsnumber flow conditions. The development of a more accurate 

algebraic turbulence model in areas of separated flow will also 

be addressed. 

Approach 

Given the original design and various constraints, a small 

appendage on the underside of a submarine bow has been 

redesigned using the incompressible RANS flow-solver, IFLOW 

Accomplishment Description 

The calculated pressure coefficient on the hull in the vicinity 

of the appendage and the longitudinal vorticity component 

downstream from the appendage for both the original and the 

modified designs are shown in the accompanying figure. The 

modified design produces a significantly smaller variation in 

pressure on the appendage and induces less longitudinal 

vorticity. Typical flow analyses required 10 megawords of 

memory and 30 Cray Y-MP minutes. Over 100 different design 

modifications were considered and analyzed. This is an important demonstration of the capability of the current RANS flow

solver to aid in the design of critical components of naval ships 

and weapons. In additional research, a new algebraic turbulence 

model has been developed for the prediction of three-dimensional 

separated flows. Preliminary results indicate that the new turbulence model predicts flow-field quantities for an unappended 

submarine with greater accuracy than previously developed 

models. Typical flow analyses required 5 megawords of memory 

and 10 Cray Y-MP minutes. 

Significance 

Submarine design has relied on a combination of classical 

analysis, potential-flow computation, and viscous–inviscid 

interactive computation. The current RANS solver provides fast 

and accurate flow-field analyses and previously unobtainable 

detailed information about longitudinal vortices, flow nonuniformity at the propeller plane, forces, and moments at large 

angle of attack during maneuvering. This solver will make it 

possible to produce inexpensive and efficient submarine designs. 

Future Plans 

We will further develop turbulence models for submarines and 

complete validation of IFLOW and additional applications of 

IFLOW to design components critical to naval fleets. A method 

for computing three-dimensional flows about an arbitrary ship 

form advancing in regular waves is being developed. Calculations to test and validate this method will be performed. 

ZIL

MODIFIED DESIGN 

-0.02 

MODIFIED ORIGINAL -5.00 

Cp 

DESIGN DESIGN

-20.0 

-0.04

?'' 

• 1.00 0.00 0.02 0.04 y/L 

• 0.10

I-0.8Ot z/L 

U -1 .70k a -0.02

ORI(4L DESI 

—. — - 4 -5.00 

- 

2 -15.0 

(II

-0.04 

0.00 0.02 0.04 ylL 

LOCATION OF APPENDAGE ON HULL LONGITUDINAL VORT1CITY AT xIL = 05 

Design of a submarine appendage using the Reynolds-averaged Navier–Stokes flow solver.

81
Generic National Aero-Space Plane Fuselage Configuration Study 

Lawrence D. Huebner, Principal Investigator 

Co-Investigator: Kenneth E. Tatum 

NASA Langley Research Center/Lockheed Engineering and Sciences Company 

Research Objective 

To develop the capability to predict the three-dimensional flow 

field about a generic National Aero-Space Plane (NASP) fuselage 

and engine module under simulated power conditions and to 

determine the three-dimensional effect of using a tetrafluoromethane (Cl4) and Argon (Ar) gas mixture as the exhaust 

simulant instead of air. 

Approach 

For the two exhaust-gas simulations (air or CF 4—Ar), threedimensional viscous flow fields were computed by solving the 

parabolized Navier—Stokes equations using the General Aerodynamic Simulation Program. The resulting flow-field quantities, 

surface quantities, and integrated forces and moments were 

examined to quantify differences between the exhaust-gas 

simulations. 

Accomplishment Description 

Computational solutions for the fuselage and engine module 

were obtained for the two types of exhaust. The first solution was 

with both the external and internal flow represented by air 

(perfect gas) and the second with the external flow represented 

by a mixture of nitrogen and oxygen (simulating air) and the 

internal exhaust simulant with CF 4—Ar. All four gas species 

were present in the aftbody zone. Solutions were computed at 

M,, = 6.0, Re, = 1.0 x 1 06/ft, a = 0 degrees, and static nozzle 

pressure ratios of 135 and 232. The accompanying figure shov 

surface pressure contours and flow-field Mach number contours 

on the model centerline and downstream grid planes. The figure 

inset shows a comparison of plume expansion at the end of the 

model for the two exhaust gases. For both cases, the plume has 

expanded beyond the edge of the body, indicating that the 

plume will likely impact lifting and control surfaces that are on 

the fuselage. The CF4—Ar plume expands farther from the aftbody 

than the air plume, although some of the air plume expands to 

Mach 8, well beyond the free-stream Mach number. Typical-air 

parabolized Navier—Stokes (PNS) solutions required about 

9 megawords of memory and 8.3 Cray Y-MP hours, while the 

multiple-species PNS solutions required 12 megawords of 

memory and 9.2 Cray Y-MP hours.

Significance 

There are differences in aftbody flow fields. Also, regardless of 

the exhaust gas modeled, the plume extends beyond the edge of 

the aftbody and will likely impact the lifting and control surfaces 

of a NASP vehicle. 

Future Plans 

Following minor grid improvements on the existing computational model, cases will be computed on the configuration with 

the wings installed to confirm the plume effects on lifting and 

control surfaces. Additional solutions will be obtained by 

varying the inlet representation and free-stream and internal (jet) 

conditions. 

Computational solution of a National Aero-Space Plane fuselage 

configuration. Tetrafluoromethane and argon exhaust solutions 

computed at = 6.0, Re = 1.0x 10611t, a = 0 degrees, and 

static nozzle pressure ratio = 232, show surface pressures 

(blue = 0.05 psi, red = 2.50 psi) and flow-field Mach number 

contours (blue = 0.0, red = 6.0). The inset shows streamwise 

velocity-component contour comparisons at the fuselage trailing 

edge for the two types of exhaust simulations. 

82
Hypersonic Scramjet and Detonation Flows 

Joseph W. Humphrey, Principal Investigator 

Co-Investigators: Darrell W. Pepper, Thomas H. Sobota, Frank P. Brueckner, Barry Dyne, and Juan C. Heinrich 

Advanced Projects Research, Inc./University of Arizona/NASA Langley Research Center 

Research Objective 

To develop the capability to analyze and design a scramaccelerator technology demonstrator. A scramaccelerator is a 

projectile launching device that uses scramjet propulsion and 

oblique detonation to accelerate projectiles. Hypersonic 

chemically reacting flow is encountered in the high Mach 

number (5-16) regimes. 

Approach 

A finite-element Navier—Stokes solver in both two and three 

dimensions specifically oriented toward solving compressible 

high-Mach-number scramjet-type flow fields was developed to 

incorporate finite-rate chemistry. 

Accomplishment Description 

A two- and three-dimensional finite-element model for viscous 

high-speed compressible flow is being applied to the scramaccelerator configuration. The scheme is based on a shockcapturing Petrov—Galerkin formulation where a different 

anisotropic diffusion is added for each conservation equation. 

A precise criterion is available to choose the direction and the 

magnitude of the added diffusion for superior stability and 

accuracy. The semi-discrete Petrov—Galerki n approximation has 

been implemented in conjunction with a lumped-mass secondorder Runge—Kutta time-integration scheme. It is based on 

isoparametric bilinear elements in space and a weightedresiduals formulation where the weighting functions for the 

convective terms are different for each transport equation. The 

finite-rate chemistry model developed by D. T. Pratt at the 

University of Washington has been incorporated into the flow 

solver. Reacting-flow results are obtained by operator splitting 

or by a method of fractional steps. Time-accurate calculations 

require an iteration at each time step. This approach allows 

independent development of flow and chemistry solvers, using 

methods optimal for each system. A typical axisymmetric 

inviscid reacting-flow simulation requires approximately 

2-3 Cray-2 hours and 1 megaword of memory. 

Significance 

NASA Langley hopes to develop a hypersonic ballistic range 

for ground-based free-flight hypersonic testing. The scramaccelerator propulsion technology is also a direct benefit to the 

National Aero-Space Plane scramjet efforts. 

Future Plans 

Plans include implementation of adaptive gridding, continued 

parametric studies of the scramaccelerator, dynamic design 

iteration of hypersonic nozzles, continued optimization through 

vectorization and parallelization, three-dimensional studies 

of spinning projectiles, fin effects and stabilization, and 

sabot—projectile separations. 

(b) 

Mach 8 scramaccelerator simulation: (a) pressure contours around an entire projectile, and (b) detail of a multidimensional oblique 

detonation wave and flow accommodation in flow annulus.

83
Coherent-Structure Interactions with Turbulence 

Fazle Hussain, Principal Investigator 

Co-Investigator: Mogens V. Melander 

University of Houston/Southern Methodist University 

Research Objective 

To investigate the core dynamics of an isolated coherent 

structure and its interactions with turbulence by simulating 

idealized configurations. 

Approach 

Direct numerical simulations of the three-dimensional timedependent incompressible Navier—Stokes equations are utilized. 

The simulations use highly accurate and efficient pseudospectral 

Fourier methods. Critical parameters such as Reynolds number, 

amplitude of core nonuniformity, and intensity of background 

turbulence are varied. 

Accomplishment Description 

Using an axisymmetric vortex tube with a nonuniform cross 

section, we have shown that core dynamics lead to axial waves. 

The frequency of the waves decreases as the Reynolds number 

increases, reaching a finite value in the inviscid limit and the 

wave amplitude decays slowly for high Reynolds numbers. The 

little-used, powerful mathematical approach of complex helical 

wave decomposition explains core dynamics better than 

coupling between meridional flow and swirl. Simulations of a 

vortex tube with a nonuniform cross section in the presence of 

background turbulence showed that, contrary to the usual belief, 

small scales can also cause instability of the large-scale structures, which energizes and sustains the small scales. Coherent 

structures are helical and the helicity integral can become 

nonzero during simple viscous decay. These runs typically 

required 4-8 Cray-2 hours and 10 megawords of memory. 

Significance 

Results show that core dynamics must always be modeled and 

that simulations that neglect them are suspect. We propose a 

new physical model of cascade that questions the concept of 

local isotropy. 

Future Plans 

We will validate our conceptual model of cascade through more 

high-resolution simulations. We have developed a code based 

on complex helical wave decomposition that will be used to 

investigate the dynamics of vortex rings with swirl and vortex 

breakdown. 

Publications 

1. Melander, M. V. and Hussain, F. "Coherent Structure 

Dynamics: Interaction between Large and Fine Scales." 

8th Symp. Turb. Shear Flow. Munich, GE, Sept. 1991. 

2. Melander, M. V.; Hussain, F.; and Basu, A. "Breakdown of a 

Circular Jet into Turbulence." 8th Symp. Turb. Shear Flow. 

Munich, GE, Sept. 1991.

Hussain, F. and Melander, M. V. "Understanding Turbulence 

via Vortex Dynamics." In Studies in Turbulence, eds. Gatski 

et al. Springer Verlag, 1991. 

Hussain, F. and Melander, M. V. "New Aspects of Vortex 

Dynamics: Helical Waves, Core Dynamics, Viscous Helicity 

Generation, and Interaction with Turbulence." In Topological 

Aspects of the Dynamics of Fluids and Plasma, eds. H. K. 

Moffatt et al. Accepted by Kluwer PubI., 1992. 

Organization of small scales into spirals around a nonuniform 

cross-section vortex tube. 

84
Chemically Reacting Free-Shear Flows 

Fazle Hussain, Principal Investigator 

Co-Investigators: Ralph Metcalfe and Fernando Grinstein 

University of Houston/Naval Research Laboratory 

Research Objective 

To investigate the role of coherent structures formed by the 

primary and secondary instabilities (spanwise "rolls" and 

streamwise counter-rotating "ribs") in chemically reacting 

turbulent free-shear flows and to determine how they effect 

important phenomena such as species transport, reaction rates, 

and distribution of product. 

Approach 

Using spectral numerical methods, direct simulations are 

employed to solve the time-dependent incompressible 

Navier—Stokes equations in two and three spatial dimensions 

coupled with multiple reaction—diffusion equations governing 

the chemical reactions. Important parameters such as initial 

perturbations and convective and diffusive Damkohler numbers 

are varied to determine the effect on the reaction. 

Accomplishment Description 

New simulations have extended the range of parameters under 

investigation, including the addition of temperature-dependent 

reaction rates. We have shown that the presence of ribs can, 

depending upon the Schmidt number (Sc), enhance or retard 

chemical reactions. At a Sc = 0.70 the presence of ribs enhances 

the generation of product by increasing the surface area of the 

reaction zone, whereas at Sc = 0.07 the reaction is diffusiondominated and flame shortening occurs in the ribs, thereby 

reducing the reaction zone. Simulations performed at higher 

Schmidt numbers (up to 5.0) show that reactants in the paired 

cores are segregated by a thin product layer, thus preventing 

interdiffusion of reactants. This mechanism is different from 

flame shortening for Sc < 1 where there is an exhaustion of 

reactants within the rolls. Temperature-dependent simulations 

have shown that ribs sustain reactions while the braid region is 

unable to keep pace with the influx of reactants and is cooled 

nearly to the activation temperature. Typical high Schmidt 

number three-dimensional runs required 25 Cray Y-MP hours 

and 80 megawords of memory to reach pairing. 

Significance 

Combustion efficiency can be optimized through the artificial 

excitation of coherent structures. The role of ribs in generating 

product is important. At higher Schmidt numbers, the existence 

of another reaction-controlling mechanism has been demonstrated. In the temperature-dependent case, the presence of 

reaction within the ribs may serve as a mechanism to ignite 

mixed reactants at later times. 

Future Plans 

We plan to further investigate the role of coherent structures in 

reacting flows, including the importance of the phase of the 

unstable modes and oblique modes. We will continue to 

investigate temperature-dependent reactions and heat release 

and compressibility effects.

Publication 

Metcalfe, R. W.; Hussain, F.; and Park, K. H. "Role of Rolls and 

Ribs in Reacting Mixing Layers." 8th Symp. Turb. Shear Flow. 

Munich, GE, Sept. 1991. 

Contours of reactant concentration at the first pairing for twodimensional simulation with cold reaction; Sc = 0.5. The highest 

contour level is 0.92, the lowest is 0.27, and the interval is 

0.054. 

Iso-surface of temperature for temperature-dependent reaction at 

pairing time. The surface is at 50% of peak.

85
Aero-Assisted Orbital Transfer Vehicle Flow Fields 

Scott T. Imlay, Principal Investigator 

Co-Investigators: Donald W. Roberts and Moeljo Soetrisno 

Amtec Engineering, Inc. 

Research Objective 

To develop a computational fluid dynamics (CFD) code for the 

analysis of rarefied flow about hypersonic vehicles. 

Approach 

A three-dimensional Navier—Stokes code with finite-rate 

chemistry is modified to solve the two-dimensional Burnett 

equations with rotational nonequilibrium. The Burnett equations 

are more accurate than the Navier—Stokes equations at low 

densities where the continuum assumption is beginning to jail. 

Surface-slip boundary conditions are used along the walls. 

Accomplishment Description 

The Burnett equations with rotational nonequ ii brium are sok ed 

using a hybrid implicit method. The inviscid and Navier—Stokes 

terms are solved using an implicit upwind scheme based on 

Yoon's LU-SGS algorithm. The Burnett terms are treated 

implicitly using two steps of a line Gauss—Seidel relaxation. The 

code has been applied to the calculation of a one-dimensional 

shock wave structure and two-dimensional hypersonic flow past 

a two-dimensional circular leading edge. The accompanying 

figure compares the translational temperature contours calculated using the Navier—Stokes and Burnett equations for a 

circular leading edge with a Knudsen number of 0.5 and a Mach 

number of 20. The shock wave is much more dispersed with the 

Burnett equations than with the Navier—Stokes equations. 

Significance 

Since wind tunnel simulations of hypersonic vehicles is difficult 

and expensive, future hypersonic vehicle design will increasingly depend on CFD flow analysis. Most CFD codes solve the 

Navier—Stokes equations. For low density flows, however, the 

Burnett equations are more accurate. 

Future Plans 

The usefulness of the hypersonic flow analysis code will be 

increased by adding catalytic wall-boundary conditions.

Publication 

Imlay, S. T. "Solution of the Burnett Equations for Hypersonic 

Flows near the Continuum Limit." AIAA Paper 92-2922, 

July 1992. 

Comparison of the translational temperature contours for a 

circular leading edge; Knudsen number = 0.5; Mach = 20. 

86
Wing—Nacelle--Pylon Installations 

Gregory A. Intemann, Principal Investigator 

Co-Investigator: Todd R. Michal 

Douglas Aircraft Company/McDonnell Aircraft Company 

Research Objective the nacelle—pylon region. The wing—nacelle—pylon mesh is 

To perform initial calibrations and validations of a multi-zone comprised of 25 grid zones and contains approximately 

Navier—Stokes code for use in predicting powered propulsion- 1.3 million points. A typical solution with this mesh requires 

system installation effects for wing—nacelle—pylon configurations 16 megawords of memory and approximately 25 Cray Y-MP 

on transport aircraft, hours.

Approach 

The three-dimensional Navier—Stokes code used for this study 

was NASTD. The code algorithm is a time-dependent finitevolume implicit scheme based on upwind differencing. A 

Baldwin—Lomax turbulence model was used in this study. To 

identify potential areas of propulsion—airframe interaction and jet 

interference, the wing—nacelle—pylon combination results were 

compared to wing-alone and isolated nacelle—pylon results. The 

computational results were also compared to flight-test data. 

Accomplishment Description 

Several cruise angles of attack and corresponding engine-power 

settings were evaluated using a representative transport aircraft 

wing—nacelle—pylon configuration in conjunction with two 

different engine configurations. The wing-alone and isolated 

nacelle—pylon configurations were also analyzed at typical 

cruise conditions. Comparisons between flight-test results and 

wing—nacelle-pylon computations indicate good agreement in

Significance 

NASTD analyzes transport aircraft wing—nacelle—pylon configurations with power effects. Good correlation of the NASTD 

predictions with flight-test data lends confidence to the computed solutions. Comparisons between the isolated and wingmounted nacelle—pylon NASTD solutions predict sizeable wing 

influence on the pylon shelf and core exhaust flow fields. The 

pylon-shelf designs developed using a typical wind tunnel test 

set-up, where an isolated (powered) nacelle—pylon without a 

wing is used, may not perform satisfactorily in the presence of 

the wing. 

Future Plans 

We will continue to apply the NASTD code to predict jetinterference effects for wing-mounted nacelle—pylon installations 

to establish the changes in interference characteristics due to 

increasing engine-bypass ratio (up to 15), and due to nacelle 

positioning variations (longitudinally and vertically). 

/ 

1

Y 

NASTD-computed surface-pressure contours on a transport aircraft wing—nacelle—pylon; Mach = 0.828, a = 2.85 degrees, and 

altitude = 35,000 ft.

87 
Analysis of High-Speed Civil Transport Configurations 

Kenneth M. Jones, Principal Investigator 

Co-Investigators: Victor R. Lessard and M. A. Takallu 

NASA Langley Research CenterNiGYAN, Inc./Lockheed Engineering and Sciences Company 

Research Objective 

To determine the accuracy, applicability, and efficiency of three 

NASA Langley three-dimensional Navier–Stokes solvers for 

analyzing the low-speed aerodynamic characteristics of HighSpeed Civil Transport (HSCT) configurations. 

Approach 

The three-dimensional Navier–Stokes codes (CFL3D, TLNS3D 

and FMC1) were used to analyze an HSCT configuration 

operating at low speed. The codes were run using the same grid, 

and comparisons were made of the results obtained from the 

three different codes. 

Accomplishment Description 

The analyzed HSCT configuration was optimized for supersonic 

cruise at Mach 3. The three codes were used to analyze the 

configuration at subsonic free-stream conditions (Mach 0.3). All 

codes obtained converged solutions on the vehicle, but with 

varying degrees of accuracy and run times. The grid did not have 

optimal grid-clustering near the body for the free-stream initial 

conditions. A new grid has been generated and the analysis will 

be run again. Even with a slightly coarse grid, the results from 

the CFL3D code compared well with the available experimental 

data. The analyzed grid had approximately 850,000 points and a 

typical run required about 40 megawords of memory and 

10 Cray-2 hours. In support of the NASA Langley High-Speed 

Research (HSR) program, a different HSCT configuration was 

analyzed with CFL3D. This configuration was designed for 

supersonic cruise at Mach 2.4 and consisted of a cranked arms' 

wing with a 73 degree inboard leading-edge sweep and a 

60 degree outboard sweep. The results from the analysis a 

shown in the accompanying figure. The color contours sho 

total pressure and the lines on the surface are particle traces 

simulating flow near the body of the HSCT. The vortex-dominated flow field that is characteristic of these configurations at 

low speed is clearly shown. These results were compared with 

experimental data and are shown in the second figure. In 

addition to excellent force comparisons (the bottom portion of 

the figure), predicted flow-field characteristics very closely 

matched results obtained with laser-light sheet-flow visua?izatL 

data from 

Significanc 

One of the 

speed characteristics of configurations optimized for high-spec 

cruise—particularly the high-lift characteristics of the vehicles. 

Experimental and computational work is being done to understand and control the dominant flow physics. The computation 

demonstrate the ability of these methods to help solve the 

problems. An initial concern regarding the use of these codes 

was the convergence rate of the codes at low free-stream Mach 

numbers. In all cases, the convergence rate of the methods was 

acceptable.

Future Plans 

We plan to complete our analysis and comparison of the three 

codes on the Mach 3 configuration. A high-lift version of the 

Mach 2.4 configuration will be analyzed. 

Total pressure contours from the CFL30 solution on the AST-210 

configuration. 

Comparison between CFL3D results and experimental data for 

the AST-210. 

88
Structure and Dynamics of Multidimensional Flames 

K. Kailasanath, Principal Investigator 

Co-Investigator: G. Patnaik 

Naval Research Laboratory 

Research Objective 

To develop a better understanding of the differences in the 

propagation and extinction of premixed flames in Earth's gravity 

and in microgravity. 

Approach 

A detailed, time-dependent, multidimensional, multispecies, 

numerical model is used as a tool to simulate flames in different 

gravitational environments. The model includes finite-rate 

detailed chemical kinetics coupled with algorithms for convection, thermal conduction, viscosity, molecular diffusion, and 

external forces such as gravity. 

Accomplishment Description 

The effects of gravity and heat loss have been systematically 

investigated in a series of simulations of hydrogen—air flames. 

These simulations show that both gravity and heat loss to the 

confining walls are simultaneously required to cause the

downward-propagation flammability limit observed on Earth. 

The calculated limit of 9.75% hydrogen in the air and the details 

of the extinguishment process are in excellent agreement with 

experimental results. A typical calculation takes 30 Cray Y-MP 

hours. 

Significance 

This is the first multidimensional numerical simulation that has 

considered effects such as heat loss and gravity along with 

detailed chemical kinetics. The excellent agreement with the 

experimental results suggests that such simulations can now be 

used to more confidently predict the behavior of flames in 

reduced-gravity environments. 

Future Plans 

The code will now be used to simulate and understand a 

number of interesting flame phenomena observed on Earth and 

in space. 

Comparison of upward- and downward-propagating flames to zero-gravity flames in a 10% hydrogen—air mixture.

89
Unsteady Nozzle Flow Fields 

K. Kai lasanath, Principal Investigator 

Co-Investigator: R. Ramamurti 

Naval Research Laboratory 

Research Objective 

To study the transient flow field in a nozzle and gain a better 

understanding of the nonlinear interactions between the shock 

waves generated in the start-up procedure and the nozzle 

geometry. 

Approach 

The three-dimensional time-dependent conservation equations 

describing compressible flows in complex geometries are solved 

using unstructured grids and a finite-element flow solver. An 

adaptive refinement strategy in conjunction with an arbitrary 

Lagrangian—Eulerian form allows the tracking of complex shockstructures and moving objects in the flow field. 

Accomplishment Description 

A series of simulations have been carried out to study the effects 

of initial conditions such as nozzle exit clearance and flow 

leakage through baffles on the transient flow in the nozzle. A 

configuration has been identified that avoids excessive pressure 

buildup in the nozzle and unwanted loads on the nozzle—motor

junctions. The accompanying figure shows a typical threedimensional flow field in the nozzle at a particular time. A 

typical simulation takes 50 Cray Y-MP hours. 

Significance 

These simulations have been crucial in understanding the 

unsteady start-up process in nozzles used for propulsion systems 

such as the Trident II. Furthermore, a capability now exists to 

simulate unsteady compressible flows in complex geometries 

with moving parts. 

Future Plans 

No further simulations are planned. However, the code can now 

be used to simulate other problems involving transient compressible flows in complex geometries. 

Publication 

Ramamurti, R.; Kailasanath, K.; and Lohner, R. "Numerical 

Simulation of Unsteady Flow in a Nozzle." 1992 JANNAF 

Propulsion Meeting. Indianapolis, IN, Feb. 1992. 

-- - 

'_._ .-..

..-

Instantaneous pressure distribution in a nozzle from 0.7 atm (blue) to 8.7 atm (magenta). 

90
Simulation and Control of Slender Wing Rock 

Osama A. Kandil, Principal Investigator 

Co-Investigator: Ahmed A. Salman 

NASA Lewis Research Center/Old Dominion University 

Research Objective 

To simulate the limit-cycle rock motion of slender delta wings 

and investigate active control of the rock motion using antisymmetric forced oscillation of the wing leading-edge flaps. 

Approach 

The problem has three sets of equations. First, the unsteady 

compressible Euler equations are written relative to a moving 

frame of reference. Second, the unsteady linearized Navierdisplacement equations are used in the moving frame of 

reference to compute the grid displacements whenever the 

leading-edge flaps oscillate. Third, are the Euler equations of 

rigid-body motion for the wing or for the wing and its flaps. This 

set computes the wing motion for the wing-rock problem and is 

solved in sequence with the first set. To control wing rock, the 

third set is solved in sequence with the first and second sets. 

Accomplishment Description 

A delta wing with a sweep-back angle of 80 degrees at 

35 degrees angle of attack and Mach 1.4 is considered. The

wing-mass moment of inertia about its x axis is 0.02, the bearing 

damping coefficient is 0.2, and the spring stiffness is 0.74. The 

unsteady Euler equations are solved for locally conical flows. For 

these flow conditions, the steady flow is asymmetric. The Euler 

equations of fluid flow and of rigid-body dynamics are sequentially integrated in time. The accompanying figure shows the 

results of this case. The time responses show the long time 

needed to build up the unstable response. The roll angle and 

roll-moment coefficient increase in time with an increase in 

frequency. The limit-cycle response is reached at t 21. To 

control the wing-rock response, a leading-edge flap hinge is 

assumed to be at the 76% location of the local half-span length. 

Flap motion is antisymmetric. The equation for the wing-flap 

motion is sequentially integrated in time. The second figure 

shows the time responses for the wing. 

Significance 

The codes treat multi-disciplinary problems where fluid-flow 

equations and rigid-body-dynamics equations interact. 

Future Plans 

Aeroelastic equations have been added to the existing codes to 

solve for multidisciplinary problems where fluid-flow and 

aeroelastic equations interact. The vertical tail-buffet problem is 

being simulated. 

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Roll-angle and roll-moment-coefficient responses for an unstable Active control of wing-rock motion using an anti-symmetric 

rolling motion (wing rock). oscillation of the leading-edge flaps.

91
Ic rilperattire Carbon 

I)ixidc

Soot 

Volume 

I- ft t f1

Magnitude of 

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lruensitv 

84 

(cm 

Unsteady Ethylene Jet-Diffusion Flames 

Carolyn R. Kaplan, Principal Investigator 

Co-Investigators: Elaine S. Oran and Seung W. Baek 

Naval Research Laboratory/Korea Advanced Institute of Science and Technology 

Research Objective 

To examine the effect of radiation transport on the development, 

structure, and dynamics of an unsteady luminous ethylene jetdiffusion flame. 

Approach 

The model solves the two-dimensional time-dependent 

Navier—Stokes equations with finite-rate chemical reaction, 

soot formation, and radiation transport models to simulate 

fluctuating axisymmetric ethylene—air jet-diffusion flames. The 

convection terms are solved using the barely implicit correction 

to flux-corrected transport algorithm. The radiative heat flux is 

found from a solution of the radiative-transfer equation using the 

discrete ordinates method. The soot formation and growth model 

is based on a set of coupled ordinary differential equations 

describing nucleation, surface growth, and coagulation. The 

effects of thermal conduction, molecular diffusion, and viscosity 

are included through explicit finite-differencing formulations. 

Accomplishment Description 

The model is validated through comparisons with experimental 

data for a low-velocity laminar ethylene—air diffusion flame. 

Simulations conducted for the high-velocity (10 meters/ 

second) fluctuating flame show that soot is the dominant 

absorbing—emitting medium and that radiation transport changes

the temperature, soot concentration, and chemical species 

distributions in the flame. Radiation transport reduces the 

maximum flame temperature by 140 K, reduces the width of the 

high-temperature and sooting regions in the flame, and causes 

the flame to accelerate slightly. The increase in the Reynolds 

number due to both lower viscosity and higher velocities results 

in more and stronger vortical structures in the inner jet shear 

layer. Results show the opacity of the flame is in the transitional 

regime between optically thin and optically thick, and heat 

transfer by radiation dominates opacity by conduction or 

convection. These computations require 6-10 megawords of 

memory and 10-20 Cray Y-MP hours. 

Significance 

Although radiative properties of weakly radiating nonluminous 

diffusion flames can be calculated with reasonable accuracy 

using the laminar flamelet concept, prediction of strongly 

radiating luminous flames has remained a greater challenge. 

In this research, the effects and the interactions between the 

individual processes have been studied. 

Future Plans 

The model will be extended to study stabilization mechanisms of 

ethylene jet-diffusion flames during the process of flame lift-off 

and reattachment to the burner. 

Contours of temperature, radiative intensity, and soot-volume fraction in a 3 cm x 34 cm region of an ethylene jet-diffusion flame. 

92
' ..

i_ 

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rid 

Unstructured Grid-Generation/Flow-Solver Calibration 

Steve L. Karman, Jr., Principal Investigator 

Co-Investigator: Gregory S. Spragle 

General Dynamics, Fort Worth Division 

Research Objective 

To demonstrate the capability to numerically model threedimensional steady-state and time-dependent multi-body flow 

fields using state of the art unstructured-grid methods. 

Approach 

Unstructured grid generation codes are used to discretize complex 

two- and three-dimensional configurations. Unstructured-grid 

flow solvers then compute time-dependent Euler/Navier—Stokes 

equations on the initial mesh or use grid refinement and 

adaption to enhance the mesh to improve solution accuracy 

and/or track moving boundaries within the mesh. 

Accomplishment Description 

A robust two-dimensional unstructured-grid generation code has 

been developed for triangular meshes with arbitrary geometries. 

A two-dimensional axisymmetric Navier—Stokes code with a 

two-equation turbulence model has been developed and 

calibrated on several two-dimensional and axisymmetric test 

cases. An initial three-dimensional unstructured-grid generation 

code has been developed and has successfully generated 

tetrahedral meshes for selected simple test cases. A threedimensional Navier—Stokes code with a two-equation turbulence 

model has been developed and tested on simple geometries 

using tetrahedral grids created from structured three-dimensional 

grids. Calibration of the three-dimensional flow solver has been 

limited due to the inability to generate a three-dimensional 

unstructured mesh. Further calibration work has been done on 

the turbulence model for two-dimensional geometries. The first 

case was a high-speed inlet flow field and the second case was a 

multiple-slot ejector flow field. Comparisons with experimental 

data and other computational fluid dynamics codes were made. 

An alternative approach to analyzing complex geometries 

involves the use of Cartesian grids generated by recursively 

subdividing a root cell that encompasses the entire domain. 

Results from initial two-dimensional work shows promise of 

overcoming many of the grid generation problems encountered 

with the traditional unstructured triangular and tetrahedral 

schemes. Extension of the method to three dimensions should be 

relatively easy. 

Significance 

The development of this unstructured-grid capability is significant in the evaluation of store separation from aircraft and 

complex problems such as inlet unstart. 

Future Plans 

The calibration of the Cartesian-based unstructured-grid method 

in two dimensions will continue. The development of the threedimensional code will begin. Once the three-dimensional 

capability becomes available, the code will be tested on various 

complex geometries and flow fields.

Publications 

1. Karman, S. L., Jr. and Spragle, G. S. "Development of an 

Unstructured CFD Method." AIAA Paper 91-0019, Jan. 1991. 

2. Karman, S. L., Jr. "Development of a 3D Unstructured CFD 

Method." Dissertation, University of Texas at Arlington, 

May 1991. 

3. Karman, S. L., Jr. and Spragle, G. S. "Calibration of 2D 

Unstructured Grid Methods on Propulsive Flow Fields." 

AGARD FDP Symposium, Aerodynamic Engine—Airframe 

Integration, Fort Worth, TX, Oct. 1991. 

Aircraft store grid and flow field.

93
Turbulent Flow over Riblet-Mounted Surfaces 

George E. Karniadakis, Principal Investigator 

Princeton University 

Research Objective 

To quantify the drag reduction caused by riblets and to understand the responsible physical mechanism. 

Approach 

The incompressible Navier–Stokes equations are discretized in 

space using a mixed spectral-element Fourier method and in 

time using a high-order splitting method. The computational 

domain is a channel with a lower wall mounted with streamwise riblets and a smooth upper wall. A typical run requires 

3-5 Cray Y-MP seconds and approximately 20 megawords of 

memory per time-step. 

Accomplishment Description 

Simulations were performed in the Reynolds-number range 

500-5,000, corresponding to laminar, transition, and turbulent 

regimes. The effect of riblet geometry was also investigated as a 

means of finding an optimum drag-reduction design; V-grooves 

(triangular), curved, and shark-scale-like (dermal dendrite) riblets 

were tested. Results suggest that in the laminar regime there is 

no drag reduction, but in the transitional and turbulent regimes 

drag reduction of up to 6% exists for the V-groove ribletmounted wall compared to the smooth wall. A detailed study 

of turbulence structure was conducted for triangular riblets. All 

turbulence statistics including turbulence intensities, Reynolds 

shear stresses, skewness, and flatness factors were accurately 

computed. Based on our results, a physical model for the dragreducing mechanism was devised. 

Significance 

This computation is the first direct numerical simulation of 

turbulence in a complex geometry domain and demonstrates 

the flexibility of the hybrid spectral-element Fourier method in 

efficiently simulating such flows. Riblets are used for civil 

transport airplanes and result in significant savings. Therefore, 

the accurate prediction of such flows is very important. 

Future Plans 

Simulations at higher Reynolds numbers for the optimal riblet 

geometry—the shark-like riblets—will be addressed. A largeeddy-simulation approach based on renormalization groupsubgrid models will be followed.

Instantaneous streamwise vorticity contours (vertical component) 

on a plane perpendicular to the flow. Red represents positive 

values and white represents negative values. 

94
Turbulent Flow over a Backward-Facing Step 

G. E. Kamiadakis, Principal Investigator 

Co-Investigator: S. A. Orszag 

Princeton University 

Research Objective 

To rest the subgrid model obtained with the renorrnalization 

group theory (RNG), especially in the early turbulence regime, 

and to investigate the flow physics in transitional and fully 

turbulent states. 

Approach 

The incompressible Navier—Stokes equations are integrated using 

the spectral element method and a subgrid model derived by the 

RNG of turbulence. Approximately 1 million degrees of freedom 

were employed in the computation, requiring 20 megawords of 

memory and 3 Cray Y-MP seconds per time step. 

Accomplishment Description 

Following our earlier work on equilibria and transition in flow 

over a backward-facing step, a numerical study of turbulent flow 

up to Reynolds number 9,000 was conducted. The size of the 

separation zone is a decreasing function of the Reynolds number 

in the transition regime and is constant in the fully turbulent

regime. We performed simulations at three Reynolds numbers 

(2,500, 4,000, and 9,000) corresponding to early transition, late 

transition, and fully turbulent regimes, respectively. Our simulation results agree with the experimental results and predict the 

variation of the separation length in the transition regime. 

Significance 

The flow over a backward-facing step represents a prototype for 

many internal flows where sudden expansions cause massive 

separation. It also serves as a good test case for both discretization 

and subgrid models for turbulent separated flows. 

Future Plans 

Future work will address the robustness of the RNG subgrid 

model with respect to discretization variations and will extend 

the simulation to Reynolds number 20,000. A companion 

experiment being performed at NASA Langley will provide 

detailed data for mean flow quantities and turbulence statistics. 

Instantaneous streamwise velocity contours in a flow over a backward-lacing step at Reynolds number 9,000. Red represents positive 

values and white represents negative values (separated flow).

1,1
Aeronautical Vehicle Radar Cross Sections 

Jack H. Kennedy, Principal Investigator 

Co-Investigator: Daniel A. White 

General Dynamics, Convair Division 

Research Objective 

To develop advanced computational radar cross-section (RCS) 

analysis techniques to design the next generation of high-stealth 

cruise missiles. 

Approach 

The work is centered around two electromagnetic scattering 

codes, the representation theorem method (RTM) and the finiteelement radiation model (FERM). RTM is a two-dimensional code 

that models linear dielectric and magnetic materials used in 

modern aeronautical systems. FERM is a three-dimensional code 

primarily used to analyze scattering from conductive bodies. Both 

codes are computationally intensive for even moderate-sized 

bodies. Work with RTM is concentrating on developing a 

capability to optimize geometry and material parameters for 

minimum RCS. FERM is being used to display induced surface 

currents on targets in order to model the special currents created 

by interactions between edges and interfaces on aeronautical 

vehicles. 

Accomplishment Description 

We validated RTM predictions for radar absorbing materials 

(RAM) with significant magnetic characteristics (large permeabi Iities). RAM-coated flat plates were constructed and the scattering 

response was measured for comparison with RTM predictions. 

Work was also done on validating the FERM code. Two fins of 

the type used on the Tomahawk cruise missile were modeled

using the code. RCS measurements were carried out over a range 

of 3-10 GHz at both horizontal and vertical polarizations for 

360 degree orientations of the targets. Predictions were generated and are currently being evaluated. A capability was also 

developed for extracting and displaying surface currents 

predicted in the code. In the accompanying figure the target 

segmentation and the results of current predictions for one of 

the fins analyzed are shown. Relative current amplitudes and 

directions are shown as short line segments over the surface of 

the fin. These results typically require about 1.5 Cray-2 hours 

and 20 megawords of memory. 

Significance 

Developing high-stealth aeronautical vehicles is a top priority of 

the United States military. As requirements increase, empirical 

design approaches have proved costly and are limited in their 

effectiveness. New analytic methods are required to speed the 

design effort and to find better designs. 

Future Plans 

After completing validation studies of the RTM and FERM codes, 

prediction capabilities will be improved. Improvements in RTM 

algorithms are under development. Studies of RTM and FERM 

target segmentation will be carried out to determine which type 

of segmentation produces the most accurate predictions for 

different target shapes, sizes, and radar frequencies. 

'••'I\ '\,'"—_ 

S -

\\\ \_

\ 

\ 

\ -

- 

H1 

Inc3d..,tfields A \ 

E1 

(HorI

1Z_.. ^ 45 deg

Cruise-missile fin segmented for analysis of the finite-element radiation-model 

code and the currents predicted for oblique field incidence. 

96
Orographically Forced Oscillations of the Martian Atmosphere 

Christian L. Keppenne, Principal Investigator 

Co-Investigator: Jean 0. Dickey 

Jet Propulsion Laboratory 

Research Objective 

Low-frequency oscillations caused by barotropic instability over 

orography cause significant variations of the length of day on 

Mars. Since the surface relief irregularities of Earth and Mars are 

similar in their scale heights, orographically forced modes 

effecting the Martian rate of rotation should also be expected. 

Approach 

A barotropic dynamical model in spherical coordinates with 

realistic topography is run at high resolution in several 6,000-501 

(Martian day) integrations (see the accompanying figure). The 

model is based on the primitive-equation system used in 

operational general-circulation models, but it is formulated 

without accounting for thermodynamic processes such as the 

cycles of water on Earth or 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 is spectrally analyzed. Composite 

maps of the free-surface height-anomaly fields are keyed to a 

filtered AAM time series to analyze the spatial details associated 

with the oscillations detected in the global AAM time series. 

Accomplishment Description 

The code is multitasked and the prognostic equations are 

integrated in spectral spherical harmonic space, using a semiimplicit scheme, but the nonlinear terms are computed on a 

latitude—longitude grid. At high-resolution (192 x 150 gridpoints), 

performance comes close to 1 GFLOP using all eight processors. 

At a more modest resolution, performance decreases to 

500 MFLOPS or 0.2 Cray Y-MP second per sol and about 

8 megawords of memory. Two robust oscillations with periods 

near 85 and 125 sols were identified in the model AAM time 

series. They persist over a significant range of model parameters 

or when seasonal variations are introduced in the forcing. The 

corresponding spatial variability shows standing oscillations with 

centers of action upstream and downstream from the highest 

Martian mountains (see the accompanying figures). 

Significance 

Because the total AAM of the Mars atmosphere system is 

conserved, it produces variations in the Martian day length on 

the order of tens of milliseconds. Predicting such changes is 

important for space navigation and will be beneficial to future 

space missions.

Future Plans 

Interference between the observed oscillations and comparatively shorter timescale phenomena will be examined. This will 

require a model with higher vertical resolution and the appropriate parameterization of some processes neglected in this study. 

Barotropic dynamical-model topography projected onto a 

rectangular grid.

kr%0-.

. imm, 

Composites of a free-surface height anomaly showing the four 

phases of the (a) 85-sol oscillation and (b) 125-so! oscillation. 

The composites are keyed to the maxima and minima of a 

filtered atmospheric angular momentum time series and its first 

derivative.

97
Low-Thrust Chemical Rockets 

Suk C. Kim, Principal Investigator 

Sverdrup Technology, Inc ./NASA Lewis Research Center 

Research Objective 

To study the flow fields and performance of low-thrust space 

propulsion rockets. 

Approach 

A full Navier—Stokes code with a finite-rate chemistry code 

(RPLUS) is used to calculate the flow fields and performance of 

low-thrust chemical rocket nozzles. The combustion processes 

of hydrogen and oxygen are modeled by an 8-species, 18-step 

reaction mechanism and are compared with experimental data. 

Accomplishment Description 

Calculations were made for the 25 lb1 gaseous hydrogen—oxygen 

thruster designed for Space Station Freedom, which uses filmcooling hydrogen injection to cool the thruster wall. The results 

for various mixture ratios and fuel film-cooling percentages show 

that the vacuum-specific impulse decreases by increasing the

mixture ratio and film-cooling percentages and agrees well with 

the experimental data. A typical run requires approximately 

4 megawords of memory and 11 Cray Y-MP hours. 

Significance 

The study demonstrates that the full Navier—Stokes code with a 

finite-rate chemistry can be used for performance prediction of 

various chemical rockets. 

Future Plans 

The flow fields and performance of various space propulsion 

rockets will be studied. 

Publication 

Kim, S. C. and VanOverbeke, T. J. "Performance and Flow 

Calculations for a Gaseous Hydrogen/Oxygen Thruster." 

J. Spacecraft and Rockets 28, no. 4 (1991): 433-438. 

CONTOUR L[LIL1S 701x9fl 1). 00000 

11.20000 

11. 400011 

Ii 6000(1 

II. 130001) 

• 1)0001) 

200(31) 

1. 30000 

1 6131)01) 

[.ROOD)) 

7.1)001111 

7. 70001) 

Mach number contours; mixture ratios = 3.12, fuel film-cooling = 85.5%. 

98
Fluid Flow of Jets in Cross Flow 

S. W. Kim, Principal Investigator 

Co-Investigator: T. J. Benson 

NASA Lewis Research Center 

Research Objective 

To develop and validate a numerical analysis method to 

investigate the large-eddy mixing phenomenon occurring in 

separated and recirculating fluid flows of jets in cross flows. 

Approach 

The Reynolds-averaged Navier—Stokes equations are solved 

using a finite-volume method that incorporates a partial differential equation for incremental pressure to obtain a divergence-free 

flow field. The turbulence is described by a multiple-time-scale 

turbulence model. 

Accomplishment Description 

Numerical results show that large-eddy mixing occurs in the 

wide region along the jet edge and that fluid particles in the 

center region of the jet do not easily mix with the cross flow. The 

multiple-time-scale turbulence equations successfully predict the 

horseshoe vortex located in the forward region of the jet while 

several numerical investigations using k-c turbulence models and 

Reynolds stress-turbulence models fail to predict the horseshoe 

vortex. The calculated fluid flow and the concentration field 

show that the local maximum concentration occurs where there 

is minimum velocity along the jet. This trend is in excellent

agreement with the measured data. The accurate numerical 

results obtained in the present investigation are attributed to the 

capability of the multiple-time-scale turbulence equations to 

resolve the nonequi I ibrium turbulence phenomena. Calculation 

of the fluid flow with 0.8 million grid points and 

7.3 million degrees of freedom requires 60 megawords of 

memory and 20 Cray Y-MP hours. 

Significance 

Turbulent flows similar to jets in cross flows can be found in 

a number of engineering applications. For example, in gas 

turbine combustors, a number of circumferentially distributed 

jets are used to ensure correct combustion and to dilute the hot 

gas entering the turbine. The accurate numerical results help to 

understand the large-eddy mixing and the turbulent transport of 

heat and mass in jets in cross flows. The capability to accurately 

solve complex fluid flows is useful for design and analysis of 

advanced fluid machinery. 

Future Plans 

Numerical investigations of heat transfer and topological 

analysis of separation and reattachment locations for jets in 

cross flows will be done. 

The computed flow field for an incompressible fluid injected into a cross flow with a 2:1 velocity ratio. The penetration of the jet is 

shown by the light gray cloud being swept downstream while the colored particle traces show the three-dimensional interaction of 

the streams.

99
Multistage Tu rbomach I nery Flows 

Kevin R. Kirtley, Principal Investigator 

Co-Investigators: John J. Adamczyk, Tim A. Beach, Mark L. Celestina, William A. Maul, and Rick A. Mulac 

Sverdrup Technology, Inc./NASA Lewis Research Center 

Research Objective 

To more fully understand the development of secondary flows 

and their impact on mixing and performance in multistage 

turbomachinery flows. 

Approach 

The average-passage Navier—Stokes equations accurately 

describe the time mean flow in each blade-row embedded in a 

multistage environment. Therefore, these equations are solved 

using an explicit Runge—Kutta time-integration scheme with a 

multigrid for analysis of axial, radial, and mixed-flow compressor 

and turbine stages. 

Accomplishment Description 

The two-stage Pratt & Whitney space shuttle main engine fuelturbopump turbine was simulated using experimental test-rig 

conditions. Previous simulations using design conditions were 

validated by a good comparison with the data at the cold-flow 

test-rig condition. The two-stage turbine simulation required 

21 megawords of memory and 40 Cray Y-MP hours. Two 

transonic mixed-flow turbine-stage builds were also simulated. 

The results matched experimental data quite well. The simulations were used to study development and migration of secondary flows in mixed-flow machines, identify potential loss 

mechanisms, and guide future redesign efforts. The single-stage 

simulations required 16 megawords of memory and 11 Cray Y-MP 

hours. 

Significance 

The average-passage Navier—Stokes equations are efficient and 

useful tools for studying the development and impact of secondary flows and the effect of deterministic unsteadiness on the time 

mean flow in multistage turbomachinery. The results can be used 

to improve performance. 

Future Plans 

Improved numerical and mesh-generation algorithms and 

average-passage stress modeling will be developed. In particular, 

a diagonalized lower—upper multigrid technique is being 

developed for multiblock mesh topologies for more accurate 

simulations of turbomachinery with splitter blades, part-span 

shrouds, and tip-clearance gaps.

Publications 

1. Kirtley, K. R. and Beach, T. A. "Deterministic Blade-Row 

Interactions in a Centrifugal Compressor Stage." ASME 

Paper 91 -GT-273, 1991. 

2. Kirtley, K. R.; Beach, T. A.; and Rogo, C. "Aeroloads and 

Secondary Flows in a Transonic Mixed-Flow Turbine Stage." 

Presented at the 37th International Gas Turbine Conference, 

June 1992. 

Hub static pressure for a single-stage mixed-flow turbine.
Turbulent Boundary Layers with Suction 

W. Kollmann, Principal Investigator 

Co-Investigator: P. Mariani 

University of California, Davis 

Research Objective 

To simulate turbulent boundary layers with suction and statistically evaluate the boundary-layer structure in Euleriari and 

Lagrangean frames. The investigation of the structural properties 

is based on iso-surfaces for fluctuating scalar variables and the 

statistical quantities relevant for second-order closures are to be 

determined. 

Approach 

The direct simulation of the boundary-layer flows is carried out 

with the spectral method developed by Spalart at NASA Ames 

using Fourier modes for the directions parallel to the wall and 

Jacobi polynomials for direction normal to the wall. The suction, 

or blowing, velocity is included as a single mode. The structural 

analysis is based on triangularized iso-surfaces allowing the 

calculation of topological and geometrical properties as a 

function of the level values. The statistical analysis is based on 

averaging in time, horizontal planes, and averaging over material 

points tracked during the simulation. 

Accomplishment Description 

The direct simulation of a flat-plate boundary layer with zero 

pressure gradient and a suction velocity vc,/u,,. = 0.0038 (corresponding to an available set of measurements) and a Reynolds 

number of 500, based on the displacement thickness, was 

10 

+ 

0

0 .25 .50 .75 1.00 1.25 1.50

V/DEL 

The kinetic energy of turbulence for the boundary layer.

completed. The statistical analysis of the resulting flow was 

carried out and all quantities for the budget of the Reynolds 

stress components and the dissipation rate were determined. 

Evaluation of the results shows good agreement with the 

experiment. Comparison with the Spalart results for the boundary layer without suction shows significant modification of the 

flow structure in the inner layer. The simulation of the boundary 

layer with a Reynolds number 011000 was also finished and the 

statistical evaluation completed. The first simulation (Reynolds 

number= 500) required 2 megawords of memory and 50 Cray-2 

hours. 

Significance 

The understanding of the relation between mass transport 

through the wall and the turbulence structure of the boundary 

layer should be improved considerably. The development of 

closure models for the prediction of turbulent boundary layers 

with transpiration should be aided by the statistical evaluation of 

the simulation results. 

Future Plans 

The project objectives were achieved and evaluation of the 

results is near completion. 

3.0 

2.5 

2.0 

1.5 

1.0 

.5 

0 

iO1 100 101 102 103 

YPLUS 

The effect of suction on normal stress components with suction 

are experimental (with suction, Watts and Brundrett, 1979; 

without suction, Spalart, 1988).

ITIT
Receptivity, Transition, and Turbulence Phenomena 

Linda D. Kral, Principal Investigator 

Co-Investigators: William W. Bower and Alexander Pal 

McDonnell Douglas Research Laboratories 

Research Objective 

To investigate the physical processes in fluids by simulating 

transition mechanisms or turbulence phenomena occurring in 

flows ranging from incompressible to supersonic speeds. 

Approach 

Three areas of research are studied: boundary-layer transition, 

shear-layer dynamics, and boundary-layer receptivity. In the 

first two areas, direct numerical simulations (DNS) of the 

Navier—Stokes equations are performed, and in the third area, 

linearized Navier—Stokes equations are recast in spectral form 

and solved. 

Accomplishment Description 

A parametric study of boundary-layer receptivity was performed 

to consider the excitation of Tollmien—Schlicting waves in a 

laminar boundary layer by a downstream-traveling plane soundwave interacting with the flow through a porous strip which links 

the layer to a fluid-filled cavity acting as a receptivity site. By 

varying the Reynolds number for the fixed-strip width, it was 

found that the receptivity coefficient decreases with increasing 

Reynolds numbers. Resonance occurs when the ratio of cavity 

length to acoustic wavelength is 0.25. By varying the frequency 

of the sound wave for a range of strip widths normalized by the 

Tollmien—Schlicting wavelength, resonance again occurs (with 

maximal values of the receptivity coefficient) for the ratio of 

cavity length to acoustic wavelength equal to 0.25. The evolution of two-dimensional spatially evolving disturbances in a 

heated-water boundary layer was simulated. Previous investigations showed discrepancies between the direct simulations and 

stability theory for a heated boundary layer. A comparison 

between DNS, theory, and experiment is shown in the accompanying figure. Amplification rates are shown for several levels of 

heating. Comparison of the Navier—Stokes solution with parallel 

theory shows excellent agreement. Amplification rates are shown 

in the figure at several frequencies for AT = 3.48° F. The simulation results are in good agreement with both parallel and 

nonparallel theory and follow the experimental trends. 

Significance 

Understanding boundary-layer transition and receptivity is 

crucial in the design of aerodynamic vehicles. Through investigations of the flow physics, the influence of transition and 

turbulence on skin friction, mixing enhancement, and separation 

can be predicted and controlled.

Future Plans 

We are investigating the physics of the compressible, boundarylayer transition process. An algorithm is being developed to 

perform DNS of compressible flows, including curvature effects, 

during the nonlinear stages of transition. Analysis of wall 

compliance on boundary layer receptivity to free-stream 

disturbances is being conducted. 

o Navier-Stokes (Inner max U) 

o Navler-Stokes (max v) 

o Parallel theory [Lowell] 

o Parallel theory [EI.Hady at al.] 

—1

(a)

—2

AT (°F) 

I I

0 

o Navler-Stokes (Inner max u) \ 

a Navier-Stokes (max v) 

- Parallel theory [El-Hady at al.] \ \ 

• - - - Nonparallel theory (Inner max U) 

[EI-Hady at al.] 

• Experiments (Inner max u) 

(b) [Strazlsar at al.] 

.6 .8 1.0 1.2 1.4 

Fu2 

Comparison of amplification rates from direct numerical 

simulations with linear stability theory and experiments for 

different levels of wall heating at (a) F = 1 and (b) different 

frequencies for AT = 3.48° F. 

.6 

.4 

.2 

0 

—.2 

—.4

10 

1.6 

102
Turbulence Modeling for Three-Dimensional Flow Fields 

Linda D. Kral, Principal Investigator 

Co-Investigator: John A. Ladd 

McDonnell Douglas Research Laboratories/McDonnell Aircraft Company 

Research Objective 

To investigate turbulence models used in numerical simulations 

of complex viscous flows for critical aircraft components. 

Approach 

A three-dimensional zonal Navier—Stokes code with a twoequation turbulence model is used to predict the turbulent flow 

fields around flight vehicles. An implicit approximately factored 

upwind scheme is employed to solve the three-dimensional 

compressible Favre-averaged Navier—Stokes and energy equations together with the k-s two-equation turbulence model. 

Three zero-equation algebraic models are also available in the 

code. Four low-Reynolds-number forms of the k-c turbulence 

model are available. 

Accomplishment Description 

Comparison of the k-s model is made with the algebraic models 

for complex turbulent flows. Validation of the low-Reynoldsnumber models has been made for several turbulent fiat-plate 

boundary layers (Mach = 0.3, 2.0, 3.0, and 4.5). The behavior in 

the near-wall region is best predicted by the Jones—Launder and 

Speziale low-Reynolds-number models. A calculation of a multislot ejector nozzle shows that the Jones—Launder k-s model is 

better able to predict flow characteristics than the Thomas 

algebraic model. Mach contours for a nozzle pressure ratio of 14 

and secondary total-pressure ratio PTOTp = 0.175 are shown in 

the accompanying figure for the k-s and algebraic model. Dark 

blue contours represent Mach 0.1 and magenta contours 

represent Mach 2.2. The interior static surface pressure and test 

data were compared. The algebraic model falsely predicts a 

large separated region at the last ejector slot, but the 

k-s calculation does not show this tendency and agrees well with 

the test data. An investigation of a two-stream hypersonic mixing 

layer with finite-rate chemistry was conducted with the twoequation turbulence model and the Thomas shear-layer algebraic model. The k-s model predicted the correct ignition delay, 

while the algebraic model immediately ignited at the injection 

point. 

Significance 

It is unclear how these turbulence models will perform for more 

realistic geometries. Emphasis is placed on comparing the 

capabilities of several turbulence models for more complex 

geometries and flow fields of engineering interest. Simulation of 

turbulent flows found on aircraft would provide an extremely 

valuable design tool.

Future Plans 

An assessment of the capabilities of different models will be 

made for a single-slot ejector nozzle. The investigation of 

turbulent flow about an F/A-i 8 fuselage forebody with wing 

leading-edge extension at high angle of attack will continue and 

the different low-Reynolds-number models will be evaluated for 

this flow and the contributions of the corrections for rotation and 

curvature. An algebraic Reynolds-stress transport model will also 

be implemented. The ability of the k-s model to predict transition is under investigation and modifications to the turbulence 

model will be sought for the transition zone. 

POT N9dWW ---

Mach number contours for (a) a multi-slot ejector nozzle using 

the k-t and (b) Thomas algebraic turbulence models.

103
Translating-Strut Scramjet Inlet 

Ajay Kumar, Principal Investigator 

Co-Investigator: Dal J. Singh 

NASA Langley Research Center 

Research Objective 

To investigate the effects of change in geometry by strut movement on the overall performance of an opposite-sweep sidewall 

compression inlet. 

Approach 

A three-dimensional Navier—Stokes code is used to numerically 

simulate the flow through a translating-strut scramjet inlet. The 

inlet has variable geometry to allow for variation of the contraction ratio (CR). At low speed the CR is decreased, while at high 

speed the CR is increased for efficient performance of the inlet 

over a wide range of Mach numbers. Effects of boundary-layer 

ingestion on the overall flow features are also investigated. 

Overall flow-field features such as the corner flow, topwall 

separation, shock-wave coalescence, cowl pressure increase, 

and flow distortion at the throat are investigated. The inlet shows 

good characteristics at both low- and high-CR strut locations. 

Accomplishment Description 

Numerical and experimental data are in good agreement, except 

in the separation zone where the calculations overpredict the 

pressure. The calculations show the need for a better turbulence 

model in the separation zone; the Baldwin—Lomax model is not 

appropriate. The ingestion of a top-surface boundary layer 

reduces inlet performance for both cases in terms of mass

capture, average-throat Mach number, and pressure recovery. 

The presence of the boundary layer also decreases the uniformity of the air profile entering the combustor. The accompanying figure shows Mach-number distribution in the inlet at various 

axial locations and along the cowl plane for the low-CR case. It 

shows a complex system of shock waves, expansion waves, and 

their interaction with each other and the boundaries. The two 

counter-rotating vortices and the large region of low-momentum 

flow near the top wall are clearly visible. An average run used 

20 megawords of memory and took 4 Cray-2 hours. 

Significance 

The present numerical investigation of a translating-strut inlet, 

when complemented with the wind tunnel tests, could have a 

favorable impact on the National Aero-Space Plane inlet design 

in particular, and high-speed inlets in general. 

Future Plans 

Efforts are under way to investigate the inlet unstart. 

Publication 

Singh, D. J.; Trexier, C.A.; and Hudgens, J. A. "Three-Dimensional 

Simulation of a Translating-Strut Inlet." AIAA Paper 92-0270, 

Jan. 1992. 

Mach number translating-strut inlet. 

104
High-Lift Aerodynamic Flow 

Dochan Kwak, Principal investigator 

Co-Investigators: Stuart Rogers and and Lyn Wiltberger 

NASA Ames Research Center 

Research Objective 

To accurately and efficiently perform numerical simulations of 

flow over high-lift aerodynamics configurations. 

Approach 

Existing two- and three-dimensional incompressible Navier—Stokes 

flow solvers (INS2D-UP and INS3D-UP) were used with a multizonal, overlaid Chimera-grid approach to simulate the flow over 

two-dimensional multi-element airfoils and three-dimensional 

wing configurations. Both Baldwin—Lomax and Baldwin—Barth 

turbulence models were used. 

Accomplishment Description 

The INS2D code was run on a number of different multi-element 

airfoil configurations. Good agreement with experimental results 

was obtained for most cases; however, there is a problem with 

the turbulence model's ability to handle significantly separated 

cases. The Baldwin—Barth model performed better than the 

Baldwin—Lomax model. Most of the computing time was used 

by the INS3D-UP code to compute a number of different wing 

flows. Good quantitative agreement with the experiment was 

obtained for several basic wing shapes. Some high-lift computations were performed over a rectangular wing with vortex fences

(shown in the accompanying figure). The fences supplemented 

the lift of the wing by trapping a vortex on the top surface 

of the wing. A typical coarse-grid wing configuration using 

100,000 grid points requires a half-hour of Cray-2 time and 

5 megawords of memory. A finer grid calculation using 

500,000 grid points requires 3-4 Cray-2 hours and 25 megawords 

of memory. A fine grid calculation of flow over a four-element 

airfoil using 35,000 points can be obtained in 4 Cray-2 minutes. 

Significance 

We have shown the capability to compute significantly complex, 

high-lift flow problems. These problems are important to aircraft 

designers and are a pacing item of research in the aircraft 

industry. The ability to run many two-dimensional jobs in a very 

short time will significantly impact the design process. Currently, 

there is a great need for this capability in three dimensions. 

Future Plans 

We will further test and develop turbulence models for high-lift 

aerodynamic flows in two and three dimensions. Testing the 

flow over a three-dimensional wing in landing configuration 

(with flaps and slats deployed) is planned. 

Particle traces showing the vortex formed in the flow over a wing with vortex fences. The particle 

traces are colored by pressure; blue = low pressure and orange = high pressure.

105
nal Plane 

--

Computational

"I 

Grid 1e1o(itv Vectors 

Space Shuttle Main Engine Flow 

Dochan Kwak, Principal Investigator 

Co-Investigators: S. Yoon, C. Kiris, L. Chang, and R. Williams 

NASA Ames Research Center 

Research Objective 

To extend, validate, and apply incompressible Navier—Stokes 

solvers to computational fluid dynamics analysis of liquid rocketengine pump components. 

Approach 

The three-dimensional incompressible Navier—Stokes equations 

in generalized coordinates are solved in steadily rotating frames 

of reference. The solution is based on a pseudo-compressibility 

approach with two different numerical algorithms. The first 

algorithm, INS3D-UP, uses higher-order upwind differencing 

and the Gauss—Seidel line-relaxation scheme. The second 

algorithm, INS3D-LU, is based on the lower—upper symmetric 

Gauss—Seidel implicit scheme. 

Accomplishment Description 

Two major components of a rocket-engine pump have been 

investigated: the Rocketdyne inducer and the advanced pumpimpeller. As a benchmark problem, the flow through the 

Rocketdyne inducer is numerically simulated. A typical calculation requires 8 megawords memory and 12 Cray-2 singleprocessor hours. A coarse-grid solution is obtained with a single 

zone by using an algebraic turbulence model. Numerical results, 

in the form of relative velocity and relative flow angles, compare 

fairly well with experimental data. Multi-zone fine-grid computation uses a one-equation Baldwin—Barth turbulence model 

derived from a simplified form of the standard k-c model 

equations. In recent computations, a bull-nose cavity analysis 

is included. 

Significance 

An efficient and robust design-stage solution for three-dimensional 

pump flows was developed and validated. Computational 

analysis of new design variations for rocket engines will expedite 

the development of future heavy-lift launch vehicles. 

Future Plans 

We are continuing the study by generally simulating future 

rocket engines. 

Geornet,y

Surface Colored

by Pressure

Computed results of the flow through the Rocketdyne inducer. 

iIII.
Hydrogen—Air Reacting Flow Fields in Drag Reduction External Combustion 

H. T. Lai, Principal Investigator 

Sverdrup Technology, Inc. 

Research Objective 

To numerically simulate three-dimensional reacting flow fields in 

external combustion to reduce drag. 

Approach 

The finite-volume RPLUS code recently developed for reacting 

flows is employed to compute the solutions of compressible 

laminar Navier—Stokes and species equations. The numerical 

scheme has a lower—upper solver requiring only scalar inversions of the flow equations, but blocks diagonal inversion of the 

species equations because of the chemical source terms. Finiterate chemistry of hydrogen and air is used based on a combustion model containing 18 reaction steps and nine species with 

nitrogen assumed inert. 

Accomplishment Description 

Numerical flow fields have been computed to simulate an 

external burning experiment to study external combustion as a 

mechanism for drag reduction in aerospace vehicles when 

operating in the transonic regime. The model consists of a small, 

flat plate connected to a 12 degree expansion ramp, representing 

a typical underside of hypersonic nozzle walls facing the 

ambient flow. An air stream flows over the model at Mach 1.26 

and hydrogen fuel is sonically and transversely injected along a 

row of 26 orifices upstream from the expansion corner. A flame 

holder to stabilize combustion is installed at the expansion 

corner downstream from the injection. Self-sustained combustion behind the flame holder can be numerically achieved 

using ignition by increasing fuel temperature initially and then 

reducing it to the desired level after combustion has been 

established in the flow field. Results are compared to measured 

data for code validation. The computations are performed on a 

multiblock grid of 48 x 50 x 96 simulating only half of the 

model. The converged solution requires 350 Cray-2 hours and 

25 megawords of memory.

Significance 

Numerical prediction is consistent with the external burning 

experiment, confirming possible hydrogen—air combustion at 

low speeds. This can be used for drag reduction in hypersonic 

vehicles when operating at the off-design transonic-flow 

conditions. 

Future Plans 

This study is complete. 

Publication 

Lai, H. T. "Computation of H2/Air Reacting Flowfields in Drag 

Reduction External Combustion." AIAA Paper 92-3672, July 1992. 

AK flow Hyd.ogen injection (U = 126,T410 R, P=8ps) (2€066c.s,U= 1,T410R,P35Opsâs) 

Fkamahokler 

Thermal exp.non 

The model shows external combustion. Near-surface temperatures are colored and the particle traces are white.

107
Steady and Unsteady Turbomachinery Flow Fields 

Budugur Lakshminarayana, Principal Investigator 

Co-Investigators: S. Fan and R. Kunz 

Pennsylvania State University/General Motors Technical Center 

Research Objective 

To develop and apply explicit steady three-dimensional 

Navier—Stokes and unsteady Euler procedures for computation 

of complex turbomachinery flow fields. To improve the understanding of particular flow phenomena, including secondary 

flows and radial mixing, tip-clearance effects, wake decay and 

transport, and the influence of Reynolds-stress anisotropy on 

mean flow fields. To develop an efficient time-accurate solution 

to resolve rotor/stator interaction and develop turbulence 

modeling approaches for these applications. We will investigate 

stability and accuracy issues associated with the numerical and 

modeling strategies employed. 

Approach 

A compressible explicit flow solver was developed to perform 

inviscid and viscous two- and three-dimensional steady and 

unsteady turbomachinery flows. Numerical and turbulencemodeling strategies were developed, implemented, and evaluated to achieve more robust and more accurate mean-flow 

predictions for a wide class of turbomachinery applications. 

Numerical topics addressed directly include artificial diffusion 

issues (stability and accuracy), time-accurate boundary-condition 

treatments, and numerical compatibility of two-equation and 

Reynolds-stress transport models. Several complex turbomachinery flows for which experimental data were available 

were analyzed to calibrate the methods developed and to 

investigate flow phenomena of interest. 

Accomplishment Description 

A three-dimensional explicit steady full-Navier—Stokes capability 

has been developed that incorporates two-equation and algebraic Reynolds-stress transport models. The numerical stability 

issues associated with incorporating transport turbulence models 

in an explicit solution procedure were investigated and quantified. The procedure was used to perform viscous simulation 

studies in several three-dimensional turbomachinery blade-row 

configurations. The code has been modified to obtain timeaccurate solutions. Wake transport through a flat-plate cascade 

has been computed. 

Significance 

These efforts should be useful to the turbomachinery research 

and development community. They include accuracy and 

numerical compatibility issues of current and new turbulence 

models and the ability of Navier—Stokes codes to reconcile 

important three-dimensional and unsteady-flow physics. 

Future Plans 

Efforts include development and installation of turbulent heattransfer modeling for compressible turbine flows, enhanced 

coupling of the unsteady algorithm with an unsteady boundarylayer code, and further centrifugal compressor flow research.

It 

I 

Computed contours of dynamic pressure just aft of midchord in a 

compressor cascade with tip clearance. High-momentum jet-like 

fluid in the tip-gap region and low-momentum fluid associated 

with the tip-clearance vortex and endwall boundary layer are 

evidenced. 

I I I 

- Unsteady velocity

T.E. 

Suct

- 

-

TE 

Pres 

L. E. 

I I I I

-1 0

X 

Vector plot of unsteady velocities for a fiat plate cascade with 

incoming gust. 

108
Steady and Unsteady Viscous Flow 

Budugur Lakshminarayana, Principal Investigator 

Co-Investigators: Yin-Hsiang Ho and Anton Basson 

Pennsylvania State University 

Research Objective 

To develop an efficient and accurate numerical analysis for the 

prediction of steady and unsteady turbulent flows in turbomachinery applications. 

Approach 

The technique developed is for incompressible flow and is an 

extension of the pressure-based method. A finite-volume 

approach with a non-stagger grid formulation is used in the 

analysis. A predictor—corrector type algorithm is applied to the 

governing equations to ensure the time accuracy and the 

convergence are improved using a multigrid scheme. A lowReynolds form of a two-equation turbulence model is used to 

simulate the turbulent flow. 

Accomplishment Description 

The unsteady flow field in a flat-plate cascade (at seven different 

reduced frequencies of the inlet gust, two Reynolds numbers, 

and two different steady loadings) has been computed. Detailed 

interpretation of the effects of reduced frequency and Reynolds 

number on the aerodynamic response function is carried out. In 

addition, unsteady flow through a compressor cascade, for 

which flow data are available, is computed and correlated with 

the existing experimental data. The agreement between the 

predicted and the measured blade-pressure and the wake data is 

good. The flow field over an airfoil with a separation bubble is 

computed with good agreement between the time-averaged 

values and the data. In addition, three-dimensional viscous flow 

through the tip-clearance gap of a turbine is computed and good 

agreement with the measured data is shown in the accompanying figure. Each two-dimensional unsteady-flow computation 

required approximately 1-2 Cray-2 hours and 4 megawords of 

memory. For three-dimensional tip-clearance flow, each run 

required 4-6 Cray-2 hours and 16 megawords of memory. 

Significance 

The ability to simulate the unsteady viscous flow inside the 

airfoil cascade is a prelude to the prediction of rotor/stator 

interaction and noise generation in turbomachinery. A better 

understanding of the flow field inside the turbomachinery 

through three-dimensional viscous flow simulation can improve 

the performance and design of turbomachinery.

Future Plans 

We will extend the code to include the three-dimensional 

effects. To simulate the unsteadiness caused by the upstream 

stator wake flow on a rotor, an instability wave model will be 

employed to estimate the characteristic length and time scales of 

the wake flow. The small-scale motion can be modeled by the 

traditional two-equation model which provides a statistical 

description of non-isotropic, nonhomogeneous turbulence. 

Publications 

1. Ho, Y. H. and Lakshminarayana, B. "Computation of 

Unsteady Viscous Flow using an Efficient Algorithm." AIAA 

Paper 91-1597, 1991. 

2. Basson, A. H.; Kunz, R. F.; and Lakshminarayana, B. "Grid 

Generation for Three-Dimensional Turbomachinery Geometries including Tip Clearance." AIAA Paper 91-2360, 1991. 

Computed pressure coefficient contours on the endwa!l of the 

tip-clearance region.

109
Integrated Hypersonic Vehicle Flow-Field Analysis 

Scott L. Lawrence, Principal Investigator 

Co-Investigators: Bradford C. Bennett and Gregory A. Molvik 

NASA Ames Research Center/MCAT Institute 

Research Objective 

To develop the capability to compute the entire aerothermal 

environment of a vehicle configuration (tip-to-tail capability), 

such as a the National Aero-Space Plane (NASP), in hypersonic 

flight, including the forebody external flow, and the flow into the 

inlet, through the combustor, and onto the afterbody/nozzle 

portion of the vehicle. 

Approach 

The compressible Navier—Stokes (CNS) and upwind parabolized 

Navier—Stokes (UPS) codes solve the Navier—Stokes and 

parabolized Navier—Stokes (PINS) equations using algorithms 

appropriate for high-Mach-number flows. The UPS code is used 

for streamlined regions of a configuration. The zonal methodology employed by the CNS code makes it well suited for complex 

portions of the vehicle such as inlets, wings, and fins. Both codes 

include turbulence, transition, equilibrium and finite-rate air 

chemistry models. In addition, the codes TUFF and STIUFF have 

recently been developed for flows requiring a fully coupled finiterate chemistry capability. 

Accomplishment Description 

Computations were completed examining the effects of various 

parameters such as bluntness, gas model, grid density, and 

marching-step size on forebody performance. Perfect-gas results 

of the bluntness study were compared with results of a wind 

tunnel study and generally showed good agreement with 

measured surface-pressure and heat-transfer data (see the 

accompanying figure). Each of the forebody computations 

required approximately 2 Cray-2 hours, including time required 

for the time-dependent nose solution, and 4 megawords of 

memory. Computational design studies were performed for a 

wave-rider configuration using TUFF. A hydrocarbon combustionchemistry model is being implemented in this code for future 

propulsion-system design studies. 

Significance 

Because ground-based facilities are unable to simulate the 

environment that the NASP will encounter, numerical simulations are required to assess the importance of real-gas effects on 

the NASP flow field, particularly their impact on the flow 

properties at the engine inlet.

Future Plans 

A tip-to-tail computation is under way and should be completed 

during the next NAS Operational Year. This computation will be 

followed by more detailed investigation of the combustor flowfield physics pursuant to the implementation of an advancedturbulence/chemistry-interaction model for reacting shear layers. 

Finally, advanced wall-catalysis boundary conditions will be 

included within the PNS solver. 

An equilibrium air-chemistry model showing Mach contours 

and surface streamlines for Mach 16 flow past the Generic 

Option II forebody. 

110
Numerical Investigation of Vehicles Similar to the National Aero-Space Plane 

C. C. Lee, Principal Investigator 

Co-Investigators: William Bower, Shawn Hagmeier, Brad Hopping, Morzata Mani, Scott Van Horn, 

Charles Vaporean, Patrick Vogel, and Darrell Weber 

McDonnell Douglas Corporation 

Research Objective 

To calibrate computational fluid dynamics (CFD) codes related 

to the National Aero-Space Plane (NASP) using existing data and 

to perform parametric studies using the calibrated CFD codes to 

aid in the design of vehicles similar to the NASP. 

Approach 

Three-dimensional Navier—Stokes codes (CFL3DE, eM1k, 

NASTD, MDCFPNS3D, MDCPNS) are continually upgraded 

and calibrated for external- and internal-flow analysis. These 

calibrated codes are applied in the design and evaluation of 

hypersonic NASP vehicles to improve the vehicle shape and 

supply vehicle performance parameters. 

Accomplishment Description 

A Mach 10 hypersonic cruise vehicle was selected for preliminary design analysis using the parabolized Navier—Stokes (PNS) 

code MDCPNS. The preliminary analysis was performed on a 

simplified vehicle (no wings, tails, or scramjet engine package). 

The analysis covered the majority of the flight envelope from 

M0,, = 2-10 at a = 4 degrees. Runs were made on a 1.97% scale 

model at wind tunnel conditions (ReJL = 2.0E6 1/ft) and for the 

full-scale vehicle along the flight trajectory, including equilibrium air thermochemistry. The results included (1) the bow 

shock shape, which was used to design and place the wings and 

tails, (2) pre-tunnel flow fields, to estimate sting loads and aid in 

the placement of instrumentation, (3) tunnel-to-flight scaling for 

aerodynamic forces and moments, and (4) differences in inlet 

flow fields between tunnel and flight, especially the effect of 

boundary-layer thickness. 

Significance 

The CFD work is essential for hypersonic vehicle design. The 

lower-order physics (the PNS methodology) have proven to be 

useful in the placement of wings and tails and for performing 

rapid screening of hypersonic vehicles over a flight envelope. 

Future Plans 

The vehicle configuration is undergoing constant refinement. 

Tip-to-tail analysis is being planned using the CFL3DE, NASTD, 

and MDCFPNS3D codes. Work to significantly advance the state 

of the art in hypersonic vehicle design and performance prediction using CFD is being done.

Mach 10 hypersonic cruise vehicle. 

Predicted surface and flow-field pressures; M00 = 10, 

a =4 degrees, h = 120 km/ft.

111
Subsonic/Transonic Flutter Boundaries 

Elizabeth M. Lee, Principal Investigator 

Co-Investigator: Mike Gibbons 

NASA Langley Research Center/Lockheed Engineering and Sciences Company 

Research Objective 

To evaluate the aeroelastic version of the three-dimensional 

Euler/Navier—Stokes CFL3D code for wing-flutter predictions. 

The code was modified for the aeroelastic analysis of wings. 

Approach 

The configuration studied is the first Advisory Group for Aeronautical Research and Development standard aeroelastic 

configuration for dynamic response that was tested. The wing 

had a quarter-chord sweep angle of 45 degrees, a panel aspect 

ration of 1.65, a taper ratio of 0.66, and a NACA 65A004 

constant-airfoil section. The mounted model is shown in the 

accompanying figure. Flutter data for this model are reported 

over a range of free-stream Mach numbers from 0.338 to 1.141 

The flutter-speed index versus Mach number for this model 

exhibits the characteristic transonic dip near Mach 1. 

Accomplishment Description 

The aeroelastic version of CFL3D in the Euler mode was used to 

compute a complete flutter boundary for the 45 degree sweptback wing. The grid contained 261,129 grid points and required 

approximately 26 megawords of memory for a transient calculation. Flutter analyses were performed at free-stream Mach 

numbers of 0.338, 0.678, 0.90, 0.96, 1.07, and 1.14. The 

determination of the flutter point at each Mach number required 

one steady-state calculation which was used as a starting point

for several transient calculations at varying dynamic pressures. 

The dynamic pressures were chosen so that the wing response 

varied from stable, at a lower dynamic pressure, to unstable, at a 

higher dynamic pressure, thus bracketing the flutter point. Each 

flutter-point prediction required approximately 26 Cray-2 hours. 

The results of the analyses are compared to the experimentally 

determined flutter-speed index and flutter-frequency ration and 

are shown in the figure. This shows that at subsonic free-stream 

Mach numbers, the computed flutter-speed index agrees well 

with the experimental value while the computed frequency ratio 

is slightly below the experimental values and the frequency 

ration agrees well with the experimental value. Past Mach 1, 

where the flutter boundary is particularly sensitive to Mach 

number, the computed boundary indicates a sharper rise than 

the experimental boundary. 

Significance 

The good agreement between the computed and experimental 

values tends to validate the unsteady-Euler aerodynamic method 

for subsonic and transonic Mach numbers. 

Future Plans 

We will investigate the differences in the predicted flutter 

speeds for the supersonic free-stream Mach numbers using the 

Navier—Stokes equations and apply this aeroelastic analysis 

method to a National Aero-Space Plane configuration. 

Subsonic/transonic flutter boundary computed using an unsteady-Euler aerodynamic method. 

112
25 [ Numerical 

Local-asymptotic I J 

Parametric-resonance 

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25 

20 

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10 

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210 

10 

Resonant-Triad Interaction in an Adverse Pressure-Gradient Boundary Layer 

Sang Soo Lee, Principal investigator 

Co-Investigator: M. E. Goldstein 

Sverdrup Technology, Inc.JNASA Lewis Research Center 

Research Objective 

The nonlinear resonant-triad interaction for a Blasius boundary 

layer is analyzed for an adverse pressure-gradient boundary 

layer. 

Approach 

Our interest is in the nonlinear interactions that arise from the 

continued downstream growth of a resonant triad of initially 

linear instability waves. We assume that the adverse pressure 

gradient is weak and, therefore, that the instability growth rate 

is small. This ensures that there is a well-defined critical layer 

located somewhere within the flow and that the nonlinear 

interaction is effectively confined to that layer. In the initial 

interaction, the oblique instability waves exhibit faster than 

exponential growth and the growth rate of the two-dimensional 

mode remains linear. This is known as "parametric resonance." 

The interaction becomes fully nonlinear and the growth rates 

become fully coupled once oblique-mode amplitudes become 

sufficiently large. However, the coupling terms are now quartic. 

More importantly, new nonlinear interactions now come into 

play and eventually have a dominant effect on the instabilitywave development. 

Accomplishment Description 

The amplitude equations, which are fully coupled integrodifferential equations, are solved numerically. The results are

obtained by using the Adams—Moulton method with variable 

orders (up to the twelfth) to advance the solutions downstream 

from the prescribed upstream linear states. The integral terms on 

the right sides of the amplitude equations were computed by 

using the eleventh-order (9 point) Newton and Cotes integration 

formula with the upstream "tails" evaluated analytically from 

the upstream linear solutions. The code required about 

4-8 megawords of memory and approximately 2 Cray Y-MP 

hours. 

Significance 

Linear growth of the two-dimensional mode allows the amplitude to reach a level that produces a parametric resonance in the 

oblique modes, which then allows them to grow at an accelerated rate until they become large enough to interact with 

themselves. This self-interaction produces a further enhancement 

in their growth, ultimately ending in a singularity at a finite 

downstream position. This explosive growth is then transferred 

to the plane wave through the mutual interaction and backreaction terms. 

Future Plans 

In order to compare experimental results, we need to include the 

viscosity effect. We will study that by solving a set of partial 

differential equations. 

Magnitude of the oblique instability-wave amplitude versus Magnitude of the plane instability-wave amplitude versus 

streamwise distance for different initial amplitudes (from left to stream wise distance. 

right, 0. 1, 0.01, 0.001, 0.0001); solid lines = numeric, dotted 

lines = local-asymptotic, dashed lines = parametric-resonance.

113
Unsteady Flows for Naval Applications 

Yu-Tai Lee, Principal Investigator 

Naval Surface Warfare Center 

Research Objective 

To develop a numerical design tool for an axial turbomachine to 

predict flow unsteadiness and associated aeroacoustics in order 

to improve the turbomachine's efficiency and reduce its noise 

level. 

Approach 

A three-dimensional time-accurate Euler method was developed 

for inviscid flow through rotor/stator geometries. The rotor and 

stator are defined separately in local moving coordinates that are 

attached to each of them. The time-stepping technique is used to 

carry out the time-accurate inviscid flow analysis. To enhance 

computational efficiency, which is crucial to a time-accurate 

calculation method for such complex geometries, a new 

Gauss—Seidel successive iterative procedure is used. This is 

necessary to solve the massive system equation devised for the 

geometrically thin rotor/stator blade surfaces. A linear relationship between the perturbed pressure on the blade surface and 

the wake vortex is presented in an analytical-matrix form that 

most effectively and efficiently enables us to implement the 

nonlinear pressure-type Kutta condition. 

Accomplishment Description 

Numerical results for a stator/rotor cascade were carried out and 

strong interaction between the shed vortex and the downstream 

blade is seen. An unsteady calculation takes approximately 

5 Cray Y-MP hours for 300 time steps. The method was used to 

redesign a shipboard high-pressure ventilation fan. The original 

fan could not meet the aeroacoustic and aerodynamic performance requirements of a narrow range of flow rate versus 

pressure rise required by the system it was installed in. Utilizing 

the present method enabled designers to screen candidate 

geometries accurately and efficiently. The noise level for the 

new fan (shown in the accompanying figure) is about 

20 decibels lower than the original fan. A complete design 

case takes about 50 Cray Y-MP hours. 

Significance 

These solutions are important in understanding the physics of 

rotor/stator flows and provide an efficient and effective analysis 

tool for early stages of design.

Future Plans 

We will deduce numerical modeling and predictions for marine 

propellers. 

Publications 

1. Lee, Y. T.; Bein, T. W.; Feng, J. Z.; and Merkle, C. L. "Unsteady 

Rotor Dynamics in Cascade." ASME Paper 91-CT-1 47, 1991. 

2. Lee, Y. T.; Feng, J. Z.; Slipper, M. E.; and Merkle, C. L. 

"Redesign of an Incompressible High-Pressure Fan using 

CFD." 37th ASME International Gas Turbine Conference, 

Cologne, GE, June 1992. 

3. Lee, Y. T. and Bein, T. W. "Unsteady Pressure and Flow of 

Blade Interactions." Displayed at the 4th Navy IR/IED 

Symposium, Laurel, MD, 1991.

M.-` 

Redesigned shipboard high-pressure ventilation fan. 

114
Skewed Compressible Mixing Layers 

Sanjiva K. Lele, Principal Investigator 

Stanford University 

Research Objective 

To determine the effects of free-stream skewing on the behavior 

of compressible mixing layers. 

Approach 

Three-dimensional unsteady Navier—Stokes equations are solved 

for the skewed mixing-layer flows using the linear-stability 

analysis of the skewed three-dimensional flow to prescribe 

inflow disturbances. The unsteady dynamics of the large-scale 

structures and the overall statistical evolution of the flow is 

studied. 

Accomplishment Description 

A pre-existing computer code for solving the three-dimensional 

unsteady Navier—Stokes equations was adapted for our purposes. 

To prescribe the inflow conditions, the linear stability problem 

for a three-dimensional skewed mixing-layer laminar basic flow 

was formulated and solved. The instability results were analyzed 

to isolate the effects of free-stream skewing and flow compressibility. Under incompressible conditions the skewing effect was 

found to increase the maximum spatial amplification rate by 

a factor of three for U 2/U 1 = 0.5 and by a factor of nine for 

U2/U 1 = 0.75, with the most amplified wave propagating in the 

direction of effective mean shear. Compressibility decreased the 

skew-induced enhancement. For unskewed mixing layers at a 

convective Mach number of 0.8, the effect of skewing was a 

mild increase in the growth rate. A scale was developed to 

quantitatively explain the skewing effect and this allowed an 

effective convective Mach number to be defined and isolated the 

compressibility effect. Results of the analysis were incorporated 

into the three-dimensional nonlinear simulations and further 

work on the simulations is currently under way. Preliminary 

results from the simulations show that the vortical structures 

developing in a skewed mixing layer are more complex and 

three-dimensional. 

Significance 

The skewing effects on mixing layers can now be estimated. 

Enhanced instability may lead to more effective mixing and its 

control. 

Future Plans 

The simulation will be extended to systematically explore the 

effects of skewing and compressibility on the mixing layer.

Publication 

Lu, G. and Lele, L. K. "Inviscid Instability of Skewed Compressible Mixing Layers." Submitted to Journal of Fluid Mechanics, 

April 1992. 

Temporal simulation of a skewed compressible mixing layer. 

Contours of constant pressure have a cut of passive scalar at 

y = 0. Iso-surface pressure showing the rotational region is 

drawn with a wire mesh. The plotted pressure level is 0.55, 

while the minimum, maximum, and ambient pressure levels are 

0. 38, 1. OZ and 0.71, respectively. A cut of passive scalar at the 

mid-position of the mixing layer is overlaid; blue = lower stream 

and red = upper stream. The velocity ratio = 0.6, the temperature 

ratio = 1, the convective Mach number = 0.8, the Reynolds 

number = 500, and the grid size = 121 x 64 x 64.

115
Turbulence Effects on Stagnation-Point Flow 

Sanjiva K. Lele, Principal Investigator 

Stanford University 

Research Objective 

To study the effects of free-stream turbulence on stagnation-point 

heat transfer. 

Approach 

Direct numerical simulations (DNS) are performed to study the 

stagnation-point heat transfer with free-stream turbulence. 

Accomplishment Description 

A pre-existing computer code for solving the three-dimensional 

unsteady Navier—Stokes equations with periodic boundary 

conditions is being adapted. The stagnation-point flow requiring 

the incorporation of a no-slip surface was developed. Preliminary simulations of the interaction of a two-dimensional jet, with 

a core-flow Mach number of 0.5, impinging on a surface were 

performed. The flow exhibited self-sustained large-amplitude 

oscillations corresponding to rolled-up vortices in the jet flow. 

These vortices are expected to significantly change the heat

transfer at the plate surface. Simulations where the organized 

vortices are absent but the free-stream turbulence is imposed 

(such as isotropic turbulence at the inflow boundary) are 

currently being developed. To aid this process, asymptotic 

analysis of the compressible viscous stagnation-point flow was 

developed. The results of this analysis are being combined with 

the inflow free-stream turbulence prescription to conduct DNS. 

The impinging jet-flow simulations are being refined. 

Significance 

Understanding the processes that increase the heat transfer when 

free-stream turbulence is present will lead to the better design of 

gas turbines. 

Future Plans 

The simulations will be extended to three dimensions and the 

effect of free-stream turbulence on stagnation-point heat transfer 

will be studied. 

Numerical study of free-stream turbulence effects on stagnation-point heat transfer. The temperature contours in a laminar jet (entering 

from the left) impinging on a flat plate (right) are shown. A thermal boundary layer can be seen as a cluster of red contours along the 

right wall. The data were generated with an explicit, full Navier—Stokes calculation on a 281 x 281 mesh. The jet inlet Mach number is 

0. 5, yielding a Reynolds number of 1,000 based on the width of the jet. 

116
Three-Dimensional Parabolized and Full Navier—Stokes Techniques 

Clark H. Lewis, Principal Investigator 

Co-Investigator: Bilal A. Bhutta 

VRA, Inc. 

Research Objective 

To develop fast and accurate three-dimensional parabolized 

Navier—Stokes (PNS) and Navier—Stokes techniques to predict 

effects of gas chemistry, hypersonic flows over complex lifting 

configurations with compression surfaces, fins, flaps, stabilizers 

and other complicated control surfaces. 

Approach 

A unique PNS scheme is used as the baseline. Unlike conventional PNS schemes, this three-dimensional PNS scheme does 

not include any sub-layer approximation. A zonal approach is 

used to divide the flow field into a streamwise sequence of 

smaller regions. In each zone, a PNS solution is done and 

then, if necessary, the solution is further improved by using 

Navier—Stokes iterations consisting of a streamwise relaxation 

algorithm. A new flux-vector-splitting approach is used to treat 

the convective flux terms for a general real-gas. An implicit 

shock-fitting approach is used that treats the bow shock as a 

sharp discontinuity and predicts its location. A hybrid differencing approach is used and involves central differencing in the 

smooth shock-free regions and fully upwind flux-vector splitting 

across embedded shocks. The computational grids are generated 

using a fast elliptic grid-generation scheme that requires only 

15% more computing time than a corresponding cylindrical 

grid-generation scheme. 

Accomplishment Description 

The PNS and Navier—Stokes techniques developed have been 

used to study perfect-gas and equilibrium-air hypersonic flows 

over various lifting and maneuvering configurations, re-entry 

vehicles (RVs), and other high-speed missiles. Hypersonic flows 

over a wide variety of three-dimensional multiconic shapes with 

and without a bent forecone, cuts and flats, strong flares, fins, 

and spin tabs were computed, and the predicted flow fields were 

found to be in excellent agreement with available data. Supersonic and hypersonic flows over typical high-speed finnedmissile configurations were also analyzed to study gas-chemistry, 

fin-thickness, and fin-deflection effects. Calculation required 

1 Cray Y-MP hour and approximately 10 megawords of memory. 

Significance 

The three-dimensional PNS/Navier—Stokes techniques developed 

in this study will provide the basic numerical tools necessary to 

successfully design and operate various future and present 

supersonic and hypersonic aircraft, missiles, planetary probes, 

and RVs. 

Future Plans 

We will continue to develop and extend the basic PNS and 

Navier—Stokes algorithm for enhanced stability and accuracy, 

and extend the PNS algorithm to study blunt-body flows. We will 

focus on extending chemical and/or thermal nonec1uilibrium, 

complex three-dimensional shapes, surface ablation, and 

thermal-radiation effects.

Publications 

1. Bhutta, B. A. and Lewis, C. H. "Comparison of Hypersonic 

Experiments and PNS Predictions, Part I: Aerothermodynamics; Part II: Aerodynamics." journal of Spacecraft and Rockets 

28 (July-Aug. 1991): 376-393. 

2. Bhutta, B. A. and Lewis, C. H. "New Technique for Low- to 

High-Altitude Predictions of Ablative Hypersonic Flow 

Fields." Journal of Spacecraft and Rockets 29 (Jan.-Feb. 

1992): 35-50. 

Predicted pressure distributions at the body-end of a typical 

finned missile at Mach 15 and 5 degrees angle of attack. 

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Temperature contours for a Mach 20 flow over a 45 degree flare. 

IWA 
Electron-Beam Injections from the Space Shuttle 

C. S. Lin, Principal Investigator 

Co-Investigator: M. Muller 

Aurora Science, Inc./Southwest Research Institute 

Research Objective 

To examine the mechanism by which an electron beam radially 

expands after injection along magnetic-field lines. The electron 

beam is a radial expansion for better control of active experiments on the Space Shuttle. Computer simulations are conducted to study the injection of a high-density electron beam 

from the Space Shuttle into ambient plasma and neutral gas. 

Approach 

To understand the radial-expansion mechanism of an electron 

beam injected from a highly charged spacecraft, two-dimensional 

particle-in-cell simulations are conducted for a high-density 

electron beam injected parallel to magnetic fields from an 

isolated equ i-potential conductor into a cold background 

plasma. Realistic modeling of beam injection from a spacecraft 

requires injecting an electron beam from a finite isolatedconductor. Therefore, we use the capacity matrix method to treat 

the spacecraft surface as a finite isolated equi-potential conductor in an ambient plasma. Charged particles in the simulation 

system are advanced in time according to the equation of motion 

under the electric potentials. The procedure is repeated until the 

electron beam propagates to the end of the simulation system. 

Accomplishment Description 

The configuration space plot in the accompanying figure shows 

that the electron beam expands radially to the beam's electron 

gyroradius. The radial expansion occurs near the stagnation 

point close to the conductor surface. The initial expansion 

determines the beam envelope after the stagnation point. The 

radial expansion is caused by charge build-up at the stagnation 

point, producing a large transverse electric field. Accelerated by 

the transverse electric field, the beam electrons injected parallel 

to magnetic fields receive a large transverse kick. The maximum 

perpendicular velocity gained by the beam electrons approaches 

the beam-injection velocity. An average job run takes about 

2 Cray Y-MP hours and 4 megawords of memory. 

Significance 

Beam radial expansion is important because it affects the beam 

diameter and density—two critical parameters in determining the 

propagation and instability conditions of a finite-radius electron 

beam. The simulation results are applicable in understanding the 

measurements from several electron-beam injection experiments 

on the Space Shuttle.

Future Plans 

The code will be used to determine the dependence of the beam 

radius on beam density and other plasma parameters, and will 

be extended to simulate the motion of spacecraft. 

Publication 

Koga, J. and [in, C. S. "Simulation of Radial Expansion of an 

Injected Electron Beam." To be published in proceedings of the 

1991 Cambridge Workshop in Theoretical Geoplasma Physics. 

3.0

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The phase space plot (top) and configuration space plot (bottom) 

of the beam electrons showing the simulation results. Electron 

beams are injected from a plate at one-fourth of the system 

length from the left boundary. 

118
Turbopump Rotor/Stator Flows 

S. I. Lin, Principal Investigator 

Rockwell International, Rocketdyne Division 

Research Objective 

To utilize and extend the rotor-family code to calculate dynamic 

loading and analyze thermal characteristics of Space Shuttle 

main engine (SSME) and Space Transportation main engine 

(STME) turbopumps. 

Approach 

The rotor code was extended to be a full Navier—Stokes solver. 

It solves the two- and three-dimensional compressible unsteady 

Navier—Stokes equations on a system of patched and overlaid 

grids. 

Accomplishment Description 

The Rotor-2 code was extended to have multistage capability 

and was used to do a two-dimensional unsteady simulation for a 

two-stage SSME high-pressure fuel turbopump multi-rotor/multistation turbine blade. Good agreement was obtained between 

calculated results and NASA Marshall SSME cool-air test data. 

The Rotor-3 code was extended to be a full Navier—Stokes solver 

and was modified to compute three-dimensional steady-state 

subsonic SSME and supersonic STME turbine blades. Obtaining a 

three-dimensional steady-state solution for an isolated turbine 

takes 3-5 Cray-2 hours and uses approximately 100,000 grid 

points. 

Significance 

The current STME baseline engine design utilizes a gas-generator 

cycle. Because of the high-pressure ratio involved, the fuel 

turbine is required to operate in the supersonic regime. To date, 

only limited three-dimensional computational fluid dynamics 

analysis has been performed for supersonic turbine blades. 

Successful simulations of supersonic turbine-blade flows (steady 

or unsteady) will provide accurate predictions of blade dynamicloading and thermal characteristics that will help improve the 

performance and design of turbopumps. 

Future Plans 

The rotor code will be extended to calculate dynamic loading 

and thermal characteristics of the turbine blade in three dimensions. The upgraded code will be validated for supersonic 

turbine-blade flows and applied to support STME fuel-turbine 

design.

The velocity field indicates flow separation due to shock! 

boundary-layer interaction. The flow-separation region is at the 

suction side and near the blade leading edge. 

Mach number distribution at the midsection of the Mark-44 

supersonic turbine. A shock occurs at the suction side and near 

the blade leading edge.
Advanced Computational Fluid Dynamics Applications for Complex Configurations 

James M. Luckring, Principal Investigator 

Co-Investigators: Farhad Ghaffari and Brent L. Bates 

NASA Langley Research Center 

Research Objective 

To extend the Navier—Stokes computations about the F/A-i 8 

configuration to include the aerodynamic effects of wing 

leading-edge flap deflection. 

Approach 

A high-fidelity surface definition for the complete F/A-i 8 aircr,1i! 

has been generated in the form of cross-sectional cuts. Geom 

cal configuration simplifications include fairing over the in 

splitter plate, diverter, and leading-edge-extension (LEX) si 

The deflected-flap geometry has been simplified over the 

inboard 15% semispan by smoothly blending the deflected-i 

geometry to the undeflected-flap wing—body intersection. 

Subsequently, a multiblock structured-volume grid has been 

generated using a transfinite interpolation technique. The enti 

domain is represented by approximately 1 .25 x 106 grid point 

which are contained within 20 blocks using combinations of 

C-O, H-O, and H-H topologies. The flow field was computed 

with a version of CFL3D which has been extended for a gener,

ized surface patching. Turbulent effects are represented by an 

extended version of the Baldwin—Lomax algebraic turbulen 

model. The computations are done without the empenna) 

geometry. 

Accomplishment Description 

The thin-layer Navier—Stokes results with turbulent flow assumption have been obtained at a = 19 degrees, R? = 13.5 x 106, 

and M,,, = 0.34 and correspond to flight test conditions for the 

F/A-i 8 High Alpha Research Vehicle (HARV). The computations 

are performed for both 0 and 25 degree wing leading-edge flap 

deflections. A photograph of the HARV in flight (showing the 

tuft-flow survey over the wing, and the LEX and part of the 

fuselage along with the LEX vortex core visualized via smoke) is 

shown in the accompanying figure. Similarly, the unrestricted 

surface-flow pattern and the LEX vortex-core streamlines 

obtained from the computational results for both the blended 

(starboard) and the undeflected (port) flap configuration are 

shown. The simulated surface tuft-flow pattern, particularly over 

the blended-flap configuration, qualitatively resembles those 

observed in-flight that indicate a massively separated and stallen 

flow field on the outer wing panel. However, the computational 

results do not predict the LEX vortex breakdown observed in 

flight. This result is consistent with experimental wind tunnel 

data obtained on an F/A-i 8 model and demonstrates that the 

vortex breakdown is induced by the presence of the vertical tail. 

Excellent surface-pressure correlations between computations 

and flight-test results are achieved on the forebody with reasonable agreement over the LEX. A typical numerical solution 

corresponds to two-orders-of-magnitude reduction of the residuals, 

requiring 3,000 cycles and approximately 20 Cray-2 hours. 

Significance 

Meaningful Navier—Stokes analyses are performed on a realistic 

aircraft configuration including the aerodynamic effects from the 

wing leading-edge flap deflection.

Future Plans 

We will validate the computational results through wind tunnel 

experiments and perform the computations under different flow 

conditions. 

(a)

/ 

Surface/off-surface (low correlations with (light results. (a) cornputation;a= 19 degrees, R = 13.5x 106, and M_ = 0.34. 

(b)(light; a 20 degrees, R 11.5 x 16, and M,,, = 0.30. 

120
Ignition and Structure of a Diffusion Flame with Vortex Interaction 

M. G. Macaraeg, Prinicipal Investigator 

Co-Investigators: T. L. Jackson and M. Y. Hussaini 

NASA Langley Research Center/Old Dominion U niversity/ICASE 

Research Objective 

To study the fundamental physics of turbulent combustion or 

combustion in vortical flows relevant to high-speed propulsion 

devices. 

Approach 

The study involves a detailed asymptotic analysis and numerical 

simulation of the continuous evolution of the mixing between a 

fuel and oxidant which are allowed to mix and react in the 

presence of a viscous vortex. Emphasis is placed on the ignition 

time and location as a function of vortex Reynolds number (R) 

and initial temperature difference (PT) between the reacting 

species. 

Accomplishment Description 

The study focuses on understanding the controlling parameters 

necessary for ignition in a vortical flame. Typical results show 

the influence of the initial temperature differential and vortex 

strength on ignition onset and location. Ignition is seen as a hot 

spot in the T 1 field, displayed in the first figure. The vortical 

structure is slightly asymmetric due to an imposed perturbation 

temperature differential between the fuel and oxidant and the 

relatively weak vortex. As the vortex strength is increased the 

location of ignition moves to the center of an axisymmetric 

structure as depicted in the second figure. Increasing the vortex 

strength also reduces ignition time, which approaches a constant 

dependent on the initial temperature differential between the 

reactants. As the flame evolves, ignition occurs as a single hot 

spot in the viscous core and rapidly grows as time increases. At a 

later time, two diffusion flames begin to merge with the expanding, almost circular, flame. A typical calculation for this model 

on a fine grid requires 4 megawords of memory and 1 Cray-2 

hour. The calculations that will be performed after the development of the ignition model for high speeds will increase this 

memory and time to 12 megawords and 8 Cray-2 hours. 

Significance 

Reaction in a laminar diffusion flame depends on simultaneous 

processes of interdiffusion and chemical reaction. Some turbulent combustion processes may be described as a collection of 

laminar-flame structures that retain their identity but are distorted 

by the turbulence. No theory of ignition exists for high-speed 

flows. This study shows that the onset of ignition and its location 

in a laminar flame is strongly dependent upon the initial 

temperature differential and the vortex strength. 

Future Plans 

Analysis and parametric studies will continue with the addition 

of acoustic effects and compressibility in more complex flow 

fields. Systematic understanding of the relevant physics will 

precede a model incorporating detailed chemical kinetics in a 

full numerical simulation.

Publications 

Macaraeg, M. G.; Jackson, T. L.; and Hussaini, M. V. "Ignition 

and Structure of a Laminar Diffusion Flame in the Field of a 

Vortex." To be published in Combustion Science and Technology, 1992. 

Surface plot ofT1 for R = 28.0 and 13T = —2. 

-

R =20 

R = 28 

sI 

Contour plots of T, for = —2 and increasing vortex strength. 

121
Computer Simulation of Water—Membrane Interfaces 

Robert D. MacElroy, Principal Investigator 

Co-Investigators: Andrew Pohorille and Michael A. Wilson 

NASA Ames Research Center 

Research Objective 

To provide a molecular-level description of proto-biological 

processes and to explain phenomena occurring at interfaces 

between water and the simple membranes that form the walls 

of protocells. These membranes, built of chain molecules 

composed of polar-head groups and nonpolar chains arranged 

as bilayers, play a crucial role in promoting the organization of 

proto-biological organic material. 

Approach 

We will simulate the behavior of the atoms in the system by 

molecular dynamics (numerical integration of Newton's equations of motion). The results are used to compute the thermal 

and transport properties of the system, which can be compared 

directly with experimental results. In addition, the molecular 

dynamics results provide a detailed and accurate microscopic 

description of the system that is often not available experimentally 

Accomplishment Description 

Molecular dynamics simulations of the interface between a 

membrane and water have been performed. The atomic-level 

structure of the bilayer has been examined, and is in qualitative 

agreement with x-ray and neutron-scattering data from related 

hilayers. The structure and fluidity of the membrane have been

characterized at the microscopic level, and the membrane 

interior was found to be disordered. The orientation of water 

molecules at the water—membrane interface and the penetration 

of water into the membrane has been analyzed. The electrostatic 

surface potential across the interface has been calculated. 

Simulations require about 10 megawords of memory and at 

least 200 Cray Y-MP hours to obtain a sufficiently long trajectory. Subsequent data analysis requires 50 megawords of 

memory, but only moderate Cray Y-MP time. 

Significance 

The preferencial orientation of water near the water—membrane 

interface and the electrostatic potential across the interface 

determine the orientations of solute molecules at the surface and 

influence their transport across membranes. The preferential 

orientations of solute molecules at the interface can enhance 

reaction rates, thereby serving as a catalyst for the synthesis of 

new molecules in the proto-biological environment. 

Future Plans 

We will continue our studies of water—membrane interfaces, and 

will investigate interface chemical reactions and the transport of 

ions and small molecules across membranes. 

Membrane bilayer between two lame//ac of water. Polar-head groups of chain molecules (magenta) are located close to the interface. 

The terminal groups of the tails (pink) and C=C double bonds in the middle of the tails (yellow) are distributed throughout the membrane interior indicating the high fluidity of the membrane. The remaining tail groups are in blue. The ragged appearance of the figure 

edges is because only complete molecules are displayed, highlighting the degree of bilayer disorder. 

122
Supersonic Boundary-Layer Transition on a Cone at Incidence 

M. R. Malik, Principal Investigator 

Co-Investigator: P. Balakumar 

High Technology Corporation/NASA Langley Research Center 

Research Objective 

To study the stability and boundary-layer transition in supersonic 

flow past a sharp cone at incidence and compare the results with 

experimental data. 

Approach 

The flow field around the cone is decomposed into mean and 

perturbation parts. The mean flow was computed using the thinlayer Navier—Stokes code, TLNS3D. Calculations were performed 

for a 97 x 129 x 49 grid and were repeated for 97 x 257 x 49 

grid to check convergence. The highly refined grid normal to the 

cone surface was needed for boundary-layer stability analysis. 

This flow is unstable to the first-mode and cross-flow instability. 

The amplification of these disturbances is correlated with the 

onset of transition by way of the e method. 

Accomplishment Description 

Calculations were performed for a Mach 3.5 flow past a 5 degree 

semi-vertex cone at 2 degrees incidence. Even this small angle 

of attack introduces strong inflectional profiles along the leeward

line of symmetry. Compressible linear stability was studied for 

this flow and the eN method was used to estimate transition 

onset. The accompanying figure shows the computed N-factor 

traces up to N = 10. The computed transition location moves 

upstream with the azimuthal angle going from the windward line 

of symmetry to the leeward line of symmetry. This is in qualitative agreement with the experiment performed in the NASA 

Langley Mach 3.5 tunnel. The stream movement of transition is 

due to the increased cross flow away from the windward line of 

symmetry. 

Significance 

Transition in a fully three-dimensional supersonic boundary 

layer has been studied and compared with experiments. 

Future Plans 

Calculations will be made for higher Mach numbers and the 

reasons for some discrepancies near the leeward line of symmetry will be investigated. 

N•Idctor traces 

I.' . Angle of attack 2

Experimental results 

- - . 

________

• . -------.--..

') oi 

-

i_ .- ------,.-

-. -.

. r,( — 10 .°° 

N-factor computations for a cone at 2 degrees incidence and a comparison with the experiment.

123
.24 

.20 

.16 

.12

.30 

.20 

0

.10 

0

—.10

600 800 1000 

H 

M = 1.6, F = 0.06 x oblique breakdown 

Compressible Boundary-Layer Transition 

M. R. Malik, Principal Investigator 

Co-Investigator: C. L. Chang 

High Technology Corporation/NASA Langley Research Center 

Research Objective 

To study and understand the mechanisms involved in compressible boundary-layer transition and to provide a capability for 

boundary-layer transition prediction in both "quiet" and 

"disturbed" environments. 

Approach 

The evolution of disturbances in compressible boundary layers is 

governed by partial differential equations (PDEs) which can be 

derived from the complete Navier—Stokes equations. The original 

PDEs are reduced to a set of parabolized stability equations 

(PSEs) which are parabolic along the dominant flow directions 

so that the solution can be obtained by single-sweep marching. 

Both nonparallel and nonlinear effects of growing boundary 

layers are studied up to the transition stage using the PSE 

approach. The computational time required is an order of 

magnitude less than that for direct numerical simulation. 

Accomplishment Description 

Linear PSE calculations are performed for a Mach 1.6 flat-plate 

boundary layer. For the linear case shown in the accompanying 

figure, the growth rates computed by PSE agree quite well 

with those computed using the multiple-scale approach. For 

comparison, quasi-parallel results are also shown. The results 

indicate that nonparallel effects are more important for oblique 

waves (w = 50 degree) than for two-dimensional waves 

(w = 0 degree). Nonlinear PSE calculations are performed for 

the same Mach 1.6 boundary layer. Three-dimensional waves 

are the most amplified modes for supersonic boundary layers. 

A viable route to transition consists of two oblique primary 

waves interacting with each other. Our PSE results show that the 

initial amplitude required for oblique breakdown to occur is 

significantly lower when compared to that for the secondary 

mechanisms, which are dominant in incompressible flows. The 

figure shows the oblique breakdown process for three amplitudes at the given frequency. Transition location is clearly 

indicated by the rise of wall shear. 

Significance 

The PSE method can be used to study the nonparallel and nonlinear evolutions to disturbances in compressible boundary 

layers from the linear stage to transition. Coupled with receptivity phenomena (the initiation of disturbances in the boundary 

layer), this approach offers a computationally viable means for 

studying and predicting the complex and intricate phenomenon 

of compressible boundary-layer transition. 

Future Plans 

We will develop an axisymmetric version of the code to be used 

for transition studies in conical flows and studies at hypersonic 

Mach numbers.

M = 1.6, F = 0.4 x 10 4, '41 = 0°, 50° 

1800 2400 3000 

A 

Evolution of linear and nonlinear disturbances in Mach 1.6 

boundary-layer flow at the given nondimensional frequency 

(R = (a) Linear analysis; parallel (solid line), nonparallel 

PSE (dashed line), and multiple-scale approach (symbols). 

(b) Nonlinear analysis; wall shear versus Reynolds numbers for 

oblique-wave breakdown procedures with three different initial 

amplitudes. 

124
Turbulence in Compressible Fluids 

Nagi N. Mansour, Principal Investigator 

Co-Investigator: Gary N. Coleman 

NASA Ames Research Center 

Research Objective 

To create a data base for compressible turbulent flows and to 

improve the accuracy of compressible turbulence models. 

Approach 

Direct numerical simulations (DNS) of homogeneous compressible turbulence subjected to bulk compressions are generated, 

resolving all of the relevant scales so that no subgrid scale 

closure is required. The DNS results are used to test and guide 

improvements to models of compressible turbulence. 

Accomplishment Description 

Flow fields and one-point statistics were obtained for various 

turbulent states and compression histories; a typical run required 

20 Cray Y-MP hours and 2 megawords of memory. The manner 

in which the simulation data were used for model development 

is shown in the accompanying figure by the case of rapid 

isotropic (spherical) compression when the turbulent velocities 

are much smaller than the mean sound speed. Simulation results 

and linear theory were used to improve the two-equation model 

so that it correctly predicts the flow history. Comparisons of the 

rates of dissipation of turbulent kinetic energy given by the 

standard two-equation model, the simulation data, and the 

improved two-equation model are shown in the figure. We find 

that changes in the fluid viscosity, caused by the temperature 

increase induced by the compression, must be included in the 

turbulence model equations to correctly predict the development of the turbulence. 

Significance 

Models of the turbulence in compressible fluids can help predict 

astrophysical phenomena, and help in the design of internal 

combustion engines, hypersonic flight vehicles, and supersonic 

combustors. Unfortunately, the accuracy of the turbulence 

models is, at times, less than acceptable. By adjusting the models 

so that they are correct for idealized flows, the models will be 

more accurate when applied to more general flows. 

Future Plans 

We will continue model testing by considering new combinations of compression histories and flow parameters and attempt 

to extend the accuracy range of the models.

Publications 

1. Coleman, G. N. and Mansour, N. N. "Modeling the Rapid 

Spherical Compression of Isotropic Turbulence." Physics of 

Fluids A. 3, no.9 (Sept.] 991): 2255-2259. 

2. Coleman, G. N. and Mansour, N. N. "Simulation and 

Modeling of Homogeneous Compressible Turbulence under 

Isotropic Mean Compression." Presented at the Eighth 

Turbulent Shear Flows Conference, Munich, GE, Sept. 1991. 

3. Mansour, N. N. and Wray A. A. "Direct Numerical Simulation Data: A Tool for Turbulence Modeling." Presented at the 

Fourth International Symposium on Computational Fluid 

Dynamics, Davis, CA, Sept. 1991. 

Comparisons of the rates of dissipation of turbulent kinetic 

energy given by the standard two-equation model, the simulation data, and the improved two-equation model.

125
Flow-Field Calculations using Unstructured Grids 

David L. Marcum, Principal Investigator 

Co-Investigators: Ramesh K. Agarwal and David W. Halt 

Mississippi State University/McDonnell Douglas Research Laboratories 

Research Objective 

To develop a computational capability using adaptive unstructured grids to predict unsteady store separation and related flow 

fields. 

Approach 

The Reynolds-averaged Navier—Stokes equations are solved on 

an unstructured grid of tetrahedral elements. Space discretization 

is obtained from a Galerkin-weighted residual approximation. 

Time discretization uses either an explicit two-step Lax—Wendroff 

scheme, or an explicit multistage Runge—Kutta scheme. Boundary conditions are implemented using a method-of-characteristics 

based procedure. Turbulence effects can be modeled using a 

two-equation k-c turbulence model. An advancing-front gridgeneration scheme is used to generate solution-adaptive 

unstructured grids. The overall solution procedure has been 

initially validated for two- and three-dimensional inviscid and 

viscous flow fields. 

Accomplishment Description 

The finite-element Navier—Stokes code, MDFENS, was modified 

to include the capability to calculate unsteady flow fields and 

store trajectories for cases with multi-body separation. For these 

cases, the unstructured grid changes in time. Surrounding each 

body is a region of elements that move rigidly with the body. 

Between bodies, elements deform when there is relative motion. 

Periodically, a new solution-adaptive grid is generated to 

minimize distortion of deforming elements. An unsteady inviscid 

calculation of a hypersonic vehicle dispensing a store was 

performed to evaluate the overall procedure. For this configuration, the dispensing vehicle is considerably larger than the

separating store and the flow field near the store can be computed separately. The steady flow field about the dispensing 

vehicle was first calculated without the store. This flow field was 

used to determine inflow conditions for time dependent calculation of the flow field near the separating store. The store 

trajectory was computed from a time integration of the aerodynamic forces acting on the separating store. A typical calculation 

with 400,000 elements requires 14 megawords of memory and 

0.5-1 Cray Y-MP hour per grid remeshing cycle. A complete 

trajectory can require 40 Cray Y-MP hours. 

Significance 

Predicting unsteady flow fields with multi-body separation is 

important to many military and aerospace programs. 

Future Plans 

Calculations of flow fields about vehicles with a separating store 

will be validated with experimental data. The solution-adaptive 

grid-generation procedure with a local remeshing capability will 

be incorporated within the flow solver. Development of techniques to improve the computational efficiency of the present 

procedure for unsteady viscous-flow calculations will continue. 

Publications 

1. Marcum, D. L. and Agarwal, R. K. "Finite-Element 

Navier—Stokes Solver for Unstructured Grids." AIAA Journal 

30, no. 3 (1992): 648-654. 

2. Marcum, D. L. and Agarwal, R. K. "A Three-Dimensional 

Finite-Element Navier—Stokes Solver with k-c Turbulence 

Model for Unstructured Grids." AIAA Paper 90-1652, 1990. 

Adapted surface-mesh and instantaneous-density contours for unsteady hypersonic flow about a separating store. 

126
Space Shuttle Flow Field 

Fred W. Martin, Jr., Principal Investigator 

Co-Investigators: Pieter Buning, Steve Labbe, Ray Gomez, Jeff Slotnick, Steve Parks, and Max Kandula 

NASA Johnson Space Center/NASA Ames Research Center 

Research Objective 

To improve the numerical modeling of the Space Shuttle launchvehicle-ascent aerodynamic environment by adding multiplespecies plume capability to the flow solver. 

Approach 

The OVERFLOW—Chimera scheme is used to obtain flow fields 

about the Space Shuttle launch vehicle in Mach number range 

0.6 to 4.5, with particular emphasis in the transonic range. The 

accuracy of the redesigned solid rocket motor (RSRM) and Space 

Shuttle main engine (SSME) plume simulation is improved by 

adding the capability to convect multiple gas species. 

Accomplishment Description 

The JANNAF plume programs RAMP and SPF were used to 

define the RSRM nozzle exit conditions and plume properties for 

a Space Shuttle trajectory condition at Mach 1.25. These results 

are axisymmetric and do not include the influence of the Space 

Shuttle launch vehicle, thus they serve as the baseline plume 

definition for an undisturbed plume. The OVERFLOW program 

was modified to include high-temperature gas effects through the 

solution of a species convection equation for each flow-field gas. 

Gases are modeled as chemically frozen and are characterized 

as thermally perfect and calorically imperfect. The specific heat 

ratio, y, at each mesh point is mass averaged from the concentrations of constituent gases in the flow. In the loosely coupled 

approach, y is used in the global-equation set to incorporate realgas effects. The internal RSRM nozzle flow field was obtained 

from the RAMP code. The nozzle exit solution was then used 

as the plume start-line boundary condition in OVERFLOW. 

Downstream from the nozzle, computed external plume flowfield properties, including shear layer development and shock 

structure, were compared to SPF results. 

Significance 

The Space Shuttle launch-vehicle aerodynamic loads are 

strongly influenced by the RSRM and SSME plume/flow-field 

interactions. In particular, the RSRM plume acts to increase the 

pressure on the external tank base, and the Orbiter lower-aft 

fuselage/wing area. The correct modeling of these effects is 

instrumental in obtaining accurate simulations of the Space 

Shuttle ascent aerodynamic loads. 

Future Plans 

The improved plume-model capability will be evaluated by 

completing a three-dimensional calculation of the Space Shuttle 

with RSRM plumes with and without real-gas effects. Comparisons of hot-gas plume wind tunnel data and flight measurements 

from Columbia will be used to check the accuracy of the new 

plume model. 

SPF12 1-Phase Frozen Chemistry 

lnvscsd core solution wI turbulent mixing layer overlay 

OVERFLOW Perfect Gas (Gamma=1.4) GOOD 

988 lnv?scd FlowfieIc 

500011 

00000 3 80 DOE 300031 

39J01133 -- - 00000

OVERFLOW Variable Gamma (Solution of Species Convection Equations) 

lr?v,sc,d FlowlielO 

Comparison of axis ymmetric redesigned solid-rocket-motor solutions.

127 
Multigrid Solution of the Euler Equations 

Dimitri J. Mavriplis, Principal Investigator 

ICASE/NASA Langley Research Center 

Research Objective 

To develop an accurate and efficient method for computing 

steady-state compressible flow about complex three-dimensional 

configurations. 

Approach 

The steady-state three-dimensional Euler equations are solved on 

an unstructured tetrahedral mesh by a Galerkin finite-element 

scheme. The flow variables are stored at the vertices of the mesh, 

and an edge-based data structure is employed to minimize 

memory requirements. An unstructured multigrid technique is 

employed to accelerate convergence to steady state. This 

procedure operates on a sequence of non-nested coarse and fine 

meshes and the patterns for interpolation between the various 

meshes of the sequence are determined in a preprocessing step 

using an efficient search algorithm. For large applications, which 

require a significant portion of the Cray Y-MP main memory, a 

multitasked version of the solver has been developed. 

Accomplishment Description 

The present methodology is capable of providing accurate and 

efficient solutions for steady-state inviscid three-dimensional 

flows using very fine meshes. The accompanying figure illustrates 

the computation of transonic flow over a wing—body—nacelle 

configuration. A multigrid sequence of four meshes has 

been employed. The finest mesh of the sequence contains 

804,000 vertices and 4.5 million cells. The final solution was 

obtained in 100 multigrid cycles on the finest grid, during which 

the residuals were reduced by six orders of magnitude. This run 

required 96 megawords of memory and 16 Cray Y-MP minutes 

running in a dedicated mode on all eight processors. A computational rate of 750 MFLOPS was achieved using eight processors, 

which is 7.5 times faster than the single-processor. 

Significance 

The ability to accurately and efficiently predict compressible 

flows over complex three-dimensional geometries is important to 

the aircraft industry. The simultaneous use of multitasking and 

an efficient multigrid strategy enable the solution of large, 

complex problems in a matter of minutes.

Future Plans 

Future work will center on the inclusion of more adequate 

surface-modeling and grid-generation techniques, as well as the 

extension of the present work to viscous turbulent-flow cases in 

three dimensions, which will require the implementation of a 

multiple-field-equation turbulence model. 

Publication 

Mavriplis, D. J. "Unstructured Multigrid for the Three-Dimensional 

Euler Equations." AIAA Paper 91-1549, 10th AIAA Computational Fluid Dynamics Conference. Honolulu, HI, June 1991. 

Computation of transonic flow over a wing—body—nacelle 

configuration. 

128
atlon 

High Mach-Number Mixing in a Shock Tunnel Environment 

Charles R. McClinton, Principal Investigator 

Co-Investigator: Robert D. Bittner 

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

Research Objective 

To provide pre- and posttest computational fluid dynamics (CFD) 

support to aid in the understanding of the first series of highspeed scramjet-combustor tests using the Ames 16-Inch Shock 

Tunnel. In the process, more confidence in the ability of CFD 

to model complex hypersonic-injection flow fields will be 

developed. 

Approach 

The SPARK codes were used to model the flow field in this high 

Mach-number environment. A set of three-dimensional elliptic 

and parabolic solvers has been extensively calibrated and 

applied to scramjet .combustor problems. The model tested in 

the Ames 16-Inch Shock Tunnel was a wedge angled at 

11 degrees to a Mach 5 free stream with hydrogen injected at 

30 degrees from the model surface. Turbulence was modeled 

using a standard Baldwin—Lomax algebraic approach. Both 

mixing and reacting cases were solved. A seven-species, sevenreaction mechanism was used to model the kinetics in the 

problem. CFD results were compared with all available data 

from the tests including wall pressures, shadowgraphs, and 

finite-fringe interferograms. 

Accomplishment Description 

The computational work has been completed for both mixing 

and reacting analyses of the wedge-model tests. The results 

compared favorably with the available test data. The accompanying figure shows a comparison of CFD laterally averaged 

density versus a finite-fringe interferogram. This comparison 

shows that the basic flow physics, such as the wedge leadingedge shock, the injection bow shock, and the fuel-mixing layer, 

qualitatively agree. The fuel penetration height is indicated on 

the CFD contours and the interferogram. This agreement verifies 

facility nominal-operating conditions. In addition, the confidence 

in the ability of CFD to accurately predict scramjet-combustor 

flow fields has been increased in terms of surface loads and 

overall flow physics. A typical case requires 20 megawords of 

memory and 2-4 Cray-2 hours. 

Significance 

These studies are being used in conjunction with other highenthalpy experimental programs in an effort to provide a much 

needed Mach 10-16 data base for validation of hypersonic 

scramjet-combustor performance. Furthermore, CFD enables a 

detailed examination of the flow field, including an ability to 

quantify the mixing, combustion, and losses in a scramjet 

combustor. 

Future Plans 

Future tests are planned with more complex geometries and 

higher total reservoir pressures resulting in a test more representative of the hypersonic environment. An evaluation of this 

combustor will be analyzed with SPARK. The CFD results will be 

used to assess the combustor efficiency and flow losses. 

Comparison of computational fluid dynamics laterally averaged density contours with a finite-fringe interferogram.

129
National Aero-Space Plane Configuration Trade Studies 

Charles R. McClinton, Principal Investigator 

Co-Investigators: Arthur D. Dilley and Richard W. Hawkins 

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

Research Objective 

To perform trade studies for the National Aero-Space Plane 

(NASP) program using computational fluid dynamics (CFD) 

codes. These studies involve the airframe, propulsion system, 

and airframe—propulsion system integration. The results will 

optimize the overall NASP design and individual elements of 

the vehicle. 

Approach 

We used calibrated CFD codes to analyze external and internal 

flow fields for NASP vehicles and vehicle components. The 

codes are being coupled with linear-stability codes to study 

boundary-layer transition on the NASP vehicle. This is a new 

area of application for current CFD codes; insights into computational requirements for mean-flow and linear-stability calculations are being gained. 

Accomplishment Description 

Using CFD, we analyzed the forebody of an alternative NASP 

configuration at several flow conditions. The integrated-mass 

capture and other flow-field data were used to determine the 

viability of the alternative NASP configuration. We performed a 

parametric study on a hypersonic-flight vehicle to optimize the 

viability of several experiments located on the flare portion of 

the vehicle. The parametric study considered several flare 

lengths and analyzed the flow field immediately ahead of the 

experiments. The accompanying figure shows the flow-field 

structure including the bow-shock/ramp-shock intersection, the 

resulting slip line, the boundary-layer edge, and the inlet 

experiment location. A baseline vehicle geometry was developed using the results from this parametric study. We initiated 

mean-flow calculations for linear-stability code analysis. These 

calculations indicated a need for unusually restrictive computational requirements when performing mean-flow calculations. 

These computational requirements are from the linear-stability 

code requirement for smooth first and second derivatives of 

flow-field quantities. The net result of all these requirements is a 

highly accurate resolution throughout the boundary layer. An 

average job requires 15 megawords of memory and 2-4 Cray-2 

hours.

Significance 

We can apply current CFD codes to solve design problems 

associated with NASP and related hypersonic-flight vehicles. 

Future Plans 

We will continue to perform trade studies for the NASP program 

and develop the capability to provide accurate mean-flow 

computations for linear-stability analysis. 

I.

X 

Computational fluid dynamics applications for vehicle design. 

Mach number contours on a biconic configuration at Mach 15; 

red = high values, blue = low values. The inlet is located where 

the flow is completely turbulent and sized by the location of the 

slip line. 

130
Scramjet Inlet Interaction 

Charles R. McClinton, Principal Investigator 

Co-Investigators: Thomas N. Jentink and Arthur D. Dilley 

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

Research Objective 

To perform an inlet-unstart computational study on a full-scale, 

two-dimensional inlet (no sidewall compression) with short, setback splitter plates to determine whether the unstarted module 

will cause unstart in neighboring modules. 

Approach 

The study was performed using the computational fluid dynamics code CFL3D. First, we performed two-dimensional, turbulent 

calculations of the full-scale configuration, starting with the 

wedge-type forebody and extending through the inlet to the 

module's exit. This solution was used to study several techniques 

for causing the inlet to unstart. The localized-heat addition was 

chosen to allow adaptation to the changing mass flow rate during 

the unstart process. Next, we obtained a three-dimensional 

steady state, started a solution for a half module, and began 

the unstart sequence by unstarting an adjacent half-module. 

Additional blocks were added to the calculation as the unstart 

progressed. 

Accomplishment Description 

The problem had 1.5 modules comprised of 15 blocks and 

approximately 1.3 million points. The unstart shock moved 

10 inches upstream from the splitter-plate leading edge, then 

slowed its upstream movement as it spilled approximately 25% 

of its mass into the started module. The flow in the started 

module was highly distorted and had begun to spill a small 

amount of mass into what would have been another module 

had an additional block been added. At this point, it was 

determined that the module was unstarted and the computation 

was terminated. This study provided an excellent qualitative look 

at module-to-module interaction during inlet unstart. Approximately 450 Cray-2 hours (run in 4 to 6 hour segments) were 

used. Typical runs required about 55 megawords of memory. 

The accompanying figure shows a cowl-side view within the 

module region of both the started and the uristarted solutions. 

Pressure contours are shown, with blue denoting the lowest 

pressure, and red the highest pressure.

Significance 

Inlet unstart is an important issue in the design of scramjet 

propulsion for the National Aero-Space Plane (NASP). The 

performance of each inlet design depends on the types of 

interactions occurring between the inlet modules. If one module 

unstarts, it has a large effect on inlet and aerodynamic loads and 

may unstart neighboring modules. This study provides a picture 

of the unstart process of this inlet design and, in conjunction 

with experimental studies, will result in a better understanding of 

inlet unstart in scramjet propulsion systems. 

Future Plans 

This work has been completed and the results have been 

reported to the appropriate NASP program participants. 

-

;&•- Started solution 

-

.'- 

Pressure 'IL Y a'C Solution 

High 

Low

Tn. .70 F 

P,.,e .43 P80

Detail of a scramjet engine unstart.

131 
Scramjet Engine Design Optimization 

Charles R. McClinton, Principal investigator 

Co-Investigators: Pradeep S. Kamath and Marion Mao 

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

Research Objective 

To develop an understanding of the high-speed (Mach > 8) 

scramjet-combustor performance related to specific combustor 

design constraints in order to design effective high-speed 

scramjet combustors and develop a consistent parametric 

numerical database for potential high-speed scramjet combustor 

designs. 

Approach 

We will develop a numerical data base using the previously 

calibrated SHIP three-dimensional parabolized Navier—Stokes 

design code. The test matrix is generated using an advanced testdesign process. First, important design variables and engine 

performance response parameters are determined. A statistical 

test-design method is used to minimize the number of solutions 

required to represent the important design degrees of freedom for 

the full-factorial parametric design space. Response parameter 

results from the numerically generated data base are evaluated 

to determine trends and to develop analytical models. The 

analytical models are used to select better configurations for 

additional in-depth study. 

Accomplishment Description 

We completed about 150 large-scale scramjet combustor flowfield solutions for two of the three fuel-injector classes being 

studied. The solutions represent about 5% of the full-factorial test 

matrix for the two classes. Design parameters included injector/ 

combustor geometric issues, flight Mach number, and fuelequivalence ratio. The test matrix provided for linear, nonlinear, 

and interaction terms in regression fits of the numerical data 

base. The accompanying figure illustrates a typical combustor 

solution, including fuel concentration contours on four flow-field 

cross sections and heat flux on one wall. The resulting numerical 

data bases were used to generate analytical models for fuelmixing efficiency, fuel-injector drag, combustor-entropy rise, 

heat load, peak-heat flux, shear force, and effectiveness. Also 

shown is a regression fit for combustor effectiveness for the 

Mach 18 solutions. These models compare favorably with existing 

experimental results and Navier—Stokes solutions. Atypical run 

requires 1 megaword of memory and 1-4 Cray-2 hours. 

Significance 

The first numerical-design data base for high-speed scramjet 

combustors was developed. This data base enhances the 

understanding of scramjet performance issues such as trades 

between combustor injector design and mixing/combustion 

performance, combustion efficiency and thrust produced, and 

mixing and combustor heat load. The analytical models provide 

a method of engine design optimization, which has already 

produced a significant increase in combustor performance for 

the National Aero-Space Plane designs.

Future Plans 

We will complete our evaluation and verification of the numerical data base with Navier—Stokes solutions using the SPARK 

and General Aerodynamic Simulation Program codes. We will 

select a single-combustor performance-response parameter, 

extend the range of parametric variables for configurations on a 

boundary test matrix, and recommend direction for additional 

parametric studies. 

Inlet flow 

lnjection 

station

Combustor 

wall segment Combustor 

(a) exit 

A typical scramjet-combustor solution at Mach 15 with heat flux 

on a wall. The fuel-concentration planes show characteristic 

kidney shapes caused by the ramp-induced vortices; red = high 

values, blue = low values. 

0 

(D 

0. 

LU 

C)

1 2

CE (actual) 

Combustor effectiveness (rICE) regression at Mach 18. 

132
Aerodynamics and Acoustics of Rotorcraft 

W. J. McCroskey, Principal Investigator 

Co-Investigators: J. D. Baeder, V. Raghavan, G. R. Srinivasan, and S. K. Stanaway 

U.S. Army Aeroflightdynamics Directorate, AVSCOM/NASA Ames Research Center 

Research Objective 

To develop and validate accurate, user-oriented, viscous, 

computational fluid dynamics codes (with inviscid options) for 

three-dimensional unsteady aerodynamic flows about arbitrary 

rotorcraft configurations. The ability to calculate complex 

vortical wakes, shock waves, rotor—body interactions, and the 

associated acoustic field will be included. 

Approach 

Advanced three-dimensional unsteady multizone implicit Euler 

and Navier—Stokes codes are used to simulate rotorcraft aerodynamics and acoustics under new flight conditions that have not 

previously been treated satisfactorily. 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. Rotorcraft airframes are treated using the OVERFLOW code with Chimeraoverlapped zonal-grid technology, which is well-suited to 

moving interfaces between fixed and rotating grids. In addition, 

special solution-adaptive grid-clustering and wave-fitting 

techniques are used to capture low-level radiating acoustic 

waves and trailing tip vortices. 

Accomplishment Description 

Important results were obtained for high-speed impulsive noise. 

The hover results form a large numerical data base which is 

being used by theoretical acousticians to verify the soundness 

of the acoustic analogy approach. Results have also been 

obtained for the V-22 Osprey fuselage alone and the wing—body 

combination at three angles of attack using the OVERFLOW 

Navier—Stokes code. The results were compared with experimental airloads and potential-flow calculations. The accompanying figure shows the color-coded surface pressures and 

flow particle traces at 10 degrees. The Navier—Stokes results 

showed separation on the lower rear of the fuselage and at 

the sponson—fuselage junction. The calculated lift for the 

fuselage—wing combination agrees with the experimental value 

to within about 10%. These results will be refined during the 

next year, and the nacelles, rotors, and empennage will be 

added. 

Significance 

The ability to accurately simulate the aerodynamics and 

acoustics of rotorcraft will allow quieter vehicles to be designed 

at lower cost and with less risk. Analysis of the flow separation 

on the Osprey V-22 airframe will permit drag-reduction 

improvements to be investigated and implemented at a fraction 

of the cost of wind tunnel studies.

Future Plans 

We will develop more accurate and efficient simulations 

of rotorcraft aerodynamics and acoustics, particularly on 

blade—vortex interaction noise and rotor—body aerodynamic 

interference on tilt-rotor aircraft. 

Publications 

1. Baeder, I. D. "Euler Solutions to Nonlinear Acoustics of NonLifting Rotors in Forward Flight." American Helicopter 

Society/Royal Aeronautical Society International Technical 

Specialists' Meeting. Philadelphia, PA, Oct. 1991. 

2. Srinivasan, G. R. and Baeder, J. D. "Recent Advances in 

Euler and Navier—Stokes Methods for Calculating HelicopterRotor Aerodynamics and Acoustics." Fourth International 

Symposium on Computational Fluid Dynamics. Davis, CA, 

Sept. 1991. 

Color-coded surface pressure and particle traces on the Osprey 

V-22 wing—fuselage simulation in cruise; Mach = 0.2, Reynolds 

number = 106, a = 10 degrees.

133
High-Performance Rotor-Blade Tips 

W. J. McCroskey, Principal Investigator 

Co-Investigators: E. P. N. Duque, V. Raghavan, G. R. Srinivasan, and R. C. Strawn 

U.S. Army Aeroflightdynamics Directorate, AVSCOWNASA Ames Research Center 

Research Objective 

To compute the viscous three-dimensional unsteady flow field 

around an advanced helicopter rotor. Particular emphasis is 

placed on viscous transonic aerodynamics of high-performance 

rotor-blade tips and resolution of the vortical wake details of 

rotors in hover. 

Approach 

The unsteady three-dimensional Euler/Reynolds-averaged 

Navier—Stokes equations are solved by structured- and 

unstructured-grid methods, either in blade-fixed or inertial 

reference systems. The codes are validated through detailed 

comparisons with experimental data. 

Accomplishment Description 

The transonic unsteady-rotor Navier—Stokes (TURNS) structuredgrid code was improved and applied to the flow fields of 

complex multibladed lifting rotors in hover. The results show 

good agreement with experimental data for surface pressures, 

thrust, power, and near-field vortex trajectory. However, the 

detailed structure of the vortex core is difficult to resolve with the 

single-block formulation. The first unstructured-grid rotor-blade 

solution was obtained with an explicit upwind Euler code with a 

locally refined grid. The upper section of the accompanying 

figure shows the grid at the boundary after two refinements in 

the wake; the lower section shows color-coded vorticity 

contours arising from the convected tip vortex. Excellent 

agreement with experimental airloads data was obtained, 

although further solution-adaptive grid refinement in the wake 

seems to be necessary. 

Significance 

The TURNS code is efficient and robust for computing airloads 

on helicopter blades. The feasibility of locally refining the grid in 

the wake in order to improve the resolution of the vortical wake 

structure by means of a solution-adaptive unstructured-grid code 

was demonstrated for the first time. 

Future Plans 

Chimera multiblock overlapped zonal-grid technology will be 

incorporated into the TURNS structured-grid code for improved 

wake computations and for rotor—body interactions. The 

solution-adaptive unstructured-grid code will be applied to 

acoustic-wave computations. It will be improved for tip-vortex 

simulations and extended to forward flight.

Publications 

1. Srinivasan, G. R.; Raghavan, V.; and Duque, E. P. N. "FlowField Analysis of Modern Helicopter Rotors by Navier—Stokes 

Method." American Helicopter Society/Royal Aeronautical 

Society International Technical Specialists Meeting. 

Philadelphia, PA, Oct. 1991. 

2. Strawn, R. C. and Barth, T. J. "A Finite-Volume Euler Solver 

for Computing Rotary-Wing Aerodynamics on Unstructured 

Meshes." American Helicopter Society 1992 Annual Forum. 

Washington, DC, June 1992. 

261,133 Nodes 1,448,336 Tetrahedra 24,434 Boundary faces 

I

I7 - 

Grid and vorticity contours at the boundary of a hovering 

helicopter rotor blade; 261,133 nodes, 1,448,336 tetrahedra, 

and 23,434 boundary faces. 

134
Active Control of Ramjet Combustion Instability 

Suresh Menon, Principal Investigator 

QUEST Integrated, Inc. 

Research Objective 

To investigate techniques to control unstable combustion in 

ramjets using large-eddy simulations (LES). Numerical research 

is being conducted parallel to ongoing experimental studies. 

Approach 

The unsteady compressible Navier—Stokes equations, together 

with a thin-flame model for premixed combustion, are solved 

using an explicit fourth-order-accurate finite-volume scheme. 

The turbulent-flame speed appears explicitly in the combustion 

model as a function of the laminar-flame speed and the subgridturbulence kinetic energy. A one-equation model for the 

subgrid-turbulence kinetic energy is used in the code. 

Accomplishment Description 

The low-frequency, large-amplitude oscillation in the combustor 

is driven primarily by the inlet acoustic mode. However, in some 

cases a coupled convective—acoustic mode can also excite the 

instability. Results are in good agreement with experimental 

data. The spatial variation of the subgrid kinetic energy and the 

turbulent-flame speed is highly dependent on the presence of 

large vortical structures in the combustor. This is consistent with 

experimental observations of the large structures entraining the 

flame and wrinkling it, thereby increasing the local-flame speed. 

Secondary fuel-injection control strategies have been studied. 

When the controller is effective, the low-frequency oscillation 

is replaced by a high-frequency, low-amplitude oscillation. 

When the controller is only partially effective, the amplitude is 

reduced, but the frequency is unaffected. Similar to experimental 

observations, control of one frequency results in the excitation of 

other frequencies. Simulation of combustion instability required 

about 40 Cray-2 hours using a grid of 320 x 64 and approximately 8 megawords of memory. Each simulation of active 

control required about 15 Cray-2 hours. 

Significance 

The simulation provided detailed information on the behavior of 

unstable combustion in ramjet engines. The fully coupled flame 

and subgrid model can reproduce features observed in experimental configurations. Active control results are in good 

agreement with experimental observations and the LES results 

will be used to interpret experimental results.

Future Plans 

Active control using secondary fuel injection will be studied in 

more detail. More complex controllers for multiple frequencies 

will be investigated. The effects of flame curvature and local 

extinction will be incorporated to generalize the LES capability. 

Finally, fully three-dimensional LES will be done to investigate 

the importance of three-dimensional vortex stretching on the 

instability mechanism. 

Publications 

1. Menon, S. "Active Combustion Control in a Ramjet using 

Large-Eddy Simulation." Comb. Sci. and Tech. 84 (1992): 

51-79. 

2. Menon, S. "A Numerical Study of Secondary Fuel-Injection 

Techniques for Active Control of Combustion Instability in 

Ramjets." AIAA Paper 92-0777, 30th Aerospace Sciences 

Meeting, Reno, NV, 1992. 

Combustion product

FWM 

fuel

1111V 7 

.10 1(c) 

.65 .75 75 aG as 95 95 1.00 1.05 1.10

0(m) 

Combustion instability in a ramjet engine; (a) flame structure, 

(b) vortex structure, and (c) subgrid kinetic energy.

135
0 .020 .040 .060 .080 

X 

Y .040 

.020 

.080 

.060 

Bluff-body stabilized plasma in a quartz tube at power 

level = 1 kW. 

13000 

12000 

w 

cc

11000 

:1 10000 

- 9000 

8000

1 k 

2

O Experimental data 

A

Computational data 

- I I I I I 

Incident microwave power (kW) 

Coupled Navier—Stokes Maxwell Analysis for Microwave Propulsion 

Charles L. Merkle, Principal Investigator 

Co-Investigator: S. Venkateswaran 

Pennsylvania State University 

Research Objective 

To model microwave plasmas in experimental configurations. 

Comparisons and parametric studies are being performed to 

establish the validity of the model and to understand experimental 

trends. 

Approach 

Coupled solutions of Navier—Stokes equations for gas dynamics 

and Maxwell equations for microwaves are obtained by finitedifference approximation of the equations. The Navier—Stokes 

equations are preconditioned for efficient computation of the 

low-Mach-number flows, while the Maxwell equations are 

solved by an explicit time-accurate procedure. 

Accomplishment Description 

Numerical computations have been performed for resonant 

cavity and waveguide configurations. For both cases, experimental conditions were duplicated and detailed comparisons of 

plasma temperature, coupling efficiency, and overall thermal 

efficiency were made to establish the accuracy of the computational results. Nearly 40 different cases were computed to 

perform detailed parameter surveys. Grid-resolution studies 

were carried out to verify solution accuracy.

Significance 

Computational results demonstrate that resonant-cavity plasmas 

show excellent stability and high coupling efficiencies (up to 

99%) over a wide range of powers. Waveguide plasmas with a 

reflecting end wall also show good plasma stability, but the 

coupling efficiency drops almost linearly as incident power is 

increased. 

Future Plans 

We will study the effect of asymmetries in the power coupling to 

the cavity, model realistic microwave-thruster configurations, 

and investigate size scale-up issues. 

Publications 

1. Venkateswaran, S. and Merkle, C. L. "Numerical Investigation of Bluff-Body Stabilized Microwave Plasmas." AIAA 

22nd Fluid Dynamics, Plasmadynamics, and Lasers Conference, Honolulu, HI, June 1991. 

2. Venkateswaran, S. and Merkie, C. L. "Numerical Modeling 

of Waveguide-Heated Microwave Plasmas." Symposium of 

Multi-Disciplinary Applications of CFD, ASME Winter 

Annual Meeting, Atlanta, GA, Dec. 1991

Peak plasma temperature versus incident microwave power for 

different incident power levels. Experimental data are included 

for comparison. 

136
Turbulent Flow Over Riblets 

Parviz Mom, Principal Investigator 

Co-Investigators: John Kim and Haecheon Choi 

Stanford University/NASA Ames Research Center 

Research Objective Future Plans 

To study the mechanics of turbulent drag reduction by riblets. We will analyze the detailed flow field over riblets using flowThe results of this investigation could lead to the design of better visualization techniques and motion pictures of the flows. 

riblet configurations leading to more drag reduction.

Publication 

Approach Choi, H.; Mom , P.; and Kim, J. "On the Effect of Riblets in Fully 

Direct numerical simulations (DNS) are used to simulate Developed Laminar Channel Flows." Phys. Fluids A. 3 (1991): 

turbulent flows over longitudinal riblets. A fully implicit finite- 1892. 

difference method is used to solve unsteady incompressible 

Navier—Stokes equations in a generalized coordinate system. S = 40 

Accomplishment Description 

We have computed the flow in a channel with riblets on one 

wall and have considered four cases, S = 20 and S = 40, where 

S is the spacing of the riblets in wall units, and ridge angles are 

45 and 60 degrees. The calculations with S = 40 and S = 20 

have converged and show about 10% drag increase and 5% 

drag decrease, respectively. This is in agreement with the 

experimental findings. Wall-shear rates on most regions of the 

cross-sectional perimeter of riblets are smaller than those of the 

corresponding plane turbulent-channel flow. It appears that the 

drag reduction by riblets is caused by alteration of the location 

of the streamwise vortices with respect to the wetted surfaces 

(see the accompanying figure). With riblets of proper dimensions, streamwise vortices are located farther away from the 

wetted surfaces than in the flat-plate boundary layer and 

produce fewer high skin-friction regions. The total time per 

time step on the Cray Y-MP was 1 minute. Thus, with S = 20, 

approximately 200 Cray Y-MP hours were required to integrate 

500 nondimensional time units. The required core memory and 

disk scratch spaces were 10 megawords and 40 megawords, 

respectively. 

Significance 

The drag variation of this study with the spacing and height of 

the riblets matches well with the experimental results. The 

difficulty in the measurement of velocity fields near riblets and 

the modest change in skin friction and turbulence statistics have 

prevented a complete understanding of the mechanics of drag 

reduction by riblets. The data bases generated by DNS are of 

considerable value for the quantitative and qualitative studies of 

the structure of turbulence over riblets.

Z

S 20 

Velocity vector and contours of streamwise vorticity in a 

(z, y) plane.

137
Turbulence Over a Backward-Facing Step 

Parviz Mom, Principal Investigator 

Co-Investigators: John Kim and Hung Le 

NASA Ames Research Center 

Research Objective 

To perform a direct simulation of turbulent flow over a backward 

facing step with inflow and outflow boundary conditions, to 

generate a data base for modeling, and to investigate the physics 

of turbulent reattachment. 

Approach 

Three-dimensional time-dependent simulations of incompressible turbulent flow are performed using a second-order finitedifference method on a staggered mesh and a semi-implicit 

fractional-step time advancing. The inflow fluctuations are 

imposed using broadband spectra and are scaled to match all 

second-order statistics of the channel or boundary-layer flow. 

Convective boundary condition is applied at the exit, and 

periodic boundary condition is imposed in the spanwise 

direction. Based on the step height, the Reynolds number (Reh) 

is about 5,000. Up to 770 x 194 x 66 grid points are used. 

Accomplishment Description 

The numerical results are in good agreement with experiments. 

The calculated reattachment length is found to be a strong 

function of the expansion ratio, which agrees with past experimental results. Strong negative skin friction is also seen in the 

back-flow region. When plotted in wall coordinates, the mean 

velocity profiles in the recovery region are below the log-law of 

the wall. This may be caused by the adverse pressure gradient 

created by the sudden expansion. The accompanying figure 

shows the streamwise velocity contours on a horizontal plane at 

y = 16. The superimposed black lines indicate the reattachment 

locations. The typical direct numerical simulation requires about 

2,000 Cray Y-MP hours and 10 megawords of memory. 

Significance 

The numerical simulations were a significant step for direct 

computation of complex turbulent flows with turbulent inflow/ 

outflow conditions. The simulations have provided a comprehensive database for Reynolds-averaged modeling of complex 

flows. 

Future Plans 

The simulations with fine-grid resolution are being continued 

and flow statistics, including all the terms in the Reynolds-stress 

budgets, will be computed. 

I tz 

POW - 

-(-

-.

- . - - -•---. -.-

Streamwise velocity contours at y = 16 for Reh = 5, 100. The step is at the left boundary. The mean flow is from left to right. Dimensions = 20 height x 40 height. Velocity range: blue = —0.4 to magenta = 0.5. 

138
Aerodynamic Sound Generation 

Parviz Mom, Principal Investigator 

Co-Investigators: Sanjiva K. Lele, Tim Colonius, and Brian E. Mitchell 

Stanford University/NASA Ames Research Center 

Research Objective 

To compute acoustic sources and far-field sound using the 

unsteady Navier—Stokes equations to allow direct validation of 

aeroacoustic theories and help provide noise control strategies 

Approach 

Accurate computation of far-field sound along with near-field 

hydrodynamics requires that the Navier—Stokes equations be 

solved using accurate numerical differentiation and timemarching schemes with nonreflecting boundary conditions. 

Sixth-order compact finite-difference schemes are used to 

evaluate spatial derivatives and time is advanced with a fourthorder Runge—Kutta scheme. Nonreflecting boundary conditions 

developed for the two-dimensional linearized Euler equations 

are modified for use with nonlinear Navier—Stokes computations 

of open flow problems. 

Accomplishment Description 

The numerical scheme was tested on several model problems, 

including sound radiation in uniformly sheared viscous flow, 

scattering of sound waves by a compressible vortex, sound 

generation by a pair of corotating vortices, and the spatial 

evolution of a compressible mixing layer. The accompanying 

figure summarizes the results of the sound produced by 

corotating vortices. The sound level measured in the far field of 

the direct Navier—Stokes computations agrees well with predictions based on an aeroacoustic theory. In computations of 

spatially evolving free-shear flows, large-scale vortical structures 

are found to produce large reflections at the outflow boundary 

due to nonlinear effects; these reflection errors cannot be 

improved by increasing the accuracy of the linear boundary 

conditions. The addition of an exit zone just upstream from an 

outflow, where disturbances are significantly attenuated through 

grid stretching and filtering, reduces reflections from vortical 

structures by three orders of magnitude. The figure shows the 

improvement of the dilatation field associated with the rollup of 

a compressible mixing layer. The model problems typically use 

from 40 Cray Y-MP hours for each mixing-layer test to about 

100 Cray Y-MP hours for the corotating vortices. The solid state 

device is used; therefore, core memory is typically less than 

4 megawords. 

Significance 

Computational aeroacoustics research is in its infancy and 

successful test problem simulation, including excellent agreement with theories and experiments, shows that accurate direct 

computation of fluid flow, with sound sources and far-field sound 

whose energy is 10 orders of magnitude smaller than the flow 

energy, is feasible. 

Future Plans 

The codes developed will be used to compute the sound 

generation in two- and three-dimensional turbulent shear flows.

I. ON 

Directly computed far-field pressure from a corotating vortex 

pair (left). Near-field vorticity illustrates the corotating vortices 

generating the sound. 

Spurious acoustic reflections at the outflow boundary of a 

compressible mixing layer.

139
Shock-Wave/Turbulent-Bou ndary- Layer Interaction 

Parviz Mom, Principal Investigator 

Co-Investigators: Sarijiva K. Lele, Mahesh Krishnan, and Sangsari Lee 

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

Research Objective 

To understand the mechanisms of shock-wave/tu rbu lentboundary-layer interaction. Interactions of a shock wave with 

isotropic turbulence and turbulence subject to mean shear are 

investigated. 

Approach 

Direct numerical simulation (DNS) and linear analyses are 

performed. 

Accomplishment Description 

In the interaction of isotropic turbulence with a shock wave, 

turbulent kinetic energy is rapidly amplified at the immediate 

downstream of the shock. Transverse components of vorticity are 

amplified with the normal component unchanged. Turbulence 

length scales decrease during the interaction. Boundary conditions for the simulation of inhomogeneous turbulence interacting 

with a shock wave are being developed. Unsteady boundary 

conditions were developed for the simulation of interaction of 

turbulence with oblique shock waves by applying different 

nonreflecting boundary conditions in different zones. The 

accompanying figure shows a typical dilatation field of turbulence 

interacting with incident and reflecting shock waves. A typical 

three-dimensional DNS requires approximately 200 Cray Y-MP 

hours, 8 megawords of memory, and 40 megawords of disk 

scratch space.

Significance 

Since DNS resolves all the relevant time and length scales of 

turbulence and a shock wave, it can provide all the physical 

information that cannot be measured in experiments. Thus, the 

data base generated by these simulations provides key information for the development of turbulence models to predict the 

interaction of turbulent boundary layers with a shock wave. 

Future Plans 

DNS and linear analyses on the interaction of sheared turbulence with a normal shock wave will be performed. In order to 

investigate the interaction of a strong shock wave with turbulence in practical applications, large-eddy simulations (LES) will 

be performed without resolving the shock wave. The main 

components of the LES, such as shock-capturing schemes and 

subgrid scale turbulence models, will be validated by comparing 

them with the DNS results obtained in this work. 

Publications 

1. Lee, S.; Mom, P.; and Lele, S. K. "Interaction of Isotropic 

Turbulence with a Shock Wave." Report No. TF-52, Department of Mechanical Engineering, Stanford University, 

Stanford, CA, 1992. 

2. Lee, S.; Lele, S. K.; and Mom, P. "Simulation of Spatially 

Evolving Turbulence and the Applicability of Taylor's 

Hypothesis in Compressible Flow." Phys. Fluids A. 4, no. 7 

(1992). 

A typical dilatation field of turbulence interacting with incident and reflecting oblique shock waves. An oblique shock wave is incident 

on and reflected from the bottom boundary where a slip boundary condition is applied. Along the top, boundary conditions based on 

exact characteristics and nonreflecting boundary conditions are used in different zones. 

140
Wall-Bounded Turbulent Flows 

Robert D. Moser, Principal Investigator 

Co-Investigator: John Kim 

NASA Ames Research Center 

Research Objective 

To study the physics of wall-bounded turbulent shear flows, 

including Reynolds number effects, flow control, and coherent 

structures. 

Approach 

We performed direct numerical simulations (DNS) of turbulent 

flows in simple geometries and analyzed the resulting fields. 

Accomplishment Description 

Using short-time Lyapunov exponent analysis on a computed 

turbulent channel flow, it was found that there were several 

instabilities responsible for the chaotic nature of the turbulence. 

The instabilities include the Kelvin—Helmholtz rollup of vertical 

and inclined internal shear layers as well as the "flapping" of the 

inclined shear layers. Turbulent channel flow DNS with heat 

transfer at Reynolds = 700 has been performed. The small-scale 

structure of the turbulence has been examined by considering 

the three-dimensional velocity derivative field. It was found that 

the velocity and velocity-derivative spectra satisfy the local 

isotropy relations at high wave numbers. 

Significance 

The mechanisms responsible for the unpredictability of turbulence are important for the development of turbulence control 

strategies and the development of a turbulence theory. Local 

small-scale isotropy of inhomogeneous turbulent flow is often 

assumed in developing turbulence models, but this assumption 

has not been well tested. 

Future Plans 

The analysis of the simulations is continuing. Simulations of 

higher Reynolds-number flows and of compressible flows are 

planned.

Publication 

Antonia, R. A.; Kim, J.; and Browne, L. W. "Some Characteristics 

of Small-Scale Turbulence in a Turbulent Duct Flow." J. Fluid 

Mech. 233 (1991): 369. 

Color contours indicate the perturbation energy (top) and the 

production of perturbation energy (bottom) in a plane parallel to 

the wall. Contour lines indicate stream wise velocity. The 

perturbations are growing along the shear layer indicated by the 

velocity contours.

141
Hypersonic Rarefied Flow about a Compression Ramp 

James N. Moss, Principal Investigator 

Co-Investigator: Joseph M. Price 

NASA Langley Research Center 

Research Objective 

To establish the validity of the direct simulation Monte Carlo 

(DSMC) method for calculating viscous interacting flows and to 

investigate the sensitivity of such flows to cell resolution. 

Approach 

A two-dimensional DSMC code was used to calculate the flow 

about a flat plate followed by compression ramps. Ramp angles 

of 0-35 degrees were used, and the free-stream conditions 

were those obtained in low-density hypersonic wind tunnels at 

the Deutsche Forschungsanstalt for Luft- und Raumfahrt in 

Gottengen, Germany. For selected flow conditions, grid variation 

studies were made. 

Accomplishment Description 

Comparison of the calculated results with experiment is very 

favorable. The flat plate provides a rigorous test for numerical 

solutions and the DSMC method appears to provide satisfactory 

treatment. The same can be said of the DSMC method for 

merged-layer or strongly interacting flows in the presence of 

adverse pressure gradients and separated flows. The separatedflow case for the 35 degree ramp shown in the accompanying

figure used 335,000 simulated molecules and 22,500 cells. 

The solution required approximately 40 Cray Y-MP hours and 

4 megawords of memory. 

Significance 

The design of future hypersonic vehicles will heavily rely on 

numerical simulations to complement and extend information 

obtained in ground-based and flight experiments. The validation 

process begun in this study ensures that a tested method exists 

for calculating flap-control effectiveness and heating conditions 

at high altitudes. 

Future Plans 

Three-dimensional DSMC simulations will be conducted to 

account for edge-relief effects as they influence the separation 

and reattachment phenomena. 

Publications 

Moss, J. N.; Chun, C.-H.; and Price, J. M. "Hypersonic Rarefied 

Flow about a Compression Corner—DSMC Simulation and 

Experiment." AIAA Paper 91-1313, 1991. 

DENSITY (pip_) SKIN FRICTION 

10 

40 Begin Ramp - 8 0.10

Begin Ramp

 71.4 mm 

30

x=71.4mm 

20

•

C1 0.05 

mm

10 U 

0

0.00- 

0 25 50 75 100 0

0 25 50 75 100 

X, mm

X, mm 

STREAMLINES MACH NUMBER 220 

Begin Ramp 10 Begin Ramp

19.0 

16.0 

40 x= 71.4 mm-> 40 8

x= 71.4 mm ---' 30 

o.o 

y, 30 - y,30 

mm mm 20 

7r60

2 02 

5075160 

X, mm X, mm

Direct simulation Monte Carlo calculations for a 35 degree ramp at Mach 21.6 nitrogen flow with p,, = 3.9 x 10- 4 kg/m3 and a wall 

temperature of 341 K. 

142
Transonic-Shock/Boundary-Layer Interaction to Alleviate Separation 

Jon S. Mounts, Principal Investigator 

Co-Investigator: Thomas]. Barber 

United Technologies Research Center 

Research Objective 

To develop and apply three-dimensional numerical techniques 

to describe the flow field associated with vortex generators (VG) 

used to alleviate separation. 

Approach 

The approach is to modify and couple a three-dimensional 

multiblock Euler and a thin-layer Navier—Stokes analysis to 

model the flow over and downstream from VG devices. Downstream from the VG, a shock wave interacts with an adverse 

pressure gradient producing a separated region. The analysis will 

be validated with experimental data and then used in a parametric study to design optimized VG shapes by identifying the 

dominant length scales. 

Accomplishment Description 

A simplified analysis based on the solution of the Euler equations 

over a VG coupled with a Navier—Stokes applied downstream 

has been used to simulate the effect of shock-wave/boundarylayer interaction. A mesh was used to capture the variety of 

length scales inherent to the problem (e.g., boundary layer 

height (ö), VG height (h 8/3), the axial alleviation distance 

(=30-60i), and the lambda shock triple point (5)). The inviscid 

calculation used 175,000 grid points (4 blocks) and required 

40 megawords of storage on the Cray-2. The viscous calculation 

used 225,000 grid points (3 blocks) and required 64 megawords 

of storage. The first figure shows a three-dimensional display of 

the induced vorticity, the surface streamlines, and the axial 

velocity field. The analysis results indicate a surface flow "owl'seye" pattern similar to those observed experimentally, and a 

reduction of the reverse flow region at the point of separation 

and in the thickness of the reverse-flow bubble. These calculations require approximately 60 Cray Y-MP hours. Parametric 

inviscid flow analyses were conducted to investigate the effect of 

varying VG geometrical length scales and shapes. Performance 

assessment entails evaluating the shed streamwise vorticity, the 

net circulation, and the energization of the axial component of 

velocity. The second figure illustrates the calculated induced 

vortical flow fields for several configurations studied. 

Significance 

Normal shock/boundary-layer interaction separation occurs in 

many aerodynamic devices, such as transonic inlets, diffusers, 

and airfoils, resulting in detrimental effects on drag and pressure 

recovery. Alleviating or reducing the effects of these phenomena 

will have a significant impact on the efficiency of aerodynamic 

devices. 

Future Plans 

Future work on this project has not yet been determined.

Transonic-shock/boundary-layer interaction alleviation due to 

vortex generation. 

TRIANGULAR RAMP WHEELER SINGLET 

WENNERSTROM DEVICE TRIANGULAR PLOW 

Parametric vortex generation analyses.

143
Rotorcraft Drag Prediction 

J. C. Narramore, Principal Investigator 

Co-Investigators: A. G. Brand 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 

We utilized three-dimensional Navier—Stokes codes to perform 

both the steady and unsteady analyses of rotor blades and 

fuselages in flight and correlated 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 and fuselages, with special emphasis 

on the calculation of drag levels, has been performed. Work has 

been focused in three areas: correlation of flows about fuselages, 

correlation of flows about complex rotor tips, and modeling of 

inlet flows. Navier—Stokes solutions for the flow about a helicopter fuselage at Mach 0.23 were compared to wind tunnel test 

data with grid topology, angle of attack, and yaw angle being 

varied. The best correlation was obtained from the C-O grid 

topology (see accompanying figure). The difference between the 

measured drag and the computed drag was only 4% for a low 

angle-of-attack case, 7% for a high angle-of-attack case, and 

13% for a high-yaw case when this grid topology was used. 

Significance 

Computational fluid dynamics methods offer high potential for 

improved rotorcraft design and performance capabilities. This 

new technology should reduce vehicle costs, improve performance, and reduce development cycle times. 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. The first Navier—Stokes solutions that 

correlate well with measured drag levels on a typical helicopter 

fuselage were obtained. 

Future Plans 

We will (1) modify the boundary conditions on a Navier—Stokes 

code to allow the modeling of an actuator disk; (2) correlate 

wind tunnel data for a ducted-tail-rotor model to the results from 

this Navier—Stokes method; (3) use a Navier—Stokes code to 

investigate the flow about an anhedral rotor-tip shape; and 

(4) develop Navier—Stokes solutions for a more complex 

helicopter rotor fuselage using C-O and H-H grid topologies 

and compare the results with test data.

Publications 

1. Scott, M. T. and Narramore, J. C. "Navier—Stokes Correlation 

of a Swept Helicopter Rotor Tip at High Alpha." Presented at 

the AIAA 22nd Fluid Dynamics Conference, Honolulu, HI, 

June 1991. 

2. Narramore, J. C. and Brand, A. G. "Navier—Stokes Correlations to Fuselage Wind Tunnel Data." Presented at the 

48th Annual Forum of the American Helicopter Society, 

Washington, DC, June 1992. 

i 1XV. 

. NEWS Wind Tunnel Results 

. Navier Stokes Resul 

2 

Cr

0 :1 6 

rag/t(t1

Navier—Stokes results correlate with wind tunnel drag 

measurements. 

144
Controlling Combustion in Free-Shear Layers 

David Nixon, Principal Investigator 

Co-Investigator: Laurence Keef 

Nielsen Engineering and Research, Inc. 

Research Objective 

To investigate the generic effects of heat release on free-shear 

layers, particularly theoretical predictions that asymmetrical heat 

release can promote mixing and layer spreading. 

Approach 

We will test three-dimensional direct numerical simulations of 

low-Reynolds-number, spatially developing, compressible shear 

layers. First, we will test arbitrarily specified heat release, then 

we will test a simple three-component chemical reaction. 

Accomplishment Description 

After the addition of an independently specifiable heat-release 

term to the energy equation in our Navier—Stokes simulations, 

the effect of steady, but spatially varying, heating on a twodimensional layer was investigated. Because of the particular 

spatial form chosen, the total heat release per unit of streamwise 

distance was constant, thus it was consistent with linear growth 

of a reacting shear layer. Previous experiments and simulations 

have left the impression that heating always reduces shear-layer 

growth. Our theoretical work indicates that an asymmetric heatrelease profile across the layer can actually increase layer 

growth. So far, the numerical work has confounded both 

expectations. Although mixing rate increases due to heating 

asymmetry have not been found, neither have noticeable 

decreases in mixing been detected at the heating rates employed. 

The figure displays the mixing rates of heated and unheated 

mixing layers at M = 0.4. Despite a greater than 30% rise in 

centerline temperature in the heated case, the spreading rate, 

measured by the vorticity thickness (), is essentially unaffected. 

Such results have prompted a reexamination of the theory. The 

calculations required to obtain developed three-dimensional 

shear layers and statistics on local Mach numbers took roughly 

7 Cray Y-MP hours and 16 megawords of memory.

Significance 

This work tests the basic assumptions underlying a novel theory 

linking heat-release asymmetry to increased mixing. 

Future Plans 

We will exercise the simple chemistry model and perform 

experiments where the heat release profile across a shear layer 

is tailored to increase mixing. Also, we will determine if the 

resulting flow configurations are practical to implement in a 

high-speed combustor setting. 

12

J

2-11) layer

.D layer, heatedf\,, 

0 50 100 150 200 250

X/8(0 

Mixing rates of heated and unheated mixing layers. 

S 

CIO

Mc 0.4

145
Thermomechanical 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 thermomechanical buckling and post-buckling 

responses of multilayered composite panels with cutouts. 

Approach 

The computational model of the composite panels used in this 

study is based on the geometrically nonlinear shallow-shell 

theory, with the effects of both transverse shear deformations and 

temperature dependence of the material properties included. 

Analytic sensitivity coefficients are evaluated, measuring the 

sensitivity of the post-buckling response to variations in material 

and lamination parameters. The panels are subjected to combined applied-edge displacement and temperature increase. An 

efficient reduced-base computational procedure is used for 

determining the stability boundary, generating the post-buckling 

response and evaluating the sensitivity derivatives. The procedure allows the number of degrees of freedom used in the initial 

discretization to be significantly reduced. 

Accomplishment Description 

Extensive numerical studies were performed to understand the 

effects of temperature dependence of material properties,

lamination parameters, and the size of the cutout on the stability 

boundary and post-buckling response of flat unstiffened panels. 

The first figure shows the affect of hole size on the stability 

boundary for a 16-layer quasi-isotropic panel with fiber orientation. The second figure compares the normalized total-strain 

energies in the post-buckling range for two panels, one with and 

one without a cutout, when subjected to temperature increase. 

The third figure shows the sensitivity coefficients with respect to 

material parameters. 

Significance 

The structures of future high-speed and high-performance aircraft 

will be subjected to higher temperatures for longer periods. This 

study will help in the lamination selection and identify the 

material systems required. 

Future Plans 

Detailed studies of the effect of temperature dependence of 

material properties will be conducted. Stiffened panels with 

cutouts will also be considered. The numerical simulations will 

be closely coupled with laboratory experiments to identify the 

failure mechanisms and develop failure criteria for the panels. 

p 

/ 

q 

240 - 

180

L/ UL 

Fig. 2 - Posthuckling response

TIT

.9 

.6 

.5 

.4 

.3 

.1 

3

dIL 

T/Tcrm, 

240 

o,TL'

120

aL V ET 

.LJ L 

E,h' 

Fig. 3 - Sensitivity coefficients 

The effects of temperature dependence of material properties, lamination parameters, and the size 

of the cutout on the stability boundary and post-buckling response of flat unstiffeneci panels. 

146
Space Transportation Vehicle Aerothermodynamics 

Grant Palmer, Principal Investigator 

Co-Investigators: Susan Tokarcik, Ethiraj Venkatapathy, and William J. Feiereisen 

NASA Ames Research Center 

Research Objective 

To determine if the effects of self-generated and/or externally 

applied magnetic fields have a significant impact on the flow 

behind the bow shock wave of an aerobrake returning at 

hypersonic velocities from a lunar or Mars mission and if these 

effects are of consequence to vehicle design. 

Approach 

The axisymmetric thermochemical non-equilibrium Navier—Stokes 

equations with finite-rate chemical reactions are loosely 

coupled with Maxwell's equations and are solved using 

upwind-differencing flow codes. 

Accomplishment Description 

Flow was computed over a Mars-return aerobrake at 13.2 km/sec 

corresponding to a flight trajectory point of 64.8 km altitude. 

Peak ionization levels along the stagnation streamline exceeded 

25%. This level of ionization produced a self-generated magnetic field of 1.86 Gauss. This level of magnetic induction had

no discernible effect on the other flow quantities. It was shown 

that an externally applied magnetic field of 0.1 Tesla, a moderately strong magnetic field, increased the computed shockstandoff distance by a factor of two. The effect of an externally 

applied magnetic field was found to vary as a function of the 

field strength squared. A typical calculation took 6 Cray-2 hours 

and required 5 megawords of memory. The accompanying figure 

shows self-generated magnetic induction contours over the 

aerobrake. 

Significance 

Including the magnetic terms with the equation doubles the CPU 

cost of a flow computation. This study shows that the effects of 

undisturbed self-generated magnetic induction are negligible for 

estimated trajectories of Mars-return aerobrakes. 

Future Plans 

There is no continuation of this study because the research 

objective has been definitively answered. 

Magnetic Induction (gauss) Computational Domain ,. 

CONTOUR LEVELS 

0. 000011] 

0.06000 

CL 1201%) 

0.180110 

0.280110 

0.38000 

0,36000 

0.W00 Piers Return Herobrek 11. 48000 

a. 601001) 

a. 66000 

0.72000 

0.78000 

0.84000 

0.90000 

0. 960M I) 13.2 kri/s 

Hit. 64.8 kin 

2LJ[1U 

1.38000 

1.44000 

1. 50000 

56000 

• 62000 

.68000 

74001) 

1.30000 

1.86000

Self-generated magnetic field around a Mars-return aerobrake.

147 
0.062a 0.0789 0.172a 0.1888 

-5800 

(1JnJIn.)

Thermo-Viscoplastic Analysis of an Engine-Cowl Leading Edge 

Ajay K. Pandey, Principal Investigator 

Lockheed Engineering and Sciences Company/NASA Langley Research Center 

Research Objective 

To predict the thermo-viscoplastic response of an engine-cowl 

leading edge subjected to oscillating shock—shock interaction. 

Approach 

Prediction of thermo-viscoplastic response of an engine-cowl 

leading edge involves predicting the temperature solution and 

then performing structural analysis using a unified viscoplastic 

constitutive relation. First, aerodynamic heating and pressure 

due to oscillating shock—shock interaction is determined. These 

loads vary in both space and time. For the specified aerodynamic heating, radiation, and other boundary conditions, 

thermal analysis is performed by solving the energy equation. 

Resulting transient temperature response and specified pressure 

used to predict structural response by solving the structural 

equilibrium equation. The unified viscoplastic model is used in 

the structural analysis and relates plastic strain rates in terms 

of stresses and internal state variables, combining the timeindependent and time-dependent responses into a single 

inelastic component. The internal-state variable represents 

hardness of the material and is given in time—rate form in terms 

of temperature dependent material properties and plastic work 

rate. Since the unified viscoplastic constitutive relations are 

highly nonlinear, very small time steps are required to march the 

solution in time. An adapted unstructured finite-element mesh is 

used to improve the solution accuracy and computational 

efficiency. 

Accomplishment Description 

A finite-element thermo-viscoplastic code was used to predict 

thermo-viscoplastic response. An unstructured finite-element 

mesh was adapted from an initial solution using temperature and 

circumferential stress as the key parameters. A peak temperature of 

about 3,000 °R was predicted on the outer surface of the leading 

edge during the inward and outward movement of the shock. 

Elastic analysis over-predicted the circumferential stresses on 

the leading edge and was about 50% higher than the stresses 

predicted by thermo-viscoplastic analysis in the region of peak 

aerodynamic heating and pressure. Analysis also predicted 

permanent plastic deformation on the leading edge. The 

accompanying figure shows the increase in plastic region as the 

shock moves inward and outward. The effective plastic-strain 

contour shows that the peak effective plastic strain keeps 

increasing with time.

Significance 

The leading edge experiences intense aerodynamic heating 

resulting in high temperatures where viscoplastic effects are 

significant. Stresses predicted by thermo-viscoplastic analysis are 

significantly lower than the elastic analysis in the peak heating 

region. Thermo-viscoplastic analysis identifies the regions 

undergoing plastic deformation. Use of an adaptive finiteelement method significantly reduces the problem size. 

Future Plans 

We will study the transient thermal and viscoplastic response of 

three-dimensional leading-edge structures using adaptive finiteelement methods. 

Publication 

Pandey, A. K. "Thermo-Viscoplastic Analysis of Engine-Cowl 

Leading Edge Subjected to Oscillating Shock—Shock Interaction." 

Presented at the AIAA 33rd Structures, Structural Dynamics, and 

Materials Conference, Dallas, TX, April 1992. 

Effective plastic-strain contours. 

148
Afterbody Aerodynamics with Canted Pitch-Vectoring Twin Nozzles 

S. Paul Pao, Principal Investigator 

Co-Investigator: K. S. Abdol-Hamid 

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

Research Objective 

To obtain detailed surface and field-pressure distributions, 

identify regions of flow separation, and calculate the mass flow 

rate and forces of the jet propulsion system. The installation of 

twin rectangular thrust-vectoring nozzles for advanced aircraft 

propulsion has many advantages. However, it is both geometrically and aerodynamically complex. 

Approach 

A multiblock structured-grid Navier-Stokes code, PAB3D, is 

used for the numerical analysis. Algebraic and two-equation 

k-c turbulence modeling are available within PAB3D. A multiblock grid generator has been developed in-house for the canted 

twin-nozzle/afterbody configuration. Half-body grids are used 

for studying the symmetrical pitch-vectoring configurations. Fullbody grids will be necessary for cases where pitch and yaw 

forces are effected by the nonsymmetrical diverging-flap setting 

of the twin nozzle system. 

Accomplishment Description 

A 54-block surface-fitted grid was generated for the right half of 

the twin-nozzle/afterbody with a zero degree vector angle for the 

nozzle. The unusual blocking strategy provided a grid that is 

well matched in terms of orthogonality and spacing, and is a 

continuous boundary layer over the afterbody and the nozzle 

interior. One solution, for free-stream Mach number 1.2 and 

nozzle pressure ratio 4.0, was obtained during this project 

period. Surface-pressure coefficient contours on the twinnozzle/afterbody configuration are shown in the accompanying 

figure. Significant findings include a local-shock compression 

zone upstream from the nozzle assembly (top of figure), nonsymmetrical flow separation above and below the nozzle 

assembly, and jet-plume interaction with the flow behind the 

afterbody. This half-body grid contains 550,000 grid points and 

requires 11 megawords of memory. A converged solution takes 

approximately 5,000 iterations, requiring 30 Cray Y-MP hours.

Significance 

Numerical results from this investigation are useful in understanding the aerodynamics for complex configurations. The 

results also provide a quantitative basis for drag-reduction design 

and the prevention of afterbody flow separation. Although 

experimental data for force measurements are available, surface 

pressure data have not yet been obtained for this configuration. 

Future Plans 

We will study the effect of turbulence modeling and unstructuredgrid Euler calculations, possibly including hybrid structured,' 

unstructured computations for efficient representation of the 

boundary-layer, separated-flow, and jet-plume interactions. 

BOTTOM VIEW TOP VIEW 

Canted twin pitch-vectoring nozzles-a ftethody pressure-coefficient 

contours. Mach number = 1.20, nozzle pressure ratio = 4.03, 

contour level from blue to red -0.75-0.65 at 0.05 intervals.

149
Hydrodynamics of Self-Propelled Bodies 

V. C. Patel, Principal Investigator 

Co-Investigator: F. Stern 

University of Iowa 

Research Objective 

To extend and further develop the Iowa Institute of Hydraulic 

Research (IIHR) numerical method for solution of the Reynoldsaveraged Navier—Stokes (RANS) equations to simulate the flow 

around complete hydrodynamic vehicles with appendages and 

propellers. 

Approach 

The IIHR numerical method solves the complete RANS equations in numerically generated, body-fitted, nonorthogonal 

coordinates for unsteady three-dimensional turbulent flows. 

A two-equation turbulence model is used, either with wall 

functions or in conjunction with a one-equation model for the 

near-wall flow. Several basic numerical and physical problems, 

including grid generation for complex shapes and moving 

boundaries, time accuracy for unsteady flows, and turbulence 

model development are investigated. To generalize the method 

for prediction of unsteady flow around three-dimensional marine 

vehicles and propellers, a number of interrelated model problems have been formulated. Among these are: the flow around 

bodies at incidence and two-body junctions to study the 

topology of three-dimensional flow attachment and separation, 

and evaluation of turbulence models; the flow around a propeller to resolve the viscous flow over and around the blades; and 

the flow over bodies with appendages and/or propellers. 

Accomplishment Description 

The flow on a prolate spheroid at various incidence angles and 

the flow around a hemisphere placed on a flat wall have been 

computed to clarify the topology of separation. Studies have 

been completed both for the SR-7 turboprop for on-design and 

off-design conditions and for the interaction between a marine 

propeller and a representative ship hull, including comparisons 

with available experimental data. Zonal methods, advanced 

turbulence modeling for simplified geometries, and timeaccurate unsteady flow calculations for natural and forced 

unsteady flows using overlaid grids were also implemented. 

Average job-run Cray hours and memory requirements are: 

three-dimensional steady flow computation, 7 Cray-2 hours and 

13 megawords of memory; two-dimensional fixed-grid unsteady 

flow computations, 2 Cray-2 hours and 3 megawords of 

memory; the more complex overlaid-grid unsteady flow computation, 10 Cray-2 hours and 4 megawords of memory.

Significance 

This research is being carried out to advance the technology 

base for the design of marine vehicles and propellers. The 

problems addressed relate to critical gaps in our knowledge of 

complex three-dimensional unsteady turbulent-shear flows. 

Future Plans 

The numerical method is being revised to achieve greater 

accuracy and efficiency for unsteady flows. Steady and unsteady 

flow calculations for realistic marine propeller geometries are 

being developed. Unsteady flow calculations to study the 

interaction between natural and forced unsteady flows using 

overlaid grids will also be studied. 

Publication 

Patel, V. C. "Three-Dimensional Flow Separation." Fifth Asian 

Fluid Mechanics Congress. Taejon, Korea, Aug. 1992. 

SR-7 turboprop leading-edge and tip-vortex flow pattern. 

150
Hypersonic Scramjet Flow 

Principal Investigator: Darrell W. Pepper 

Co-Investigators: Frank P. Brueckner and Kevin L. Burton 

Advanced Projects Research, Inc. 

Research Objective 

To predict the internal flow field of a hypersonic scramjet engine 

in conjunction with development of the hypersonic engine for 

the National Aero-Space Plane (NASP). Numerical simulations 

of the inlet, combustor, and nozzle sections of the engine are 

necessary to predict engine efficiency and performance. 

Approach 

Both two- and three-dimensional Navier—Stokes codes based 

on unstructured grids and finite-element methods are being 

used to model supersonic and hypersonic internal flows in 

combustor—nozzle configurations. 

Accomplishment Description 

The goal is to develop a reliable multidimensional algorithm for 

simulating compressible inviscid and viscous flows. Continued 

refinement of the finite-element code, RUBY3D, was undertaken 

last year. Incorporation and testing of a finite-rate chemistry 

module within RUBY2D was completed. In addition, parallel 

versions of the code were developed and run on a MasPar 1216 

SIMD massively parallel computer (16,384 processing elements). 

The RUBY family of codes include two- and three-dimensional 

and axisymmetric capabilities for solving Euler and viscous 

laminar-flow regimes. All the codes are explicit, employ 

Petrov—Galerkin weighting for the advection terms, and utilize 

unstructured meshes. Modifications are employed in all the 

codes to enhance solution speed and reduce storage. The codes 

interface with PATRAN, a commercial grid generation package 

for preprocessing and grid generation. PLOT3D, a NASA Ames 

program for graphical post-processing, can be used to display 

two- and three-dimensional structured mesh results generated 

by the RUBY codes. A three-dimensional mesh of a round-torectangular nozzle and surface pressure contours for Mach 3 

inflow are shown in the accompanying figures. To generate the 

mesh, 55,000 elements were used. An average inviscid job run 

required 1 Cray Y-MP hour and 7 megawords of memory. 

Significance 

Successful simulation of the internal reacting flow field is 

important in the design and understanding of engine performance and airframe integration. Sufficient wind tunnel data can 

be taken in scramjet engine tests up to Mach 8. However, the 

NASP requirements extend from low subsonic to Mach 25 

conditions. Consequently, numerical simulation is required to 

design the engine and airframe beyond present experimental 

limits. 

Future Plans 

The reacting flow capability of RUBY21D will be extended to the 

three-dimensional algorithm with the addition of turbulence 

models planned for all programs. Optimization of the RUBY 

codes for increased parallel performance and incorporation of 

three-dimensional mesh adaptation will be continued.

Publications 

1. Burton, K. L. and Pepper, D. W. "An Adaptive Finite-Element 

Algorithm for Compressible Flow on Workstations." Presented 

at the ASME WAM, Atlanta, GA, Dec. 1991. 

2. Pepper, D. W.; Jacobsen, K. P.; and Rajan, M. "A Hybrid 

Finite-Element Method for Calculating Compressible Flow 

on a Massively Parallel Computer." Presented at the Fourth 

International Symposium in CFD, Davis, CA, Sept. 1991. 

Three-dimensional mesh of round-to-rectangular nozzle and 

surface pressures for Mach 3 inflow.

151
Wave-Induced Transports 

Leonhard Pfister, Principal Investigator 

NASA Ames Research Center 

Research Objective 

To develop a better understanding of the role of atmospheric 

waves in the momentum budget of the Earth's stratosphere, 

stratosphere—troposphere mass exchange, and stratospheric 

mixing. 

Approach 

We use a linear three-dimensional time-dependent approach to 

simulate stratospheric gravity waves observed during the tropical 

phase of the Stratosphere—Troposphere Exchange Project, a 1987 

ER-2 aircraft field experiment based in Australia. The fields are 

represented as two-dimensional 160 x 160 Fourier series at 

100 altitudes. Each Fourier component is independently integrated 

in time. The convective forcing, located at the tropopause, took 

the form of temporary "mountains" poking into the stratosphere. 

The temporal and horizontal scales of these mountains were 

based on infrared imagery from the GMS Japanese Geostationary 

weather satellite. By matching ER-2 aircraft observations with the 

simulation, we were able to place constraints on the forcing 

amplitudes of tropical convection. 

Accomplishment Description 

The simulations (an example is shown in the accompanying 

figure) showed that the mesoscale upward displacement of 

material surfaces by the convection is about 450-600 meters. 

They also produced an unexpected result. By noting the difference between the observed and simulated decay phases of the 

wave, we deduced that wave breaking and subsequent nonlinear 

interaction was occurring at the aircraft flight altitude. This 

implies that the waves excited by convection can directly drive 

the flow in the lowest 3-4 km of the stratosphere. A typical run 

took about 20 Cray-2 minutes used 40 megawords of internal 

memory. 

Significance 

The calculated magnitudes of connectively induced displacements have two major implications. First, mesoscale gravity 

waves excited by convection have sufficient amplitude to make 

a significant contribution to the upper stratosphere semiannual 

wind oscillation and the lower stratosphere quasi-biennial 

oscillation. The latter strongly contributes to interannual 

variations in the ozone distribution. Second, wave breaking in 

the lowest 3-4 km of the stratosphere may have a contribution 

comparable to that of radiative heating in driving stratospheric 

upward motion.

Future Plans 

We will use a three-dimensional nonlinear model to simulate the 

interaction of convectively excited gravity waves with the mean 

lower-stratospheric circulation. This is to evaluate the contribution of these waves to tropical stratosphere—troposphere exchange. 

Publication 

Pfister, L.; Scott, S.; Loewenstein, M.; Bowen, S.; and Legg, M. 

"Mesoscale Disturbances in the Tropical Stratosphere Excited by 

Convection: Observations and Effects on the Stratospheric 

Momentum Budget." Accepted for publication in the Journal of 

the Atmospheric Sciences, 1992. 

ci) 

—D 

a 

—J

—16

136 138 140

Longitude 

Temperature field of a simulated transient linear-gravity 

wave 1.5 km above the tropopause (white = 197.5 K and 

black = 192.5 K). The convective forcing is contoured in white, 

with the highest and coldest clouds outlined by the thickest 

lines. 

152
Rocket-Base Flow-Field Simulations 

Thanh T. Phan, Principal Investigator 

Co-Investigator: Richard J. Magnus 

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-dirnenioniI Revnoft!s-averaged Navier--Stokes tIo\ 

solver, FALCON, is used to simulate the flow field in the base 

region of the launch vehicle. The code employs a finite-volume, 

cell-centered, upwind-differencing scheme that includes a 

k-kl turbulence model. A multiblock grid structure is used to 

conform to the complex geometry of the base region. 

Accomplishment Description 

A perfect-gas turbulent solution was obtained for the base region 

of an Atlas II launch vehicle (no solid-rocket boosters) at Mach 

1.8. Seven grid blocks totaling over one million points were 

used. Local time stepping at each cell was applied to accelerate 

convergence to steady state. The solution required approximately 14 megawords of memory and 70 Cray Y-MP hours. 

Significance 

Currently, the environment in the base region of a new launch 

system is predicted with very large uncertainties. These uncertainties lead to designs of thermal protection systems that 

include many over-conservatisms that cannot be used with a 

high level of confidence. The calculations in this research will 

provide valuable information to help design optimal protection 

systems that can be used with more confidence. 

Future Plans 

Real-gas effects and flow-field grid adaptation will be implemented to model plume interactions more accurately. Forebody 

effects will eventually be simulated, including the solid-rocket 

boosters and their exhaust products.

Mach number contours. 

Perfect-gas turbulent solution for an Atlas 1/ base flow at Mach 1.8. 

153
WoM -7

.4 

/ 

• 1! 1? 

Three-Dimensional Liquid-Sloshing Flows 

Richard H. Pletcher, Principal Investigator 

Co-Investigators: Franklyn J. Kelecy and Babu Sethuraman 

Iowa State University 

Research Objective 

To develop numerical simulation methods for predicting the 

complex, transient, three-dimensional motion of a sloshing 

liquid in a partially-filled container undergoing arbitrary motion. 

Approach 

The motion of a sloshing liquid is governed by the threedimensional incompressible Navier—Stokes equations. The 

presence of the free surface poses special problems as its 

position is not known a priori and has to be evaluated as part 

of the solution. Two approaches, surface fitting and surface 

capturing, are being evaluated. The surface-fitting code 

(SLOSH3D) places the free surface at the boundary of the 

computational domain. The free-surface position is estimated 

using a kinematic boundary condition. The motion of the freesurface boundary necessitates the use of a moving grid, which is 

updated once the new free-surface position is known. The flowfield solution inside the liquid is obtained by using the artificial 

compressibility approach in conjunction with a coupled strongly 

implicit procedure (CSIP) to solve the resulting system of 

equations. The surface-capturing approach also uses artificial 

compressibility, but solves conservation equations in both liquid 

and gas phases on a stationary grid, thus "capturing" the free 

surface as a density discontinuity. 

Accomplishment Description 

The CSIP algorithm was vectorized along surfaces of constant 

index sums and optimized for use on the Cray Y-MP. A typical 

t=1.498sec t=1.996sec

transient calculation (using 5,000 grid points and 3,000 time 

steps) required about 5 megawords of memory and 1 Cray Y-MP 

hour. The code has been used to compute transient solutions for 

symmetric spin-ups and asymmetric orbital motions of half-filled 

spherical containers. Selected results for the case of a simple 

orbital spin-up to a steady-state angular speed of 60 rpm are 

presented in the accompanying figures. The free-surface position 

is depicted at different times during the transient, illustrating the 

complex three-dimensional surface behavior as it approaches 

steady state. Comparisons with experimental data showed good 

agreement. 

Significance 

Liquid sloshing arises in many important areas, including the 

National Aero-Space Plane and various space and space station 

projects. Forces generated by liquid motion can result in damage 

or instability in vehicles. Numerical methods for sloshing flows 

provide an important tool for analyzing the details of the fluid 

motion and its effects on the container structure. 

Future Plans 

We will enhance the simulation capability to higher Reynolds 

numbers. Work is progressing on integrating the sloshing 

calculations with a finite-element structural model so that the 

fluid—structure system interaction can be investigated in detail. 

t=2.494sec t=3.489sec 

Selected free-surface (top) and velocity-vector (bottom) plots for the orbital spin-up of a spherical tank half filled with glycerine. 

Steady-state angular velocity = 60 rpm. 

154 
Martian Atmosphere General Circulation 

James B. Pollack, Principal Investigator 

Co-Investigator: Robert Haberle 

NASA Ames Research Center 

Research Objective 

To simulate the present and past climate regimes on Mars. A 

general circulation model (GCM) predicts the time-evolving 

three-dimensional wind and temperature fields and the surface 

pressure, temperature, and carbon dioxide ice abundance. 

Approach 

The Mars GCM is based on the primitive equations of meteorology that are used for terrestrial weather prediction. Uniquely 

Martian physics, such as the condensation and sublimation of 

carbon dioxide in the polar region, are incorporated into the 

Mars GCM. It has been interfaced with an aerosol physics model 

to simulate the transport of dust by winds and the impact of 

sunlight absorption by dust on temperatures and winds. 

Accomplishment Description 

The life cycle of global dust storms on Mars has been simulated 

with the coupled GCMJaerosol physics models for several 

plausible spatial distributions of surface sources of dust. We 

found that the strong feedback between dust heating and winds 

plays an essential role in the spreading of the dust from its 

southern hemisphere to the opposite hemisphere. Eddies, 

especially transient-baroclinic eddies, enable the dust to travel 

from northern mid-latitudes to the northern polar region. The 

results of these simulations are mostly in accord with relevant 

observations.

Significance 

The seasonal cycle of dust represents an important component 

of the present climate on Mars since it strongly influences 

atmospheric temperatures and winds. Understanding the 

development of global dust storms is an important step toward 

gaining insight into the factors that determine the amount of dust 

in the atmosphere and its seasonal and interannual variations. 

Future Plans 

A parameterization will be developed for the amount of dust 

lifted into the atmosphere by surface winds of varying strength 

and it will be used in fully interactive simulations of the seasonal 

dust cycle. The Mars GCM will be further developed for data 

analyses returned by the Mars Observer spacecraft. 

Publications 

1. Pollack, J. B.; Haberle, R. M.; Schaeffer, J.; and Lee, H. 

"Simulations of the General Circulation of the Martian 

Atmosphere: Part I. Polar Processes." J. Geophys. Res. 95 

(1990): 1447-1473. 

2. Pollack, J. B.; Haberle, R. M.; Murphy, J. R.; Schaeffer, J.; 

and Lee, H. "Simulations of the General Circulation of the 

Martian Atmosphere: Part II. Seasonal Pressure Variations." 

Accepted for publication in ICR Planets. 

SOL1 SOL3 SOL5 

to40 ^^^^ 10 

-ISO -120 -60 0 80 20 50 

LONGITUDE

-ISO -120 -80 0 60 

LONGITUDE

120 ISO -ISO -120 -60 0 60 20 180 

LONGITUDE 

SQL 10 SOL 20 SQL 50 

0 90 90 

60 60 

6.1 3C

0 0

UJ 

0

30 

D 

-

-30

8-

-30 

-60 -60 

-90 . _, ....... 

-2 -, - -- -110 -120 -60 0 60 120 ISO -ISO -120 -60 0 60 11120 150 

LONGITUDE LONGITUDE LONGITUDE

Instantaneous maps of vertically integrated visible dust opacity (r) at several times from the start of a simulated Martian global 

dust storm. The prescribed invariant surface dust source persists for 10 sols (sols = number of Mars days). Opacity values 

increase from magenta (r <0.5, including 0), violet (r = 2), blue (r = 4), green (r = 6), yellow ( = 8), to red (t^t 10).

155 
MID STR IUT DENSITY 

MID-SPAN DENS Y 

H: AT STRUT EXIT H AT COMBUSTOR EXIT 

Three-Dimensional Scramjet Combustors 

Greg D. Power, Principal Investigator 

Co-Investigator: Thomas J. Barber 

United Technologies Research Center 

Research Objective 

To identify mixing-enhancement concepts and simulate the 

complex three-dimensional interaction effects found in a scramjet 

combustor using a state of the art Navier—Stokes analysis. 

Approach 

The approach was to develop necessary turbulence and kinetics 

models and then implement them into UTSPARK, a version of 

the SPARK Navier—Stokes code developed at NASA Langley. The 

analysis was used to examine the nonreacting and reacting flow 

from an innovative scramjet-combustor fuel-injection concept. 

Accomplishment Description 

As a result of deficiencies identified through numerous validation 

studies, we developed and implemented into the UTSPARK 

Navier—Stokes code the two-equation Jones and Launder (k-c) 

turbulence model, including the recent compressibility corrections introduced by Sarkar. The analysis was applied to the 

mixing rates and combustion efficiency of a combustor fuelinjection scheme utilizing laterally spaced injector struts. The 

analysis, performed in support of a series of subscale combustor 

experiments, entailed simulation of the facility plenum nozzle 

and the test section. Fuel was injected from the sides of the strut

and at an angle to the free stream from the strut tip. The accompanying figure illustrates the strut and the fuel concentration. 

The test-section calculations, which utilized approximately 

200,000 grid points, required approximately 1 25 megawords of 

memory and about 75 Cray Y-MP hours. The fuel concentration 

indicates that the injected fuel is near the lower surface of the 

facility. Also, while the overall equivalence ratio is sufficient to 

burn all the fuel, the local equivalence ratio is either near zero 

or much greater than one, resulting in reduced combustion 

efficiency. 

Significance 

The efficiency of a scramjet combustor is limiting technology in 

successfully achieving a flyable hypersonic vehicle such as the 

National Aero-Space Plane. Critical factors influencing the 

combustor design are the combustor fluid dynamics, rate of heat 

release, and combustion efficiency. Numerically validated codes 

represent a viable design technique to augment conventional 

experimental design approaches. 

Future Plans 

Future work on this project has not yet been determined. 

Three-dimensional simulation of a combustor scramjet. 

156
0 

.6 

C) 

CA 

AC 

CA 

CA .4 - 

.2

( 

0-

-1.5

M = 17.9 

T = 205.3 K 

p = 4.3E-05 k9/rn3 

(c) -1.0 -.5 2r 

xlr 

Hypersonic-Flow Computations using Adaptive Unstructured Meshes 

Ramadas K. Prabhu, Principal Investigator 

Lockheed Engineering and Sciences Company/NASA Langley Research Center 

Research Objective 

To determine the effects of thermal and chemical non-equilibrium 

on aerothermal loads experienced by bodies in hypersonic flow. 

Approach 

Air is modeled as a reacting mixture of five chemical species 

having two temperatures—one for the translational-rotational 

mode, and the other for the vibrational-electronic excitation 

mode. The governing equations are the mass conservation 

equations for the five chemical species, the momentum equations, and the two energy equations (one each for total energy 

M = 17.9, 1,,, = 205.3 K, r = 1mm

02 

N2 

----0 

- -- NO 

-- N 

(a)

M = 17.9, T0,, = 205.3 K, r = im

and the vibrational-electronic excitation energy). These are 

solved using a finite-volume approach and a cell-centered 

upwind scheme. Unstructured meshes are employed to 

discretize the computational domain. An existing computer 

code, LARCNESS, was modified to include these changes from 

ideal gas to non-equilibrium flows. 

Accomplishment Description 

Results were obtained on several problems involving hypersonic 

flow past blunt bodies. The accompanying figure shows sample 

results obtained for an inviscid Mach 17.9 flow past a blunt 

body assuming a single temperature model. Results were 

obtained for two cases, namely r = 1 mm and r = 1 m. Mach 

number, temperature distribution, and the mass fraction distributions on the line of symmetry are shown in the figure. These 

results compared very well with the results of Desideri et al. 

Significance 

Air undergoes dramatic compositional changes at the high 

temperatures encountered in hypersonic flight. Convectional 

ideal-gas or equilibrium-gas assumptions are not valid under 

these conditions. Multiple species are needed to model the 

chemical composition and multiple temperatures are required to 

characterize the energy of air undergoing fast change. This code 

demonstrates the need for such analysis. 

Future Plans 

The computer code is being extended to include the transport 

terms of viscosity, thermal conductivity, and mass diffusion. The 

generation of unstructured meshes is also in progress. Several 

examples of practical interest will be solved, including blunt 

bodies in hypersonic flow with shock-shock interference 

configurations. 

1.0 

.8 

.6 

U 

AC 

AC 

CA .4 

.2 

0 

1.0 

.8 

Results obtained for an inviscid Mach 17.9 flow past a blunt body. (a) Mass fraction distribution; Mach = 17.9, T,,, = 205.3 K, r = 1 mm. 

(b) Mass fraction distribution; Mach = 17.9, T,,0 = 205.3 K, r = I m. (c) Problem statement.

157
Laminar Breakdown of High-Speed Boundary-Layer Flow 

C. D. Pruett, Principal Investigator 

Co-Investigator: T. A. Zang 

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

Research Objective 

To computationally examine the physics governing the transition 

of high-speed compressible boundary-layer flows from ordered 

(laminar) to chaotic (turbulent) states. 

Approach 

A variety of computational tools have been applied to the 

problem, including direct numerical simulation (DNS), "adulterated" numerical simulation (ANS), large-eddy simulation, and 

secondary instability theory (SIT). 

Accomplishment Description 

Long-duration temporal DNS computations from initial states 

derived from SIT reveal that the (0,2) Fourier harmonic, not 

present in the initial condition, eventually dominates all other 

harmonics in energy content—a result not anticipated by the 

investigators. Subsequent review of recent scientific literature 

revealed that other researchers also observed and reported the 

dominance of modes of purely spanwise wave number that are 

associated with streamwise vorticity. In DNS, the energy content 

of selected harmonics can be monitored precisely and, if 

desired, controlled. In this spirit, the ANS was conducted with 

the (0,2) mode artificially suppressed. The resulting flow fields of 

the DNS and ANS computations are compared in the accompanying figures for a Mach 4.5 boundary-layer flow on a hollow 

cylinder. In the DNS computation, the highly three-dimensional 

flow shown at dimensionless time (T) 45 soon experiences 

laminar breakdown commensurate with saturation of the (0,2) 

harmonic at large amplitude at about T = 50. In contrast, 

suppression of the (0,2) mode dramatically reduces the threedimensionality of the flow field in the ANS, completely inhibiting laminar breakdown. Because relatively coarse resolution 

sufficed in the ANS in the absence of laminar breakdown, 

roughly 120 Cray-2 hours and 5 megawords of memory were 

used—substantially less than the original DNS. 

Significance 

The results of the ANS are potentially profound and suggest that, 

rather than focusing on the roles of primary and secondary 

instability mechanisms as in the past, transition research should 

perhaps focus on the role played by streamwise vorticity. 

Future Plans 

The temporal DNS that forms the core of this work is presently 

being recast in the context of the more physically correct (but 

computationally intensive) spatial stability theory, in which 

disturbances evolve in space rather than in time.

Publications 

1. Pruett, C. D. and Zang, T. A. "Direct Numerical Simulation of 

Laminar Breakdown in High-Speed, Axisymmetric Boundary 

Layers." To be published in Theoretical and Computational 

Fluid Dynamics, 1992. 

2. Ng, L. L. and Zang, T. A. "Secondary Instability Mechanisms 

in Compressible, Axisymmetric Boundary Layers." AIAA 

Paper 92-0743, 1992. 

3. Kral, L. D. and Zang, T. A. "Large-Eddy Simulation of 

Supersonic, Wall-Bounded Turbulent Flows." In ICASE/LaRC 

Workshop on Transition and Turbulence, ed. M. Y. Hussaini. 

New York: Springer, 1992. 

Surfaces of constant streamwise velocity at T = 45; direct 

numerical simulation. 

Surfaces of constant streamwise velocity at T = 45; "adulterated" 

numerical simulation. 

158
Shock Impingement in Hypersonic Engine Inlets 

R. Ramakrishnan, Principal Investigator 

NASA Langley Research Center 

Research Objective 

To study the effect of shock impingement on aerothermal loads 

in an axial corner simulating a hypersonic engine inlet using a 

three-dimensional dynamic grid-adaptation procedure. 

Approach 

A three-dimensional Navier—Stokes code, SCRAMAD, is utilized 

to predict the complex flow features associated with shock 

impingement on an asymmetric axial corner. The code includes 

a three-dimensional dynamic adaptation procedure that is based 

on the concept of spring-mass systems to adapt the finitedifference grid to regions of strong gradients. The adaptation 

procedure uses a linear combination of spatial derivatives of 

select flow variables to relocate the grid points. Laminar and 

turbulent Mach 6 flow in the corner are analyzed and compared 

to experimental results. 

Accomplishment Description 

The three-dimensional flow features associated with shock 

impingement on a hypersonic engine inlet has been investigated 

for both laminar and turbulent flows. The overall flow features 

and heat transfer rates predicted by the numerical analysis are in 

good agreement with experimental observations. The accompanying figure shows the interactions of the oblique shock from the 

forebody with the cowl surface. 

Significance 

The three-dimensional grid adaptation procedure is capable of 

handling highly stretched grids that are prevalent in hypersonic 

viscous flows and the procedure does not encounter problems 

due to grid crossover. The coupling of the grid adaptation 

procedure to the Navier—Stokes solver makes it possible to 

analyze three-dimensional hypersonic flow problems without 

the need for a prohibitive number of grid points. The grid 

adaptation procedure used ensures adaptation to multiple shocks 

and ensures good definition of the thin hypersonic boundary 

layer.

Future Plans 

Efforts are under way to increase the efficiency of the adaptation 

procedure, investigate grid adaptation measures, and use the 

computational procedure to analyze other three-dimensional 

problems involving complex shock/shock and shock/boundarylayer interactions. 

Publication 

Ramakrishnan, R. "Numerical Simulation of Aerothermal Loads 

in Hypersonic Engine Inlets due to Shock Impingement." AIAA 

Paper 92-2605, AIAA Applied Aerodynamics Conference, 

Palo Alto, CA, 1992. 

Interactions of the oblique shock from the forebody with the 

cowl surface.

159
71!iw I

Complex Three-Dimensional Flows in the Advanced Solid-Rocket Motor 

Edward J. Reske, Principal Investigator 

Co-Investigators: Dana F. Billings and Joni W. Cornelison 

NASA Marshall Space Flight Center 

Research Objective 

To characterize the internal flow environment in the advanced 

solid-rocket motor (ASRM). Three-dimensional computational 

fluid dynamics (CFD) models are tools in addressing key design 

issues. 

Approach 

A three-dimensional Navier—Stokes flow solver has been used to 

analyze complex flows in the aft end segment and nozzle of the 

ASRM. Input on physical reference quantities and boundary 

conditions were obtained from the JAN NAF Solid Performance 

Prediction ballistics program. 

Accomplishment Description 

Preliminary two-dimensional axisymmetric analyses were 

performed with the Navier—Stokes codes CMINT and FDNS3D. 

An immediate concern has been the prediction of hinge torque 

moments and loads produced by the internal pressure field 

acting on a gimballed nozzle; such a problem is inherently 

three-dimensional. The FDNS3D code was chosen for 

the three-dimensional analyses. Three-dimensional 

220 x 64 x 26 grids were formed for various burn times and 

nozzle gimbal angles by rotating a two-dimensional, axisymmetric template grid about its centerline in 26 planes, and 

then gimballing the nozzle about its pivot point. We considered

cases for the ASRM at burn times of 19 and 115 seconds, each 

with gimbal angles of 4 and 8 degrees. Similarly, a comparative 

study was made for the redesigned solid rocket motor (RSRM) 

at burn times of 19 and 114 seconds, each at a gimbal angle 

of 4 degrees. Each three-dimensional solution required 

20 Cray Y-MP hours and 16 megawords of memory. 

Significance 

These calculations will provide a useful tool in the design of the 

ASRM. The centers of pressure for the ASRM and RSRM nozzles 

are far upstream from the nominal pivot points, thus creating 

non-restoring hinge torque moments in both motors, and the 

loads and moments acting on the ASRM nozzle are well below 

those of the RSRM. This could save millions of dollars in the 

ASRM project because existing RSRM actuators could be used 

on the ASRM. The results show that the ASRM nozzle performs 

well in straightening the flow and aligning the thrust vector with 

the nozzle axis of symmetry. 

Future Plans 

CFD analyses addressing grid sensitivity, the effect of inhibitor 

failures, vortex shedding on the the star grain, code validation by 

direct comparison with subscale cold-flow test data, and twophase flow will be performed. 

ERN

Grid, wall-pressure contours, particle traces, and Mach-number contours for the ASRM 

at the 115 second burn time with an 8 degree gimbal angle. 

160
Leading-Edge Heat Transfer in a Flow with Spanwise Variations 

David L. Rigby, Principal Investigator 

Sverdrup Technology, Inc. 

Research Objective 

To investigate the interaction of free-stream vorticity with the 

stagnation region of a leading edge. The effect on leading-edge 

heat transfer is of particular interest. Results of this research 

should improve the understanding of the mechanism by which 

free-stream turbulence causes the heat transfer to increase. 

Approach 

The three-dimensional Navier—Stokes equations are solved using 

the PARC3D code. A periodic spanwise variation in velocity is 

imposed upstream from the body. On the body surface, either 

uniform wall heat flux or uniform wall temperature is imposed. 

Accomplishment Description 

Using the PARC3D code, flow over a blunt flat plate with a 

3:1 elliptical leading edge has been studied. A steady spanwise 

velocity profile was imposed upstream from the leading edge. 

As expected, for equivalent conditions, the circular leading edge 

has a higher heat-transfer coefficient than the ellipse. As a ratio 

to the corresponding result with no spanwise variation, the circle 

and ellipse show similar increases in most cases. Increases in the 

heat-transfer coefficient on the order of 20% were seen with 

upstream spanwise velocity variations as low as 5%. On

average, each case required approximately 10 Cray-2 hours 

and 10 megawords of memory. 

Significance 

Accurate prediction of surface heat transfer is essential in the 

design of efficient and reliable turbomachines. The heat transfer 

in the stagnation region is sensitive to low levels of turbulence 

and this study indicates the stagnation-region heat transfer is also 

sensitive to small spanwise variations in the mean free-stream 

velocity. The connection between the previous two observations 

is not clear. The mechanism studied in this work may begin to 

explain how low levels of turbulence can cause large increases 

in heat transfer, since at any instant in time a turbulent flow will 

have spanwise variations. 

Future Plans 

Additional calculations will be run on a 2.25:1 and a 1.5:1 

elliptical leading edge. Experimental results will be compared to 

the calculations. Upon completion of the steady calculations, an 

investigation of the effect of unsteadiness on this mechanism will 

begin. To do the unsteady calculations, other numerical codes 

may need to be used. 

2.0 

1.5 Fr/F_—

114% 

Fr=Nu/(Re)"2 

0.0 

0.50 0.75 1.00 

Spanwise Distance, zJA 

A 3:1 elliptical leading edge with spanwise variation in velocity. The surface shows the heat transfer coefficient (red = high, 

blue = low). Reynolds number = 25,000, based on leading edge diameter. Spanwise wavelength = 0.1 D.

161
Scattering from Ocean Surfaces and Near-Surface Objects 

Charles L. Rino, Principal Investigator 

Co-Investigators: Hoc D. Ngo, Thomas L. Crystal, and Carl E. Hess 

Vista Research, Inc. 

Research Objective 

To develop a capability to simulate the scattering of electromagnetic frequency, acoustic frequency, and light from dynamically 

evolving nonlinear ocean surfaces and near-surface objects. 

Approach 

Direct numerical solutions to the surface-boundary integral 

equations remain impractical for realistic problems. However, as 

an extension of our work with multiple scattering between an 

object and a nearby rough surface, we have been exploring 

various techniques to segment the surface and coherently 

combine the subsolutions. If multiple scattering between the 

subsegments is ignored, we obtain the beam simulation technique. Bad experiences have forced us to use careful 

benchmarking. 

Accomplishment Description 

Previously, we had developed a family of codes that could 

numerically solve the surface-boundary value problem and the 

mutual interaction of an object near the rough surface. More 

recently, we have begun to explore scattering at near-grazing 

incidence, thus requiring progressively longer surfaces. To 

succeed at 5 degrees incidence, simulated surfaces require 

nearly 4,000 points and the scattering simulation needs to 

invert a 4,000-square complex matrix. This barely fits into the 

32 megaword queue (closer to grazing cases, for any significant 

roughness, we would move into the special queues). Coding 

optimization has reduced run times to 15 minutes, and the 

inversion is usable for simulating several incidence angles. 

Unfortunately, for proper statistics of rough-surface scattering, 

10-20 independent realizations are necessary. Then the whole 

parameter set must be repeated at different wind speeds for 

different rough nesses. 

Significance 

Both theory and numerical simulations have failed to satisfactorily address the grazing incidence problem. Even the small 

perturbation theory fails as strict grazing incidence is approached. 

Both the accuracy and the size of the surface needed to sustain 

resolution at small grazing angles cause problems. The beam 

simulation method, with corrections for multiple scattering, may 

provide a way to simultaneously increase numerical accuracy 

and resolution within practical computation limits. 

Future Plans 

We will continue to develop and test the methods with onedimensional simulations and, as encouraging results emerge, the 

method will be extended to two-dimensional surfaces.

Publications 

1. Rino, C. L.; Crystal, T. L.; Koide, A. K.; Ngo, H. D.; and 

Guthart, H. "Numerical Simulation of Acoustic and EM 

Scattering from Linear and Nonlinear Ocean Surfaces." Radio 

Science 26, no. 1(1991): 51-71. 

2. Rino, C. L. and Ngo, H. D. "Low-Frequency Acoustic Scatter 

from Subsurface Bubble Clouds." J. Acoust. Soc. Am. 90, 

no. 1(1991): 406-415. 

jgI1pIooIuI 

19.43 deg

kk , ^ 11 RINI om W Awu, ,] 20: 66JJJJJJJjJ1 

IIUII!I!1IAIUIi!I 

86.18 deg

Theta (deg) 

Simulation of a sinusoidal diffraction grating. Incident illumination is applied at each diffraction order. The grating lobe at 

86.18 degrees is just resolved as the incident illumination is 

brought inside of 5 degrees from grazing. 

162
Jet Interaction Aero-Optic Effects on Hypersonic Interceptors 

R. P. Roger, Principal Investigator 

Co-Investigator: S. C. Chan 

Teledyne Brown Engineering 

Research Objective 

To obtain flow-field predictions for missile control jetconfigurations that could adversely affect the performance 

of an on board infrared (IR) seeker. Computed solutions will 

provide input to a wave optics aero-optic code which predicts 

the influence of the flow field on a target signal at the seeker 

focal-plane array. Configurations of concern include lee-side 

firing of center of gravity (CG) thrusters at large angles of attack 

and firings at high altitude. 

Approach 

A three-dimensional full Navier—Stokes flow solver employing 

an upwind algorithm is being used to predict the complex jet 

interaction (Jl) flow field for advanced defense-interceptor 

concepts. Initially, computational fluid dynamics results are 

generated for wind tunnel test conditions to compare with test 

data (force and moment, surface pressure and temperature, and 

Schlieren photographs). As a result, projected tactical flight 

conditions will be simulated. 

Accomplishment Description 

One configuration is triconic with a single flat slot-cooled seeker 

window. A computational grid, modeling the window tetraconic 

forecone section and the mid and aft cone sections containing 

the CG thruster, has been constructed. The two sections are 

1.5 and 2.0 million grid points, respectively. A computational 

grid for a monoconic design containing about two-thirds as 

many grid points was also constructed. Computed results for the 

slot-cooled triconic forebody compare well with surface pressure 

and temperature distribution data. Computed results for the 

triconic thruster firing at 0 and 15 degree angles of attack also 

compared well to force and moment data and to Schlieren 

photographs. A grid resolution study was required to ensure 

accurate predictions for the extent of the recirculation region 

upstream from the jet. The tunnel JI cases did not include the 

slot-cooled window and incorporated 2.5 million grid points. 

The 0 degree case was performed in three computational blocks. 

It required 110 Cray Y-MP hours and 52 megawords of memory 

for the largest block. 

Significance 

Almost all proposed designs for advanced endo-atmospheric 

interceptor concept development employ IR seekers and 

CG thrusters. Use of CG lateral control during "end game" is 

preferred because of the reduced sensitivity to target miss 

distance. For maneuvering target intercepts, large lee-side angles 

of attack are possible. Wind tunnel tests indicate that the jet 

upstream recirculation region could extend onto the seeker 

window. We are investigating this phenomenon using our 

computations.

Future Plans 

Tactical flight-environment predictions and the incorporation of 

a transpiration-cooled mosaic-window biconic construction are 

planned. 

Publication 

Chan, S. C. and Roger, R. P. "Numerical Study of the Complex 

Flow-Field Features for a Supersonic Jet Exiting Normal to a 

Supersonic Stream." AIAA Paper 92-3675, July 1992. 

Surface and symmetry plane temperature Contours at 0 degree 

angle of attack for a tetracone triconic Configuration (nosetip 

blocked out).

163
Mixing and Reacting in 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 generate a direct numerical simulation (DNS) data base for 

reacting turbulent mixing layers to aid in the study of combustion. 

Approach 

We will focus on the low-heat-release limit of reaction in incompressible turbulence using DNS of the full three-dimensional 

Navier—Stokes equations with resolution adequate to resolve 

both the turbulence and the flame zone. Early work is based on 

reacting passive scalars in an incompressible turbulence using 

pseudo-spectral numerical methods. The mixing layers examined 

evolve temporally in an infinite cross-stream domain. 

Accomplishment Description 

A fully turbulent mixing-layer simulation with a visual thickness 

Reynolds number of up to 15,000 has been simulated. The 

simulation was begun from numerically simulated turbulent 

boundary layers with no additional disturbances. As the layer 

thickens in time, more computational modes are added in 

the cross-stream direction. Currently, the simulation uses 

512 x 210 x 192 modes (with 3/2 times as many physical 

space mesh points in each direction for de-aliasing of the 

nonlinear terms). Each sub-step takes 570 Cray-2 seconds and 

140 megawords of memory. To date, the simulation has 

required 750 Cray-2 hours. The streamwise one-dimensional

velocity spectrum exhibits nearly a decade of inertial range 

(Ic513 spectrum). Much of the vortex structure of the flow is 

similar to that of previous simulations begun from "clean" 

initial conditions, and the study of such "prototypical" flows is 

expected to aid in the understanding of this more complicated, 

fully developed turbulent case. A fast-reaction, low-heat-release 

diffusion flame is shown in the accompanying figure. The top 

three frames show side views of the layer at three different times. 

The clumping of product in "rollers" with high gradients in the 

"braid" regions is apparent. The bottom two frames show end 

views of the reaction product at the final time shown in the 

third side view. Mushroom structures of the kind observed in 

experiments are visible. At this point, the layer is evolving 

self-similarly. 

Significance 

This simulation is the first turbulent free-shear-flow DNS begun 

from realistic turbulent inlet and initial conditions and is the first 

to achieve self-similarity. The mixing layer has the highest 

Reynolds number of this flow ever achieved by DNS and is the 

only such simulation to exhibit an inertial range. 

Future Plans 

We will continue post-processing of the simulation and study the 

effects of reduced chemical-reaction rate. 

- ----------- 

wolw 

, 0,,O^ 

Product concentration (red = high and yellow = low) of a fast, low-heat-release diffusion reaction in a 

turbulent-plane mixing layer. The top three frames show a particular span wise location at three different times and the bottom two show end views at different stream wise locations at the last time shown. 

164
Hypervelocity Mixing and Combustion in Pulse Facility Flows 

R. Clayton Rogers, Principal Investigator 

Co-Investigators: David W. Riggins, Robert D. Bittner, and Glenn J. Bobskill 

NASA Langley Research Center/University of Missouri, Rolla/Analytical Services and Materials, Inc. 

Research Objective 

To determine the fluid dynamic characteristics and performance 

of hypersonic scramjet combustors, and to develop methods to 

accurately predict the combustor flow field and performance. In 

addition, we will study whether (1) flight Mach number affects 

combustor performance, (2) low-enthalpy fuel mixing simulations accurately represent fuel mixing in a scramjet combustor, 

and (3) ground-based test results represent flight performance. 

Approach 

Because of difficulty in measuring scramjet combustor performance directly, numerical solutions are performed to compare 

with limited measurements and are then used to predict combustor performance. Measurements used for comparison include 

wall pressure and heat flux, shadowgraphs, holograms, planar 

laser-induced florescence of hydroxyl, and laser absorption 

measurements of hydroxyl. 

Accomplishment Description 

Numerical and experimental results have been compared for 

several "unit" configurations with either one or two fuel injectors. Injector configurations studied include circular discrete 

flush-wall injectors, and swept and unswept ramp injectors. 

Numerical results compare favorably with the limited experimental measurements, providing confidence in the solutions. 

The current numerical solutions have been compared to 

experimental and other numerical solutions for both (1) lower 

Mach number tests and (2) low-enthalpy simulation (cold flow 

mixing at the same combustor Mach number) of the same 

configurations. A numerical flight-scaling methodology was 

developed and applied, including the effect of test pressure, 

temperature, test gas contaminants, and geometric scale in an 

incremental fashion. A typical job requires 20 megawords of 

memory and is run in 2-4 hour segments. 

Significance 

These computational studies, in conjunction with the highenthalpy experimental data, represent the first comprehensive 

examination of hypersonic scramjet combustor performance. 

Good comparison with experimental data indicates that turbulence models developed for lower Mach numbers are reasonably 

accurate at these very high Mach numbers. The accompanying 

figure illustrates a numerically generated flow field for a 

Mach 13.5 combustor with a swept ramp injector. Comparisons 

of experimentally backed numerical solutions demonstrate 

significant shortcomings in the simulation of combustor fuel 

mixing utilizing Mach similarity without accounting for velocity 

differences. Therefore, as shown in the figure, methods of scaling 

were developed to account for differences between low- and 

high-enthalpy fuel mixing.

Future Plans 

We will perform numerical sensitivity studies to determine the 

adequacy of experimental parameters that were measured to 

validate numerical models. 

Publication 

Riggins, D. W. and McClinton, C. R. "A Computational Investigation of Mixing and Reacting Flows in Supersonic Combustors." 

AIAA Paper 92-0626, Jan. 1992. 

Hydrogen-mass fraction Contours on a swept ramp injector base 

and three downstream cross-flow planes; pressure contours on 

inflow plane, symmetry plane, and bottom wall. For both 

parameters, red = high values and blue = low values. 

.8

Mach 6 

.6 

>. a 

C 

a

.4 

01 C

.2 

0 4 8 12 16 20 24

x(In.) 

Effect of combustor velocity on fuel mixing.

165
Hypersonic Inlet Flow Fields 

William C. Rose, Principal Investigator 

NASA Ames Research Center 

Research Objective 

To design aerodynamic contours for a potential wind tunnel and 

flight-test inlet that has flow fields typical of those expected to 

occur on hypersonic, air-breathing vehicles. 

Approach 

We will use the two- and three-dimensional full Navier—Stokes 

(FNS) internal flow codes run on the Cray-2 and Cray Y-MP. 

Multiple runs were made with the two-dimensional code to 

parametrically determine an aerodynamic design capable of 

operating at flight conditions at Mach numbers of 10 and 15. 

Several runs with the two-dimensional code were carried out in 

determining a suitable set of two-dimensional contours. These 

contours were then assumed to apply to a full three-dimensional 

symmetrical inlet with a sidewall swept from the ramp leading 

edge to the cowl lip. The flight-test parameters are those for the 

proposed NASA Hypersonic Flight Experiment. For these 

conditions a variable gamma was used, but no other real-gas 

considerations were invoked. The wind tunnel conditions were 

appropriate for the NASA Ames 3.5-Foot Hypersonic Wind 

Tunnel. Each three-dimensional job was run in a series of 

smaller steps, with each step taking approximately 2 hours and 

40 megawords of memory. 

Accomplishment Description 

Solutions were obtained for the Mach 15 configuration and one 

off-design Mach 10 case using the two-dimensional FNS code

and grid dimensions of 301 x 81. Mach number contours from 

the two-dimensional flow field at Mach 15 are shown in the 

accompanying figure. Because of the high dynamic pressure 

(about 7,000 psI) for one of the proposed flight trajectories, the 

structural loads on an inlet are significant. The maximum 

internal pressure loads occur when the inlet unstarts. The timeaccurate two-dimensional FNS code was run with downstream 

boundary conditions simulating an over pressurization of the 

combuster, causing the inlet to unstart. The surface pressure 

time histories on the ramp and cowl are shown in the figure 

inset. A large transient pressure rise is predicted. Approximately 

350 hours of single processor CPU time were used in the study. 

Significance 

This study has demonstrated the applicability of the twodimensional FNS code to determine aerodynamic contours for a 

Mach 15 inlet in order to properly account for large viscous 

effects. The time-accurate calculated-unstart transient pressure 

rise indicates that very large structural loads will occur. 

Future Plans 

The two- and three-dimensional FNS codes will be used to solve 

other three-dimensional inlet flow fields applicable to generic 

hypersonic vehicles. 

Contour levels 

(1. (11)000 

0.50000

Mach contours 

2 00000 

2.511000 

001)00 

D00 ou 

.L100u0 - b 500W LIE, a 

511001.1 

IJLIUI( 

I .3200 

OUIJU 

- 30000 

11100(3 

I Unstvt transient at minima,,, area 

11(0(311 

(11)013

Minimum area, time

2000 - Ramp 

/

• -' - - - -

Cowl 

1000 

Color 

Expanded vertical scale 0 64 1.28 1.92 2.56 3.20 

Time (msec)

(a) Hypersonic Flight Experiment Mach 15 inlet contours. (b) Expanded vertical scale. (c) Unstart transient at minimum area 

166
Mcnw 

=

wall 

Combustor entrance conditions 

Mach 10 hypersonic inlet.

National Aero-Space Plane Inlet Boundary-Layer Control 

William C. Rose, Principal Investigator 

NASA Lewis Research Center

Research Objective 

To compute inlet flow fields with novel boundary-layer control 

simulations. 

Approach 

We used the multiblock three-dimensional full Navier–Stokes 

code on the Cray-2 and Cray Y-MP with the boundary conditions applied to an inlet surface representative of those expected 

to occur from porous surfaces and surfaces with portions 

removed. Each three-dimensional job was run in a series of 

smaller steps with each small step taking approximately 2 hours 

and using 40 megawords of memory. 

Accomplishment Description 

The multiblock version of the SCRAM3D code was used to 

examine the nature of boundary-layer control as it applies to the 

sidewall vortical feature developed as a result of the swept 

multiple-shock-wave interactions with the sidewall boundary 

layer occurring in the Mach 5 National Aero-Space Plane 

(NASP) inlet. Conditions representative of those that occur in the 

NASA Lewis 10- by 1 0-Foot Supersonic Wind Tunnel were run. 

Calculations were carried out through the inlet throat, including 

a normal shock wave. The accompanying figure shows calculated Mach number contours to a station downstream from the 

throat. These contours indicate the persistence of the sidewall 

vortical feature throughout the inlet, so that the best control of 

the feature is either to not let it develop (cut away sidewalls) or 

to let it spill overboard (notched cowl surface). Approximately 

540 Cray-2 single processor hours were used. 

Significance 

The multiblock code has been validated for the complex threedimensional viscous-flow interactions between the captured 

internal flow and the external flow field that spilled over the 

sidewalls and cowl lip. As a result of this computation, modifications to the wind tunnel model have been made to improve inlet 

performance and these modifications are being tested in a 

second wind tunnel entry at NASA Lewis.

Future Plans 

The code will be used to investigate three-dimensional flow 

fields of interest generated in the second wind tunnel entry. 

Publication 

Rose, W. "Innovative Boundary Layer Control Methods in HighSpeed Inlet Systems—Final Report, Phase II." Prepared for NASA 

Lewis, Contract NAS3-25783, Nov. 1991. 

167
National Aero-Space Plane Inlet Flow Fields 

William C. Rose, Principal Investigator 

NASA Ames Research Center 

Research Objective 

To design aerodynamic contours for a wind tunnel test inlet 

model that has flow fields typical of those expected to occur on 

hypersonic air-breathing vehicles. 

Approach 

We used two- and three-dimensional full Navier—Stokes (FNS) 

internal flow codes on the Cray-2 and Cray Y-MP. Multiple runs 

were made with the two-dimensional code to parametrically 

determine an aerodynamic design capable of operating at 

conditions for the NASA Ames 3.5-Foot Hypersonic Wind 

Tunnel at Mach numbers of 10, 7.2, and 5. Several runs with the 

two-dimensional code were assumed to apply to a full threedimensional symmetrical inlet with a sidewall swept from the 

ramp leading edge to the cowl lip. The multiblock threedimensional code was run on these contours to examine inlet 

performance near the throat. Each three-dimensional job was run 

in a series of small steps, with each small step taking approximately 2 hours and 1 megaword of memory.

Accomplishment Description 

Solutions were obtained for the Mach 10 configuration and 

various off-design configurations for lower Mach numbers using 

the two-dimensional FNS code and grid dimensions of 301 x 61. 

Contours were generated for half of the three-dimensional 

symmetric inlet. The multiblock code was used to examine the 

three-dimensional flow field near the sidewall that is ultimately 

ingested into the inlet and remains started, thus indicating the 

viability of the design process developed in this study. Approximately 450 single processor hours were used in this study. 

Significance 

This study has demonstrated the applicability of the twodimensional FNS code in a parametric design setting to determine aerodynamic contours for an inlet in order to properly 

account for large viscous effects. 

Future Plans 

Both the two- and three-dimensional FNS codes will be applied 

to solving other three-dimensional inlet flow fields applicable to 

the National Aero-Space Plane and related vehicles. 

Cross-flow Mach-number contours and particle traces from a 3DFNS solution of a representative hypersonic inlet. 

168
Trapped-Vortex Flows on Highly Swept Wing Configurations 

James C. Ross, Principal Investigator 

Co-Investigators: Paula Lovely and Todd Riddle 

NASA Ames Research Center 

Research Objective 

To investigate the fluid mechanics involved in trapping a vortex 

on a wing. Generating high lift for landing and takeoff of HighSpeed Civil Transport aircraft with their highly swept wings is a 

difficult problem. A technique that may solve this problem is to 

trap a vortex on the upper surface of the wing. 

Approach 

The incompressible Navier—Stokes code, INS3D, was selected to 

compute the flow around wings tested in the NASA Ames 7- by 

1 0-Foot Subsonic Wind Tunnel. The tests were performed at low 

speed so the incompressible computations were appropriate. 

Emphasis was placed on gaining an understanding of the flow 

field and mechanisms involved in trapping a vortex. 

Accomplishment Description 

Computations were performed on a swept rectangular wing 

similar to the one tested in the wind tunnel. The computations 

allow a more detailed look at the flow field around the wing 

than was possible during the wind tunnel tests. The computed 

lift and drag increments due to vortex trapping were very close 

to the measured values. The accompanying figure shows top and 

end views of particle traces from the computations. The white 

surfaces are the vortex fences which help to generate and trap 

the vortex on the wing's upper surface. The character of the 

computed flow field is very close to that observed in the wind 

tunnel experiment. Closer examination of the flow showed that 

the vortex trapped on the wing was not as strong as expected 

and did not show a strong, tight core. 

Significance 

These results have shown some of the limitations of vortex 

trapping for generating high lift on swept wings. The good match 

between the computed and measured lift increments give good 

confidence in the computations.

Future Plans 

Computations will be used to examine the trapped-vortex 

concept and to look for better ways to accomplish effective and 

efficient vortex trapping. 

Publication 

Riddle, T. W. "A Numerical Analysis of Three-Dimensional 

Vortex Trapping." Master's thesis, California Polytechnic State 

University, San Luis Obispo, CA, March 1992. 

Top and side views of vortex trapping on a NACA 0012 wing 

swept 60 degrees at 0 degrees angle of attack at Re = 2,000,000. 

The white surfaces are fences to generate and trap vortex.

169
Validation of a Short Takeoff and Vertical Landing Model 

Karlin R. Roth, Principal Investigator 

Co-Investigator: Stephen Chiu 

NASA Ames Research Center/California Polytechnic State University, San Luis Obispo 

Research Objective 

To evaluate the ability of computational fluid dynamics (CFD)to 

predict the critical flow features for an aircraft configuration that 

is geometrically simplified, yet retains important aerodynamic 

and propulsive flow physics. 

Approach 

A model geometry consisting of a delta wing, tandem jets, and a 

blended fuselage is studied. The unsteady Reynolds-averaged 

Navier—Stokes equations are applied on patched or overlapped 

grids. Initially, an understanding of the flow physics and 

modeling requirements for the basic propulsive flow field is 

obtained, then the CFD technology is applied to the configuration. 

Accomplishment Description 

Numerical requirements for the propulsive flow field were 

established by computing the flow of a subsonic jet exhausting 

into a crossflow. The Chimera grid scheme was used with 

79 x 33 x 66 points in the Cartesian aerodynamic surface grid 

(positive half-plane) and 50 x 33 x 60 points for the embedded 

jet grid. Global flow features including the jet trajectory, the 

contra-rotating vortex pair, and the wake region downstream 

from the jet were simulated. The improved placement of grid 

points with the Chimera scheme coupled with Baldwin—Barth 

turbulence modeling for the flow over the surface allowed the 

horseshoe vortex upstream from the jet exit (a small scale 

feature) to be captured. Calculations using the OVERFLOW code 

required approximately 15 Cray Y-MP hours and 11 megawords 

of memory. In addition to the jet-in-cross-flow computations, the 

model geometry was extracted from the design data base and 

verified with the as-built model. Tunnel, jet, and vehicle grids 

were constructed and simulations have been initiated. 

Significance 

Validated CFD methods will aid in the design and development 

of short takeoff and vertical landing aircraft. 

Future Plans 

Simulation of the vehicle and jets will be completed with sonic 

jets at Mach 0.14 and 10 degrees angle of attack. The results will 

be compared with the measured flow field. 

Publications 

1. Roth, K. R. "A Powered-Lift Experiment for CFD Validation." 

AIAA Paper 91-1731, June 1991. 

2. Chiu, S. "Computational Investigation of a Three-Dimensional 

Navier—Stokes Model for a Subsonic Jet in Cross Flow." 

Master's thesis, California Polytechnic State University, 

San Luis Obispo, CA, May 1992.

Particle traces for computation with Mjet = 0.78 and M,., = 0. 13. 

Red particles are restricted to the surface; blue particles are 

released within the jet exit. 

An enlarged view of particle traces in the symmetry plane 

upstream from the jet exit. 

170
Coupled Rotation—Vibration—Dissociation Processes in Hypersonic Flows 

Stephen M. Ruffin, Principal Investigator 

Co-Investigators: Surendra P. Sharma and Ellis Whiting 

NASA Ames Research Center 

Research Objective 

To determine the effect of coupled rotation—vibration—dissociation 

phenomena on the behavior of dissociation and recombination 

rates in nonequilibrium relaxation, and to model the vibrational 

nonequilibrium in hypersonic flows using these rates. 

Approach 

The phenomena are investigated in two molecules: hydrogen 

(1­1 2 ) and nitrogen (N 2). State-to-state transition-rate coefficients 

for bound—bound and bound—free transitions for H 2 are determined using a quasi-classical trajectory method. This requires a 

potential energy function input which describes the interaction 

of the two rotating and vibrating H 2 molecules. Once rates for 

the transitions among all combinations of initial and final 

rotational, vibrational, and/or free states are obtained, the 

master equations are solved to yield the rotational and vibrational number densities as well as the bulk thermodynamic 

properties. For N 2 , the transition rates are computed using the 

Schwa rtz—Slawsky—Herzberg (SSH) theory. The rotation of N2 

molecules is not included in the model. To model the vibrational 

nonequilibrium in hypersonic flows, a quasi one-dimensional 

code for expanding flows is coupled with the master equations. 

An N2 molecule without rotation or dissociation is assumed. The 

transition rates at various temperatures are stored in a tabulated 

form and are interpolated and used as required by the fluid 

code. The code solves about 60 equations at each streamwise 

station and is computationally expensive. 

Accomplishment Description 

Bound—bound and bound—free transition rates for H 2 at temperatures of 10,000, 5,000, 3,000, 1,000 and 800 K have been 

computed. The master equations are solved for 10 cases in the 

500-6,000 K temperature range assuming a constant-volume 

isothermal heating environment. By monitoring the average 

rotational energy, the rotational relaxation times are deduced 

and are compared with the experimental data. Good agreement 

with experimental data is seen. Transition rates for nonrotating 

N2 and carbon monoxide molecules have been computed for 

10

vibration—translation and vibration—vibration transitions at 

various temperatures and have been compared with existing 

experimental data. Vibrational population distributions are 

computed in converging—diverging nozzles and in isothermal 

cooling simulations. The fully coupled solved simulation using 

SSH transition rates is able to predict the basic features of the 

vibrational relaxation, including population inversion; however, 

further study is needed for quantitative agreement. 

Significance 

Detailed studies of coupled rotation—vibration—dissociation 

processes are vital to our understanding of high-temperature flow 

physics. The computational procedure for the H 2 molecule will 

prove useful in predicting rotational nonequilibrium. A better 

understanding of rotational and vibrational relaxation will 

enable us to formulate accurate relaxation models and enhance 

our capability to compute nonequilibrium hypersonic flows. 

Future Plans 

The H2 model will be used to study the effects of rotation in 

expanding flows. A better model for N2 computation rates needs 

to be applied. The effects of rotation and dissociation will be 

included in the model. 

Publications 

1. Sharma, S. P. "Rotational Relaxation of Molecular H 2 at 

Moderate Temperatures." AIAA Paper 92-2854, July 1992. 

2. Ruffin, S. M. and Park, C. "Vibrational Relaxation of 

Anharmonic Oscillators in Expanding Flows." AIAA 

Paper 92-0806, Jan. 1992. 

- A/A*= 2, subsonic 

104—i -" AJA=1 

A/A= 2, supersonic 

- A/A= 4, supersonic 

- - - - A/A= 7, supersonic 

100 -L A/A= 10, supersonic

10-4

> Z

•1

L Boitnott and Warder 11 I*.0 

_ * Lensch and Gronig 

4

-a- Rabitz and Lam 0 ..... -.- Galiager and Fenn 

I .° - Present study .1 o. •_

1081 

10-12 

20 40 60 100X103 

1000 10000 Vib. energy level [cm-1] 

Temperature (K)

Computed carbon monoxide number densities in each vibraHydrogen rotational relaxation time as a function of temperature. tional quantum level in a converging—diverging nozzle.

171 
High-Alpha Flow Fields 

Christopher L. Rumsey, Principal Investigator 

Co-Investigators: James L. Thomas, Robert T. Biedron, Sherrie L. Krist, W. Kyle Anderson, Daryl L. Bonhaus, 

David L. Whitaker, Thomas W. Roberts, Gary P. Warren, Beyung Kim, and Robert P. Weston 

NASA Langley Research Center 

Research Objective CFL3D - wing alone 

To perform three-dimensional Reynolds-averaged Navier—Stokes • BART - wing alone 

computations for aerodynamic configurations at high angles of .......CFL3D - wing with tail 

attack.

BART - wing with tail 

CFL3D - tail vortex 

BART - tail vortex 

Approach 

The algorithm CFL3D solves the compressible three-dimensional 

Reynolds-averaged Navier—Stokes equations. Various upwind 

methods are used on the convective terms, including flux—vector 

splitting and flux—difference splitting. Viscous terms are centrally 

differenced and an implicit backward time discretization is used. 

Either an algebraic turbulence model or a half equation turbulence model can be used to simulate turbulent flows. The code 

has a generalized patched-grid multiblock capability to allow the 

treatment of general configurations. 

Accomplishment Description 

In an effort to explore the vertical-tail/vortex interaction phenomenon that occurs on F-i 5 and F-i 8 aircraft, computations are 

performed for a simplified geometry of a delta wing with a 

vertical tail surface downstream. An overlapped grid topology is 

employed, with the thin-layer Navier—Stokes equations solved 

on the delta-wing grid and the Euler equations solved on the tail 

grid. Preliminary computations indicate steady loads at angles 

of attack of 5 and 20 degrees and unsteady loads at 30 and 

40 degrees. Computed vortex trajectories agree qualitatively with 

trajectories seen in experiments. Computations are performed 

about an F/A- 18 wing/leading-edge-extension (LEX)/fuselage 

configuration and comparisons made with flight-test results. 

Excellent surface-pressure correlations were obtained on the 

forebody, but correlations degraded somewhat on the LEX 

surface, particularly near the wing—LEX juncture. The computed 

flow patterns on the forebody showed qualitative agreement with 

in-flight test results. Typical computations involve grids with 

between 300,000-2,000,000 grid points and require 1-8 Cray-2 

single-processor hours per case and 8-50 megawords of 

memory. 

Significance 

These computations aid in understanding the complex unsteady 

flows encountered over the F-i 8 at high angles of attack. 

Future Plans 

Future computations of the delta-wing/tail configuration will be 

time accurate at the higher angles of attack to capture the correct 

physics of the unsteady buffet phenomenon. Development of 

several unstructured-grid schemes for application to the F-i 8 are 

under way, and a hybrid scheme using Navier—Stokes equations 

on the upstream part and Euler equations on the downstream 

part of the vehicle are being developed to reduce the computational requirements.

.5 

.4 

.3 -J 

> .2 

.1 

0 

.4 

.3 

.2

N 1

0

—.1

.4 .8 1.2 1.6 

x/L 

Wing-alone and wing-with-tail configurations at 20 degrees 

angle of attack. 

Particle traces (yellow) passing through the primary and secondary vortex cores and at the leading edge of the tail and helicity 

contours at the trailing edge of the wing. 

172
Aerodynamic Inverse Design and Analysis for a Full Engine 

Jose M. Sanz, Principal Investigator 

NASA Lewis Research Center 

Research Objective 

To develop an aerodynamic design methodology for multistage 

turbomachinery that minimizes the use of experimental data. 

Approach 

The capability of the inverse design method has been demonstrated for blade sections and our goal is to extend it to fully 

three-dimensional multistage configurations. The method 

incorporates the three classic steps of aerodynamic blade design: 

preliminary design, blade geometry generation, and aero 

analysis. The procedure delivers aerodynamic blade and passage 

shapes and the corresponding flow solution. The inverse twodimensional design method is an excellent tool in expanding the 

design data base. Turning vanes with up to 60 degrees of flow 

are examples of this capability. For the two-dimensional method, 

we solved the potential flow equations because the objective 

was a shock-free, nonseparating airfoil. In three dimensions the 

assumption of irrotationality can no longer be made and the 

Euler equations are the proper vehicle for three-dimensional 

design problems. A surface pressure distribution is the threedimensional input driver. Extreme care must be used to determine a well posed problem. The input pressure distribution is 

used to drive the solution to the required design conditions. 

Accomplishment Description 

Timely use of this procedure requires its implementation on a 

parallel machine. A 25 MFLOPS processor per passage is a 

reasonable hardware environment. A typical engine may require 

at least 32 processors. To experiment with this idea, a cluster of 

seven IBM RISC/6000-550 powerstations was used to run a 

parallel version of the code. Parallel virtual machine is used as 

the message passage language; processors communicate via 

ethernet. The results show that four RISC/6000-550 processors 

perform at the equivalent CPU time of one Cray Y-MP processor 

for a coarse-grid test case. With a medium-size grid and 

enhanced vectorization, the ratio of RISC/6000-550s to one 

Cray Y-MP processor is six, and, with a fine grid, the ratio is 

eight. The two Cray Y-MPs at NASA Ames and NASA Lewis 

were added as nodes to the cluster. The host RISC/6000-550 

perfectly handles this heterogeneous cluster. The accompanying 

figure represents a schematic of the parallel cluster. NASA Lewis 

acquired 32 RISC/6000-550 processors and connected them to 

the existing configuration of RN/6000s and Cray Y-MPs. The 

cluster is performing at a sustained rate well beyond 3 GFLOPS.

Significance 

The use of parallel processing with advanced supercomputers 

makes possible the aero design of a full multistaged turbojet 

engine by direct simulation of the Euler equations. 

Future Plans 

This approach should lead to an efficient method of enlarging 

the design data base for turbomachinery blading. Various engine 

configurations will be tried to fully assess the capabilities of the 

method. 

Publication 

Sanz, J.; Pischel, K.; and Hubler, D. "The Engine Design Engine; 

A Clustered Computer Platform for the Aerodynamic Inverse 

Design and Analysis of a Full Engine." Cray Users Group, 

Washington, DC, Sept. 1992. 

ENGINE DESIGN ENGINE 

Host

-;------ 

/ 

V 

Nodes 1 !:!! I

Z_X 

=^ 0/4 

Schematic of the parallel cluster running one compressor stage 

and one and a half turbine stage.

173
NO BLOWING 

High Alpha Technology Program F-i 8 Aerodynamics 

Lewis B. Schiff, Principal Investigator 

Co-Investigators: Yehia Rizk, Ken Gee, and Scott Murman 

NASA Ames Research Center 

Research Objective 

To develop flight-validated design methodology to predict the 

aerodynamics and flight dynamics of aircraft maneuvering in the 

high-alpha regime in order to develop novel aerodynamic 

control concepts for increased aircraft maneuverability. 

Approach 

A Reynolds-averaged Navier—Stokes code (F3 D/Chimera) 

governing unsteady three-dimensional turbulent separated and 

vortical flows is used to predict the flow fields about the F-i 8 

High Alpha Research Vehicle (HARV) at high angle of attack 

flight-test conditions. The code utilizes the Degani—Schiff 

modifications to the Baldwin—Lomax algebraic eddy-viscosity 

turbulence model which accounts for the vortical flow above the 

aircraft at large incidence. 

Accomplishment Description 

The F3D code was extended to include additional geometric 

and flow features. The computational results for the aircraft at 

combined angles of attack and side slip (shown in the first figure) 

are in excellent agreement with HARV flight-test measurements, 

and computationally predict the changes in location breakdown 

of the wing leading-edge-extension (LEX) vortex with side slip. 

Time-accurate computations of the unsteady flow field downstream from the LEX vortex breakdown and the resulting 

unsteady loads imposed on the empennage are in reasonable 

agreement with wind tunnel measurements. Extensions to the 

code permit coupled simulations of the external and engine inlet 

flows. Computational results demonstrate significant effects of 

varying engine-mass flow rate on the LEX vortex structure. The 

F-i 8 isolated-forebody code was applied to investigate tangential 

slot blowing as a forebody flow-control concept for eventual 

flight evaluation. Computational solutions extended wind tunnel 

test results to full-scale flight conditions and developed a new 

slot configuration with a fourfold increase in slot efficiency. 

Recently, the slot blowing concept was evaluated in full-aircraft 

computations (shown in the second figure) that demonstrated the 

strong interaction of the displaced forebody aircraft vortices with 

the LEX and empennage. 

Significance 

Computational results obtained in the High Alpha Technology 

Program demonstrate the capability to reliably predict the highly 

complex three-dimensional separated and vortical flows 

characteristic of the high-alpha regime. Slot-blowing results 

are an initial application of computational fluid dynamics as a 

design tool, rather than an analysis tool, for the high-alpha 

regime. 

Future Plans 

The F3D code will be used for tip-to-tail computations of the 

F-i 8 configuration, with the empennage. Time-accurate computations will be used to investigate nonsteady flows and 

multidisciplinary computations coupling unsteady aerodynamics

with aircraft flight motions will be used to investigate performancelimiting phenomena. 

Flow field about an F- 18 with side-slip forebody and leadingedge-extension vortices. 

Off-surface instantaneous streamlines flow with tangential slot 

blowing. 

174
High Angle-of-Attack Vortex Flow Aerodynamics 

Lewis B. Schiff, Principal Investigator 

Co-Investigators: Yehia Rizk, Neal Chaderjian, David Degani, John Ekaterinaris, Yuval Levy, and Gabriel Font 

NASA Ames Research Center 

Research Objective 

To develop computational techniques to treat high-incidence 

flows and use them to investigate the complex unsteady 

separated and vortical flow fields surrounding geometrically 

simple wings, bodies, and wing—body combinations at large 

angles of attack. 

Approach 

Reynolds-averaged Navier—Stokes codes (F3D and NSS) governing unsteady three-dimensional laminar and turbulent separated 

and vortical flows are used to predict the flow fields about 

bodies of revolution and delta wings at large angles of attack. 

For turbulent-flow computations, the codes utilize the 

Degani—Schiff modifications to the Baldwin—Lomax algebraic 

eddy-viscosity turbulence model to account for the vortical flow 

structure which exists above the wings and bodies at large 

incidence. 

Accomplishment Description 

The F3D code was extended to investigate the effects of an 

upstream disturbance in generating asymmetric vortex flows 

about an ogive cylinder at incidence. As shown in the first 

figure, we investigated the physics governing tangential slot 

blowing as a means of generating asymmetric vortex flows on an 

ogive cylinder at large angles of attack. Also, a study was done 

investigating the ability of the Degani—Schiff turbulence model to 

compute turbulent vortex flows about an ogive cylinder at large 

incidence. The factors affecting vortex breakdown over double 

delta wings was extensively studied and was extended to 

investigate the unsteady vortical flow surrounding double delta 

wings undergoing oscillations in pitch. An analogous NSS code 

computational investigation of unsteady vortical flow over a 

delta-wing/body (shown in the second figure) undergoing largeamplitude high-rate roll oscillations characteristic of wing rock 

was also carried out. 

Significance 

The computational results yielded new insight into steady and 

unsteady high-alpha flow phenomena. Computational investigation of the onset of vortex asymmetry has led to better understanding of conflicting results in the experimental measurements. 

Computations of asymmetric turbulent separated and vortical 

flows over bodies at large incidence confirm the ability of this 

model to treat both asymmetric and symmetric separated and 

vortical flows. 

Future Plans 

The F3D and NSS codes will be used for time-accurate computational investigations of unsteady flow response to prescribed 

large-amplitude high-rate vehicle motions of simple configurations. Multidisciplinary computations coupling unsteady 

aerodynamics with vehicle dynamic equations will be used to 

investigate significant flight-dynamic problems.

sr1iit.s. RD = UUJ. lJ_._ 11.2damillr 

HiTTT:..:TT I? = 2'.\!_. = .2 tpr 'al. nt.

Surface flow comparison (side view). 

Surface flow patterns colored by pressure coefficient.

175
Viscous Non-Reacting Flows in High-Speed Combustors 

Balu Sekar, Principal Investigator 

Co-Investigators: Douglas Davis and Daniel Risha 

WLIFIMM, Wright Patterson AFB 

Research Objective 

To apply state of the art three-dimensional computer codes to 

predict the flow development and performance of high-speed 

propulsion components such as inlets, supersonic combustors, 

and nozzles. 

Approach 

The three-dimensional time-dependent compressible 

Navier—Stokes equations are solved with algebraic turbulence 

models and with fully coupled chemistry for different reaction 

mechanisms to compute the mixing and diffusion phenomena 

that occur in a typical scramjet combustor. 

Accomplishment Description 

The General Aerodynamic Simulation Program (GASP) was used 

to simulate the twin-jets staged injection into a Mach 2 cross 

flow behind a step. GASP has the capability of a general 

multiblock (zonal) feature that allows individual blocks of grids. 

Various boundary conditions and the ease of applying these for 

complex three-dimensional arbitrary internal/external configurations and implementation of efficient numerical algorithms make 

this code particularly attractive and suitable for high-speed 

combustor modeling. Numerical studies done for the twin-jet/ 

combustor configuration were used to determine key performance parameters such as penetration and mixing. The accompanying figure illustrates the computed injected air-mass fraction 

distribution inside the combustor at selected axial locations. 

Flow convergence for this configuration of two zones with 

37 x 8 x 44 and 37 x 68 x 63 grids was obtained after 

3,000 iterations using 16 single-processor Cray Y-MP hours 

and 15 megawords of memory. The computed flow compares 

moderately well with experiments. 

Significance 

Enhanced flow-prediction methods and codes benefit the design 

of all air-breathing hypersonic/supersonic propulsion systems 

and are essential for the design and development of supersonic 

combustors having a high degree of mixing and overall combustion efficiency. 

Future Plans 

Research is now under way to determine the relative performance, the improvements of different combustors, and the fuel 

injector conceptual designs applicable to scramjet propulsion. 

This will enhance the operating characteristics of the propulsion 

systems due to increased mixing and combustion and a reduction in the net flow loss.

Publication 

Sekar, B. "GASP 1 .2d Results for the Transverse Air-jet Case." 

Presented at the 3rd Scramjet Modeling Workshop by the 

CFD Group, Wright Labs/Special Projects Office, Reno, NV, 

Jan. 1992. 

Density at selected stream wise stations in the injected air mass 

fraction. 

Pressure comparison at the injector plane—PIP infinity contours. 

176

(?-3
Computational Fluid Dynamics Approach to Computational Electromagnetics 

Vijaya Shankar, Principal Investigator 

Co-Investigators: William F. Hall, Alireza Mohammadian, and Chris Rowell 

Rockwell International Science Center 

Research Objective 

To develop a three-dimensional electromagnetics capability to 

solve the time-domain Maxwell's equation employing the 

algorithmic rigor of computational fluid dynamics (CFD) 

methods. Such a capability on supercomputers is critically 

needed for many large-scale problems encountered in the 

development of low observable platforms. 

Approach 

Maxwell's equations cast in conservation form are solved using a 

finite-volume Lax—Wendroff upwind scheme. The formulation 

accounts for any variations in the material properties (time, 

space, and frequency dependent), and can handle thin resistive 

sheets and glossy coatings. The time-domain approach handles 

both continuous wave (single frequency) and pulse (broad-band 

frequency) incident excitation. Arbitrarily shaped objects are 

modeled using a body-fitted coordinate system. For treatment of 

complex internal/external structures with different material 

layers, a multi-zone framework with the ability to handle any 

type of zonal boundary conditions is implemented. The timedomain electromagnetic-field values are converted to the 

frequency domain using fast-Fourier-transform routines, and then 

a Green's-function-based near-field to far-field transformation is 

employed to obtain the radar cross section (RCS). 

Accomplishment Description 

The three-dimensional computational electromagnetics (CEM) 

code is highly vectorized and parallelized for Cray Y-MP and 

C90 architectures. On an eight CPU Cray Y-MP using the 

autotask mode the code runs near 2 GFLOPS, and over 

10 GFLOPS on the C90 with 16 CPUs. The code is presently 

undergoing the validation phase while being applied to many 

problems of interest to low observable (LO) design. The accompanying figure shows a typical RCS calculation for a finned 

projectile with over 1 million grid points. For every grid point, 

the code needs 40 megawords of memory. The execution time 

required depends on the problem size and ranges from 

a few seconds to tens of hours. Computations of complete 

LO platforms at frequencies of interest will require a teraflop 

capability.

Significance 

CFD-based methods for solving problems in other disciplines 

provide the commonality and synergism required for establishing 

a computational environment to perform multidisciplinary 

computations. This CEM capability fully exploits the power of 

CFD algorithms that have evolved over 30 years. 

Future Plans 

We will extend this CEM development to an unstructured grid 

cell and multiple-instruction/multiple-data architectures. 

Radar cross-section studies for a finned projectile.

177
Computational Fluid Dynamics of Store Separation 

Samuel P. Shanks, Principal Investigator 

Co-Investigator: Jasim U. Ahmad 

JAI Associates, Inc. 

Research Objective 

To develop a three-dimensional dynamic Chimera algorithm 

that solves the thin-layer Navier—Stokes equations over multiple 

moving bodies in order to numerically simulate the aerodynamics, missile-body dynamics, and plume of a missile separating 

from an aircraft. This procedure will be used to calculate 

powered missile separation from an aircraft and the subsequent 

effect of the missile on the aircraft and aircraft engine plume 

ingestion. 

Approach 

A three-dimensional Chimera algorithm that solves the thin-layer 

Navier—Stokes equations over multiple-moving bodies was 

modified to include a total-variation-diminishing (TVD) Roe 

upwind method, a missile with a plume, and the subsequent 

propulsion forces that move the missile. Missile and plume grids 

were generated and imbedded in the aircraft grid. 

Accomplishment Description 

To demonstrate the capability of the procedure, a missile was 

launched from a wing and a missile was launched from the 

fuselage of a F/A-i 8 aircraft. Several static solutions of the TVD 

Roe upwind Chimera scheme were run. The results of the missile 

launch from the fuselage of the F/A-118 aircraft are shown in the 

accompanying figure. Pressure contours through the axis of the 

missile/plume and on the fuselage surface are shown for several 

time slices over the first 0.4033 seconds of the missile launch 

from the F/A-i 8 aircraft. This calculation had 750,000 grid 

points and took 23 Cray Y-MP hours. 

Significance 

The powered missile separation procedure has demonstrated a 

capability to analyze missile separation from aircraft and engine 

ingestion. With improvements and code validation, the powered 

missile separation procedure could be used to generate aerodynamic data bases for new missile concepts; this data base would 

be less expensive than data bases generated from flight-test 

programs. Also, the missile need not be built to generate the data 

base. The data base could be used in faster missile separation 

procedures that have a simplified aerodynamics package. As 

computers become faster, the procedure will replace simpler, 

but less accurate, missile separation procedures.

Future Plans 

Modifications will be made to the procedure to include real-gas 

effects, and missiles with fins will be launched from aircraft. 

Eventually, non-rigid aircraft and missiles will be included in the 

procedure.

r 

Pressure contours through the axis of the missile plume and on 

an F/A- 18 fuselage surface for several time slices. 

178
Plane of 

symmetry 

Wankel Engine Flow Fields 

Tom I-P. Shih, Principal Investigator 

Co-Investigators: E. Steinthorsson, Z. Li, A. Karadag, E. A. Willis, and J. McFadden 

Carnegie Mellon University/NASA Lewis Research Center 

Research Objective 

To develop numerical codes that can be used to study the flow, 

spray, and spray combustion in Wankel engine combustion 

chambers and to use the codes to study the flow physics as a 

function of engine design and operating parameters. 

Approach 

We will develop and evaluate methods for calculating unsteady 

compressible flows in deforming spatial domains. Parametric 

studies showing the effects of engine design and operating 

parameters on the flow will also be performed. 

Accomplishment Description 

Three codes, Lewis-2D, Lewis-3D, and LeRC3D, were developed. Lewis-2D and Lewis-3D can only calculate nonreacting 

flows and are based on an alternating-direction implicit finitedifference algorithm using flux-vector splitting. LeRC3D can 

calculate reacting and nonreacting flows. This code uses an 

implicit finite-volume method that can be a lower—upper 

algorithm, a point-iterative process, or a line-iterative process 

with a variety of differencing schemes for the convection terms, 

including several different flux-vector splitting schemes and the 

advection upstream splitting method developed by Lion and 

Steffen. 

Significance 

These codes can be used to design more efficient, less polluting 

Wankel rotary engines. 

Future Plans 

We will improve the modeling of turbulent combustion and 

droplet—turbulence interactions.

Apex seal 

Exhaust '\\ + y Intake 

__________

Port

Rotor 

;j II housing 

AL 

AT 

Apex —N - Apex 

seal seal 

(a) (b) 

,%0

I 

k=11 =20 

(c) (d) 

Wankel engine schematic and flow fields; (a) Wankel engine studied, (b) grid system, (c) velocity field at plane K = 11, and (d) velocity 

field at i = 20.

179
Formation and Growth of Hairpin Vortices 

B. A. Singer, Principal Investigator 

Co-Investigators: R. D. Joslin and S. P. G. Dinavahi 

NASA Langley Research Center 

Research Objective 

To study the formation and growth of hairpin vortices in the late 

stages of boundary-layer transition. 

Approach 

We used direct numerical simulation of boundary layers with 

wall-surface disturbances to create artificial hairpin vortices and 

analyzed data bases generated from previous numerical simulations with other types of disturbances to find similar structures. 

Accomplishment Description 

Hairpin vortices were generated with a variety of wall disturbances. While the general appearance of the structures was 

similar, the development of the vortices was sensitive to the 

geometry and intensity of the wall disturbance. Very weak 

disturbances dissipated after the wall disturbance was removed; 

somewhat stronger disturbances led to the appearance of 

secondary and tertiary vortices. Preliminary work indicates that 

these stronger disturbances will form the core of a turbulent spot. 

The details of the process are sensitive to grid resolution, so 

extensive studies have been done to determine the minimum 

grid requirements necessary to capture the essential physics of 

the process. A typical grid has 261 x 161 x 121 points in the 

streamwise, wall-normal, and spanwise directions, respectively, 

resulting in computer memory requirements just under 

125 megawords. At this resolution, approximately 4 Cray-2 

hours are needed to advance the solution one nondimensional 

time unit (tU/ö*). The accompanying figure shows the flow 

structure at a time of 56.54. The gray regions show the core of a 

hairpin vortex as marked by regions of low pressure. The primary 

vortex has spawned a secondary vortex high in the boundary 

layer and a tertiary vortex rotating with the opposite sense (as 

shown by the red velocity vectors) lower in the boundary layer. 

Significance 

A better understanding of the importance of hairpin vortices to 

the late stages of transition to turbulence can yield important 

insights into the mechanisms responsible for turbulent production and spreading. It is anticipated that these insights will result 

in more realistic models for predicting flows in the transition 

region. 

Future Plans 

The simulations will be continued until the flow becomes 

turbulent, allowing us to study some of the relationships between 

mature hairpin vortices and turbulent flow. Preliminary work to 

study these localized structures in a supersonic flow has begun.

Publications 

1. Singer, B. "The Formation and Growth of a Hairpin Vortex." 

To be published in Transition and Turbulence, eds. M. Y. 

Hussaini and R. G. Voigt. Springer-Verlag, 1992. 

2. Singer, B. "Formation and Growth of Hairpin Vortices in a 

Boundary Layer." Bulletin of the American Physical Society 

36, no. 10(1991): 2644. 

Hairpin vortex flow structure. 

180
Computational Fluid Dynamics Methods for Highly Maneuverable Aircraft 

Brian R. Smith, Principal Investigator 

Co-Investigators: C. L. Reed and A. Muyshondt 

General Dynamics, Fort Worth Division 

Research Objective 

To develop a Reynolds-averaged Navier—Stokes solver for the 

analysis of propulsion-related flow fields, and to assess the 

accuracy of computational fluid dynamics (CFD) solutions on 

these complex flows. 

Approach 

A finite-volume Navier—Stokes solver with the diagonally 

inverted lower—upper implicit scheme of Yokota and Caughey is 

used. This code includes a two-equation turbulence model and 

wall-function boundary conditions. 

Accomplishment Description 

Several complex flows were studied for this project. An advanced 

compression surface inlet with a turbulent-separated-shockwave/turbulent-boundary-layer interaction was successfully 

simulated. A high-aspect-ratio nozzle with one vertical tail and 

two horizontal tails was also analyzed. The results compare well 

to an experimental investigation conducted in the NASA Langley 

16-Foot Transonic Wind Tunnel. The experiment determined the 

effects of nozzle aspect ratio, empennage surfaces, and vectoring 

on twin engine aircraft afterbody aerodynamics. Experimental 

data included both surface-pressure data and balance data. The 

nozzle analyzed computationally had an exit aspect ratio of 

5.378. The design pressure ratio of this nozzle was 5.6, which 

simulated a dry-power operating mode. For this test case, a 

Mach number of 1.2 and a nozzle pressure ratio of 5.6 were 

selected. The nozzle vector angle was 0 degrees and the angle of 

attack was 0 degrees. The tunnel conditions were Mach 1.2 and 

Reynolds number 4.17 million per foot. The solution was 

obtained by running the Navier—Stokes code, including all 

viscous terms, and making use of the two-equation turbulence 

model. The solution was executed to 4,000 iterations, at which 

point the residual of the mean flow equations had dropped 

3 orders of magnitude and the residual of the turbulence 

equations had dropped 2.5 orders of magnitude from their 

initial values. In addition to monitoring the residuals, the nozzlemass flow and external pressures were monitored during the last 

half of the run to ensure convergence.

Significance 

This project has demonstrated the accuracy of Navier—Stokes 

analysis for complex three-dimensional propulsion applications. 

This solution provides details of the flow field which are difficult 

to obtain experimentally, but can aid the designer. 

Future Plans 

Inlet flow-field prediction is the most challenging CFD analysis 

problem for highly maneuverable aircraft and will be the focus 

of the continuing project. Predictions of flows with shock/

boundary-layer interactions and strong curvature will be 

analyzed. Advanced turbulence models will be tested and 

evaluated on complex inlet flow fields. 

Publication 

Reed, C. L. and Muyshondt, A. "CFD Calibration for ThreeDimensional Nozzle/Afterbody Configurations." AGARD 

CP-498, Fort Worth, TX, Oct. 1991. 

in 

'I'll 

M1 

Analysis of a high-aspect-ratio nozzle with tails.

181
Numerical Experiments in the Formation and Evolution of Galaxies 

Bruce F. Smith, Principal Investigator 

Co-Investigator: Richard H. Miller 

NASA Ames Research Center/University of Chicago 

Research Objective 

To develop and utilize numerical experiments to study the 

important dynamic processes at work in the formation and 

evolution of galaxies. 

Approach 

The numerical experiments are based on the time development of a fully three-dimensional self-gravitational and selfconsistent particle system. The codes can follow the motions 

of 105-106 particles. The gravitational potential and forces are 

calculated on a 2563 grid. 

Accomplishment Description 

The three-dimensional particle code has been used on the 

Cray-2 and Cray Y-MP to study several important problems in 

galactic dynamics. Experiments on the nature of disk systems 

embedded in elliptical galaxies have resulted in fascinating new 

insights on the dynamic behavior of these systems. An unexpected feature is that the nucleus of the galaxy orbits around the 

mass centroid. These systematic motions can produce effects 

such as off-set centers which are often seen in observations of 

galaxies. Also found were long-lived global galactic oscillations 

which are normal modes of a stable system near equilibrium. 

These studies require large numbers of particles and stable 

experiments that extend over many crossing times. 

Significance 

Observations of galaxies give us intriguing glimpses of complex, 

evolutionary sequences. The numerical experiments are an 

effective approach in understanding the dynamic sequences in 

galactic evolution. The computational capability is necessary to 

understand the wealth of new astronomical data now available. 

The experimental results on motions at the centers of galaxies 

imply that galaxies are not in a static steady state. This affects our 

interpretation of activity seen at the centers of galaxies, including 

our own. 

Future Plans 

We will incorporate additional physics into the code. The gas 

dynamics representing the gaseous component in galaxies is 

currently being incorporated and will allow for galaxy formation 

studies and studies of galaxy interactions with highly stimulated 

star formation. High-resolution studies of the interplay between 

stellar and gas dynamics in galaxies are now possible.

Publications 

1. Miller, R. H. and Smith, B. F. "Off-Center Nuclei in Galaxies." 

Astrophysical Journal 393 (1992): 508. 

2. Miller, R. H. and Smith, B. F. "Longevity of Disks in Galaxies." Seventh Florida Workshop in Nonlinear Astronomy on 

Astrophysical Disks. 1991. 

3. Smith, B. F. and Miller, R. H. "Galactic Oscillations." 

Presented at the Third Teton Summer School, July 1992. 

U) - 

X 

U) 

0) 

C w

0 200 400 600 800 1000 1200

Time 

The lower track shows time dependence of the total kinetic 

energy from an experiment based on a spherical galaxy model. 

The upper track shows the total potential energy with its sign 

reversed. The zero is not suppressed in this plot. The kinetic 

energy variation is quite large and it shows no decrease in 

amplitude over 15 cycles of the kinetic energy oscillation. This 

graph represents about 50 crossing times (5 billion years for the 

crossing time of a typical galaxy). 

182
YAV-813 Aircraft Flow Simulation 

Merritt H. Smith, Principal Investigator 

Co-Investigators: Kalpana Chawla and William R. Van Dalsem 

NASA Ames Research Center 

Research Objective 

To validate modern computational fluid dynamics (CFD) 

methods for the simulation of the flow about complete vectoredthrust aircraft, and to investigate the interactions of the aircraft's 

major systems. 

Approach 

The flow about the YAV-813 Harrier is simulated with the 

conditions of 30 knots forward speed and 30 feet altitude. The 

Navier—Stokes equations are solved on a system of 18 overset 

structured grids containing 2.8 million grid points and using the 

diagonalized Beam—Warming option in the OVERFLOW flow 

solver. Inlet and exhaust boundary conditions are set using a 

component performance model of the Harrier's Pegasus engine. 

Accomplishment Description 

Comparisons to infrared images of the flight vehicle at low-level 

jet-borne flight conditions shows qualitative agreement regarding 

jet trajectories and forward penetration of the rear jets. The 

propulsion-induced lift loss, known as the "suck-down" effect, 

has also been predicted. A modification to the time-stepping 

procedure in OVERFLOW has resulted in a significant savings in 

CPU time. The majority of grid points are solved using a large 

constant time step to accurately convect coherent structures, 

while the few smallest grid cells are solved at a constant Courant

number to reduce nonlinear instabilities. The total solution time 

has been reduced to approximately 45 Cray Y-MP hours and 

13 megawords of memory. 

Significance 

The performance levels that will be demanded of future aircraft 

will require the careful integration of aircraft systems in order to 

minimize detrimental interactions like the suck-down effect. 

A flexible, accurate, and timely method for the prediction of 

these interactions will be an absolute necessity in this design 

environment. 

Future Plans 

A k-c turbulence model, recently developed for jet flow fields, 

will be added. Aircraft systems, in addition to the engine and 

external aerodynamics, will be modeled to study a wider range 

of interactions. Finally, more flight conditions will be simulated 

to improve confidence in the capability. 

Publication 

Smith, M. H.; Chawla, K.; and Van Dalsem, W. R. "Numerical 

Simulation of a Complete STOVL Aircraft in Ground Effect." 

AIAA Paper 91-3293, AIAA 9th Applied Aerodynamics Conference, Baltimore, MD, Sept. 1991. 

Particle traces colored by time since release for the flow about the YAV-8B Harrier at flight conditions of 30 knots and 30 feet.

183
Grid Generation for Aerodynamic Configurations 

Robert E. Smith, Principal Investigator 

Co-Investigator: Michael J. Bockelie 

NASA Langley Research Center/GEOLAB 

Research Objective 

To develop the mathematical theory and computer software to 

generate structured and unstructured grids about aerodynamic 

configurations. A second objective is to apply existing finitevolume computer codes to compute the inviscid and viscous 

flow fields about these configurations. 

Approach 

A series of solution-adaptive structured grids were generated for 

a Martian entry vehicle (MEV) design. The adaptive method uses 

a background grid control that retains the original grid spacing in 

regions of small solution gradients and employs a network data 

base representation of the vehicle surface to ensure the new 

surface grid points lie on the original surface. The thin-layer 

Navier—Stokes equations were solved on the adaptive grids using 

the TLNS3D flow solver. Unstructured grids were generated for 

several simple aerodynamic configurations using the VGRID3D 

advancing-front grid generator. Inviscid solutions were computed on the unstructured grids with the LJSM31D flow solver. 

Accomplishment Description 

The adaptive solution method was tested by computing the 

steady, hypersonic, viscous flow field about the MEV at several 

angles of attack. A typical solution for the MEV required 

300,000 grid points, 20 megawords of memory, and 1.25 Cray-2

hours. Unstructured grids and inviscid solutions were computed 

for several two- and three-dimensional geometries. About 

30 Cray-2 hours were used for the unstructured-grid computations. 

Significance 

A comparison of adaptive and nonadaptive solutions clearly 

showed that the adaptive solutions captured a sharper shock and 

more accurately predicted the stagnation point values at the nose 

of the vehicle. 

Future Plans 

We will generate unstructured grids for aerodynamic configurations. The VGRID3D code is being modified to use surfaces 

defined with data base networks of bi-cubic hermite patches in 

order to produce smoother and more accurate surfaces. 

Publications 

1. Bockelie, M. J. and Smith, R. E. "An Adaptive-Grid Method 

For Computing the Three-Dimensional Viscous Flow about a 

Re-Entry Vehicle." AIAA Paper 92-2685, 1992. 

2. Bockelie, M. J. and Eiseman, P. R. "An Adaptive-Grid Method 

For Unsteady Flow Problems." To be published in International Journal of Numerical Methods for Heat and Fluid Flow, 

1992. 

Coefficient of pressure for a Martian entry vehicle adaptive-grid solution (left) and for an M6 wing unstructured-grid solution (right). 

184
National Aero-Space Plane High-Speed Combustion Experiment 

K. Snyder, Principal Investigator 

Co-Investigators: J.-L. Cambier, D. K. Prabhu, and S. Tokarcik 

NASP Joint Program Office/NASA Ames Research Center 

Research Objective 

To study the start-up and choking phenomena in a hypersonic 

shock tunnel; to characterize the real-gas aerothermodynamic 

environment in the nozzle of the NASA Ames 16-Inch Hypersonic Shock Tunnel facility; to study the inlet performance of the 

National Aero-Space Plane (NASP) combustor to be tested in the 

shock tunnel; and to aid in the design of diagnostic experiments. 

Approach 

Two-dimensional axisymmetric time-accurate explicit/implicit 

upwind total-variation-diminishing codes with multiple grid 

capability are used to solve the Navier—Stokes equations for 

chemical nonequilibrium. The computed flow fields are used in 

line-of-sight high-resolution spectral codes. 

Accomplishment Description 

The start-up transient in the shock tunnel nozzle was computed 

starting with the primary shock at the end of the shock tube 

(driven end) and continuing until the shock reached the end of 

the nozzle. The temperature contours at various time intervals in 

the end of the driven tube and part of the facility nozzle are 

shown in the accompanying figure. The reflection of the primary 

shock off the end wall of the driven tube is clearly seen. The test 

section with a simple cone model was introduced and the 

DRIVEN TUBE NOZZLE DRIVEN TUBE 

H

REFLECTED SHOCK 

110 

II la sec 

computation continued until the shock exited the test section. 

The computed real-gas environment was then used in the 

computation of the flow in the NASP combustor inlet. The 

computed cowl shock points of impact on the body surface for 

various cowl lengths were used in the final choice of inlet 

geometric parameters. 

Significance 

The numerical computations provide information for detailed 

studies of the start-up procedure and provide good estimates of 

facility test times for various operating pressures. They also 

provide information about the real-gas environment that cannot 

be determined experimentally. The combustor-inlet performance 

study is used to determine the final size of the inlet to be tested. 

Future Plans 

The axisymmetric code and a preliminary three-dimensional 

code were given to the University of Maryland to study the 

choking phenomena for realistic three-dimensional geometries. 

The axisymmetric code will be used in the pretest calibration of 

the shock tunnel for various nozzle shapes and operating 

pressures. The performance of the actual three-dimensional 

combustor inlet to be tested in the shock tunnel 

will be assessed.

NOZZLE 

;-

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701.LSOC 

DRIVEN TUBE NOZZLE DRIVEN TUBE NOZZLE 

Start-up transient temperature contours in the NASA Ames 16-Inch Hypersonic Shock Tunnel. 

Driver pressure = 408.2 atm, enthalpy = 10.5 Mi/kg, and shock speed = 3 km/sec.

185
25 

20 

15 

10

I 

;ing 

e 

it 

Complex Turbulent Boundary Layers 

Approach 

Direct simulation (three-dimensional Navier—Stokes solutions 

with no turbulence model) will be applied using an existing 

spectral code with a "fringe" procedure that allows it to treat 

fairly general flows as long as the surface has no curvature and 

the fluid is incompressible. + 

Philippe R. Spalart, Principal Investigator 

Boeing Commercial Airplane Group 

Research Objective 

To obtain highly accurate and complete flow-field information 

for turbulent boundary layers with pressure gradients at moderate Reynolds numbers. Theories including pressure gradient are 

not complete and experiments suffer from measurement uncertainties and the inaccessibility of the wall region. Simulations 

suffer from limited Reynolds numbers and geometry. As a result, 

joint studies carry considerable weight. 

Accomplishment Description 

The study of a two-dimensional flow that Watmuff conducted at 

NASA Ames is now complete. Preliminary results show that the 

Reynolds numbers match and that 25 million grid points are 

needed. Detailed comparisons reveal differences of no more 

than 15% for the velocity and skin friction. The more striking 

finding is that the velocity profile normalized by wall variables, 

as a function of y, is seen to drift down in the buffer layer 

and log layer (8 < y1 < 70) when the pressure gradient is 

adverse. This is shown in the accompanying figure and is in 

close agreement with experiment. The simulation used several 

hundred Cray Y-MP hours, 4 megawords of memory, and 

60 megawords of solid-state-device memory. 

Significance 

Depending on the theory, the U profile in the buffer and log 

layers was believed to be universal, or even to drift up in an 

adverse pressure gradient. The present finding invalidates some 

of the theories and suggests caution in the use of wall functions. 

Combined with the experimental data, this study provides strong 

quantitative evidence that will be used to refine turbulence 

models.

Future Plans 

The triggering of instability waves in laminar boundary layers by 

acoustic waves and surface imperfections (receptivity) will be 

studied in two and three dimensions. Transition caused by large 

disturbances, such as suction holes, will also be considered. 

1 10 100 1000

Y+ 

Velocity profiles in a turbulent boundary layer with pressure 

gradient and computed by direct numerical simulation. 

tE
Free-Surface Effects on Boundary Layers and Wakes 

Fred Stern, Principal Investigator 

University of Iowa 

Research Objective 

To develop computational methods to solve the unsteady threedimensional Navier—Stokes equations for surface-piercing bodies 

with free-surface gravity waves. Areas of specific interest are: 

detailed resolution of the solid—fluid juncture boundary layer 

and wake with waves; kinematics and dynamics of waveinduced separation; bow flow; and the influences of yaw. 

Approach 

The computational method solves the Reynolds-averaged 

Navier—Stokes equations in numerically generated, body/freesurface-fitted, nonorthogonal coordinates for unsteady threedimensional flows. A regular grid is used with finite-analytic 

discretization and pressure-implicit-split-operation-type 

velocity-pressure coupling. A two-equation turbulence model 

is used with either wall functions or a one-equation model for 

the near-wall flow (two-layer model). Various treatments for 

the free-surface boundary conditions and both small-domain 

viscous—inviscid interactive and viscous large-domain approaches 

are being developed. These studies are guided and supported by 

a concurrent program of towing-tank experiments. 

Accomplishment Description 

Computational studies and comparisons with the experimental 

data were completed using first-order free-surface boundary 

conditions both for an idealized Navier—Stokes-wave/flat-plate 

(Sw/fp) and realistic geometries. The calculations for the Sw/fp 

closely agree with the foil-plate-model wake data and aid in 

explication of the near and intermediate wake characteristics, 

periodic nature of the far wake, and wave-induced separation. 

Interactive approach calculations for realistic geometries enabled 

validation of the overall approach and an evaluation of wave 

boundary layer and wake interaction. Recent calculations 

include exact treatment of the free-surface boundary conditions 

and detailed resolution of the solid—fluid juncture flow for the 

Sw/fp for laminar flow and unsteady large-domain results for 

realistic geometries. An average job run took 2 Cray-2 hours and 

13 megawords of memory for interactive calculations, and 

6 Cray-2 hours and 20 megawords of memory for large-domain 

calculations. 

Significance 

This study stresses the fundamental importance in the fluid 

mechanics of nonlinear free-surface problems and the physics of 

the solid—fluid juncture boundary layer and wake with waves 

and is of strategic importance in the identification and reduction 

of surface wake signatures.

Future Plans 

We will continue to work on Reynolds-averaged Navier—Stokes 

calculations for the Sw/fp geometry using a two-layer turbulence 

model. Also, large-domain calculations and a large-eddy 

simulation for the Sw/fp geometry will be done. 

Publications 

1. Stern, F.; Choi, J. E.; and Hwang, W. S. "Effects of Waves on 

the Wake of a Surface-Piercing Flat Plate: Experiment and 

Theory." Accepted by J. Ship Research. 

2. Tahara, Y.; Stern, F.; and Rosen, B. "An Interactive Approach 

for Calculating Ship Boundary Layers and Wakes for NonZero Froude Numbers." J. Computational Physics 98, no. 1 

(1992): 33-53. 

Wave-induced separation saddle on the plate and projected onto 

the mean free-surface flow pattern.

187
Space Station Freedom Internal Flow Analysis 

Eric T. Stewart, Principal Investigator 

Co-Investigator: Lee A. Kania 

NASA Marshall Space Flight Center/Sverdrup Technology, Inc. 

Research Objective 

To facilitate the design of the air distribution system (ADS) within 

Space Station Freedom (SSF). By performing computational fluid 

dynamics (CFD) analysis of the SSF ventilation flows, we hope to 

assess the ability of ADS to satisfy its performance requirements 

and to expose any additional problems. The ADS is a subset of 

the environmental control and life support system (ECLSS), from 

which the pertinent requirements arise. An average velocity 

between 15-40 feet per minute must be maintained in all 

elements (modules and nodes) and the flow must be well mixed 

and circulated throughout 5SF. 

Approach 

First, we perform CFD analyses of each SSF element and then 

couple the models to simulate SSF. A CFD code developed at 

NASA Ames, INS3D, is used for this study along with a PrandtlVan Driest algebraic turbulence model. 

Accomplishment Description 

A CFD model of a generic module has been developed and a 

solution was obtained. The grid for a node was generated, and 

the resulting CFD model development is in progress. The model 

contains six inlets and six outlets. The internal flow volume 

is approximately 1,050 cubic feet and the grid contains 

1.88 million grid points. The flow Reynolds number is 10,275 

and is based on conditions at the diffuser face. The solution 

required 8,000 iterations to converge using 121 Cray-2 hours 

and 32 megawords of memory. 

Significance 

Current solutions indicate that the ADS produces an average 

velocity of 30.2 feet per minute for the baseline case within a 

generic module. This easily meets the design requirements. In 

addition, short-circuiting does not appear to be a problem. 

However, visual inspection indicates that the flow does not 

appear to be well mixed for the baseline case and may necessitate redirection of the diffuser vanes. This flow solution enables 

the investigators to further examine various physical phenomena 

of interest to the ECLSS designers.

Future Plans 

We plan to complete the node model, then couple the node and 

module to form a CFD model of SSF. In addition, the computed 

flow fields will be used for contaminant tracking calculations. All 

CFD results will be validated through comparison to experimental data. 

Velocity magnitude contours for various two-dimensional planes 

within a Space Station Freedom module (magnitudes range from 

0-440 feet per minute corresponding to colors from blue to 

magenta). 

188
Parallel-Vector "Out-of-Core" Equation Solver 

Olaf 0. Storaasli, Principal Investigator 

NASA Langley Research Center 

Research Objective 

To perform structural analysis on complex models requires 

simultaneously solving millions of equations, which requires 

significant memory, speed, and large and fast data transfer from 

peripheral storage. 

Approach 

Proposed flight vehicles such as the High-Speed Civil Transport, 

the National Aero-Space Plane, and the Advanced Tactical 

Fighter require comprehensive structural analysis that can only 

be performed in sufficient detail on high-performance parallel 

supercomputers. Parallel methods to perform structural analysis 

on high-performance computers are required. The Cray Y-MP 

and gamma super-computers were explored to develop and test 

a method to solve large systems of equations. 

Accomplishment Description 

An "out-of-core" version of the PVSOLVE equation solver was 

developed and tested on large-scale structural analysis applications. Two million equations were simultaneously solved in 

139 seconds on one Cray Y-MP processor using the solid state 

disk (SSD). The algorithm was written to exploit parallelism, 

which further reduced the time by the number of processors 

used (eight on the current Cray Y-MP and 16 on the Cray C-90, 

where each processor is twice as fast). 

Significance 

Until now, large structural analysis problems could only be 

attempted if they either fit in the restricted memory size of 

supercomputers or took extensive time using conventional 

structural analysis techniques. The new equation solver overcomes size restrictions by significantly reducing solution time, 

virtually eliminating data transfer time, which dominated 

previous out-of-core techniques, and eliminating large memory 

requirements by using the out-of-core solution in concert with 

the SSD.

Future Plans 

The out-of-core equation solver is being evaluated on the 

Cray C-90. Results will be compared with the results for the 

same problems solved on the Intel Delta supercomputer. 

Publications 

1. Storaasli, 0. et al. "A Parallel-Vector Algorithm for Rapid 

Structural Analysis on High-Performance Computers." NASA 

TM-1 02614, April 1990. 

2. Nguyen, D. et al. "Parallel-Vector 'Out-of-Core' Analysis of 

Large-Scale Structures on Supercomputers." To be presented 

at the 2nd Symposium on Parallel Computational Methods 

for Large-Scale Structural Analysis and Design, Norfolk, VA, 

Feb. 1993. 

1 P -° Structural Analysis Model of a 

t

High-Speed Civil Transport 

'dim

Displacements 

00 Ir so I 97,107,762 t.rms 0 I 

68.60 solutIon S0cs 

U.g SC..y V.MP 

SO I o.. . N, 

Parallel-vector "out-of-core" solver reduces structural analysis 

time on supercomputers.

189
Harrier YAV-8B Wing Aerodynamics 

Michael W. Stortz, Principal Investigator 

Co-Investigators: Lie-Mine Gea and Gil W. Chyu 

NASA Ames Research Center 

Research Objective 

To define an aerodynamic modification to the wing—pylon 

geometry to improve the performance and combat survivability 

of the YAV-813. The long-term objective is to gain a better 

understanding of the aerodynamic interaction between wing and 

pylons for future designs, demonstrate the capability of improving flow quality by using computational fluid dynamics (CFD) 

methodology, and to provide a design tool to optimize the 

aircraft's performance. 

Approach 

The three-dimensional thin-layer Navier—Stokes code (F3D) with 

the Chimera overset grid scheme was used to calculate the 

transonic flow field around the Harrier YAV-813 wing with two 

pylons and an outrigger wheel housing. Separate grid systems 

are generated for the wing, each pylon (including fairing), and 

the outrigger wheel housing. Design iterations of the pylon and 

fairing combination were easily accommodated. 

Accomplishment Description 

An efficient CFD procedure based on the F3D/Chimera code 

was developed to compute transonic flow over the wing—pylon 

geometry. Previous calculations had demonstrated good 

agreement with flight-test data and had shown that shock waves 

were induced by the pylon-fairing geometry. A rather large 

region of flow separation outboard from the pylon was identified. A new pylon-fairing shape with a concave contour and 

smoother curvatures near the wing—pylon junction has been 

proposed to improve the flow quality. The computed results 

show that the strength of the local shock induced by the new 

fairing is weakened on both sides of the pylon and the size of 

separation region on the outboard side of the pylon is significantly reduced. Calculations required about 20 Cray-2 hours 

and 16 megawords of memory per case. 

Significance 

Since pylons are required for external-store carriage, their 

adverse affect on aerodynamic performance should be minimized. This work demonstrates the use of CFD methodology in 

addressing the problem and providing solutions to current 

problems.

Future Plans 

The pylon fairings have been redesigned and the calculations 

predict significantly improved flow quality. The next step is to 

build and flight test the new fairings, closing the loop in the 

computation-to-flight philosophy. 

Publication 

Gea, L.-M.; Chyu, W. J.; Stortz, M. W.; and Chow, C.-Y. "Flight 

Test and Numerical Simulation of Transonic Flow Around 

Harrier YAV-813 Wing." Presented at the AIAA 22nd Fluid 

Dynamics, Plasmadynamics, and Laser Conference, Honolulu, 

HI, June 1991. 

Comparison of Mach contours for pylon (airings on the outboard 

pylon (viewed from above; outboard at top; flow from left to 

right). 

190
Viscous Reacting-Flow Applications to Scramjet Propulsion 

Sundares V. Subramanian, Principal Investigator 

Co-Investigator: Richard J. Gaeta, Jr. 

General Electric Aircraft Engines 

Research Objective 

To optimize and apply state of the art three-dimensional 

computer codes for design and performance improvements of 

critical hypersonic vehicle components such as inlets, supersonic combustors, and nozzles. 

Approach 

We solved the three-dimensional time-dependent compressible 

Navier—Stokes equations with an appropriate turbulence model 

and physically realistic chemical kinetics model to describe 

combustion phenomena occurring in the component of interest. 

Accomplishment Description 

The computer program RPLUSGE was enhanced by incorporating an algebraic turbulence model, boundary conditions to 

accommodate cooled walls, and sets of new rate constants for 

hydrogen/air combustion reactions. Numerically intensive 

computations corresponding to Mach 18 and Mach 12 flight 

conditions were performed to investigate the scramjet-combustor 

performance characteristics. Using internally developed postprocessing programs, key performance parameters such as fuel 

penetration, mixing, and heat release were determined for a 

slanted-step fuel-injection geometry through high-aspect-ratio 

rectangular orifices. The accompanying figure illustrates a

product of the post-processing capabilities developed. Net heat 

release and exothermic and endothermic heat release are shown 

at a distance of /D* = 25 from the step. This combustor 

configuration includes a slanted step with "slit-slot" fuel injectors 

staggered on the the top and bottom wall. The inlet conditions 

simulate a flight velocity of Mo = 12 while the fuel equivalence 

ratio is 2.4. Using 81 x 61 x 41 grid points and the nine-species 

18-step finite-rate chemistry flow, convergence was obtained 

after 2,500 iterations that required 30 single processor Cray Y-MP 

hours and 22 megawords of memory. 

Significance 

Enhanced flow prediction methods and codes are beneficial to 

all air-breathing hypersonic propulsion systems and are essential 

in the design and development of supersonic combustors that are 

characterized by a high degree of mixing, reaction, and combustion efficiency. 

Future Plans 

Research is under way to determine the relative performance 

improvements of different combustor/fuel injector designs, with 

particular emphasis on quantifying important flow mechanisms 

that are crucial in the understanding of scramjet operation. 

GEAE S4 SCRAMJET COMBUSTOR

b, Exothermic Heat Release 

____

I

-1000e 

1 Fl ii-I2O. 1 

Endothermic Heat Absorption Net Heat Release 

- 2 - 2A 2 P - 

' - o,o 

ii ii rAl 

Illustration of the post-processing capabilities developed.

191
Microwave Hyperthermia Computer Modeling 

Dennis Sullivan, Principal Investigator 

Stanford University School of Medicine 

Research Objective 

To develop better equipment and methods for deep regional 

microwave hyperthermia, a cancer therapy that uses microwave 

heat. Computer simulation is used to design new microwave 

applicators and to simulate treatments on existing equipment to 

increase their effectiveness. A method has been developed to 

simulate treatments of patients in an annular phased array (APA), 

the most commonly used device in the treatment of deep-seated 

cancer tumors. 

Approach 

A three-dimensional frequency-dependent finite-difference timedomain method has been developed which accurately determines the electromagnetic field of microwave applicators and 

the body being heated for multiple frequencies. This method 

solves the time-dependent Maxwell's equations in difference 

equation form. 

Accomplishment Description 

A method for planning patient treatments in the APA has been 

developed and put into clinical use. The APA is a heating device 

with four independent sources that can be energized in the

60-120 MHz frequency range. Each source can be driven with a 

different amplitude and phase, allowing a wide range of input 

parameters. Therefore, it is desirable to use computer simulation 

to develop the optimum treatment plan. By using a 60-slice CT 

scan of the patient, each treatment plan is based on a true threedimensional, patient-specific model. The program to model the 

applicator and patient requires about 10 megawords of memory 

and 200 Cray Y-MP seconds. The main research effort, however, 

has switched from treatment planning on the APA to design and 

development of more effective applicators for deep regional 

hyperthermia. Methods have been developed that simulate 

illumination with pulses instead of single-frequency illumination. 

Significance 

The computer simulation has been developed to enhance the 

effectiveness of hyperthermia treatments with existing equipment 

and to expedite the development of more effective applicators. 

Future Plans 

Research continues in developing optimization methods for the 

APA treatment planning programs, and using current software to 

design more effective applicators. 

Simulated microwave energy deposition pattern from the simulation of patient treatment in the annular phased 

array. Contour lines represent 70% of normal, ranging from 70% (blue) to 90% (yellow). 

192
40.00 

27.50 

15.00 

2.50 

-10.00 

GRID: 96X4 

Cray Y-Mp: 5 

Re - 15 Millio 

Computational Fluid Dynamics for Naval Applications 

Chao-Ho Sung, Principal Investigator 

Co-Investigators: Michael J. Griffin, Chi-Wang Shu, and George T. Yeh 

David Taylor Research Center 

Research Objective 

To develop the capability to predict and understand the physics 

of the flow field about naval surface ships and submarines with 

special emphasis on complex vortical flows. This will aid in the 

innovative design of naval surface ships and submarines. 

Approach 

The three-dimensional incompressible Reynolds-averaged 

Navier—Stokes equations, supplemented by appropriate turbulence models, are solved. Presently, a scheme based on a 

central-difference finite-volume spatial discretization and an 

explicit one-step multistage Runge—Kutta time-stepping has been 

developed. Convergence acceleration techniques include local 

time stepping, implicit residual smoothing, and multigridding. 

Accomplishment Description 

The computer code IFLOW was developed for practical design 

applications. A run on a 96 x 32 x 56 grid requires 9.5 megawords 

of memory and less than 10 Cray Y-MP minutes to reduce the 

residuals by three to four orders of magnitude. A production run 

with 1.8 million grid cells requires about 89 megawords of 

memory and 2 Cray Y-MP hours. The quick return makes 

practical design iterations feasible and makes easier the task of 

tuning turbulence modeling. Applications to increase the

propulsor efficiency of a surface ship and reduce the noise level 

of a submarine have been made. 

Significance 

Since vortical flows are the major source of noise and the 

obstacle to increased propulsor efficiency, it is crucial to 

develop the predictive capability of vortical flows in order to 

design control devices or alter configurations to reduce the 

production of vortical flows. Significant progress has been made 

in this area in the past few years. 

Future Plans 

A multiblock structure allowing abrupt changes in grid spacings 

at interfaces will be completed and used to implement the local 

grid refinement and increase the ability to resolve the vortical 

flow. An effort will be made to validate the flow field computation about a fully appended submarine. 

Publication 

Sung, C.-H.; Griffin, M. J.; Smith, W. E.; and Huang, T. T. 

"Computation of Viscous Ship Stern and Wake Flow." Second 

Osaka International Colloquium on Viscous Fluid Dynamics in 

Ship and Ocean Technology, Osaka, Japan, Sept. 1991. 

IFLOW: Suboff Body With Sail, Alpha - 10 Deg 

A 

. I

Streamwise Vorticity 

Or\

Submarine body with sail simulated using the IFLOW code.

193
Simulation of Scramjet Flow Fields 

R. C. Swanson, Principal Investigator 

Co-Investigator: E. Turkel 

NASA Langley Research Center 

Research Objective 

To develop accurate, efficient, and reliable algorithms for 

solving the three-dimensional Navier—Stokes equations for highspeed internal flows. 

Approach 

In this research we use a cell-centered finite-volume spatial 

discretization, a numerical dissipation model, and a multistage 

time-stepping scheme with acceleration techniques for steadystate calculations. 

Accomplishment Description 

A central-difference scheme with standard multigrid techniques 

can be used to efficiently obtain high-speed flow solutions. In 

this project, several modifications to the scheme have resulted in 

a significant improvement in its robustness. A new initialization 

procedure with flow positivity constraints was implemented and 

is extremely beneficial in computing flows with strong shock and 

expansion waves since it prevents the pressure and density from 

becoming negative. A damped multigrid restriction operator was 

applied to control nonphysical upstream influence in the vicinity 

of a shock wave during the multigrid process. As part of the 

evaluation of the present multigrid algorithm for high-speed 

flows, we considered Mach 3 flow in a three-dimensional 

symmetric corner formed by the intersection of two wedges 

with angles of 9.48 degrees. A computation was performed for 

laminar flow with a Reynolds number of 3.07 million per meter, 

a wall temperature of 294 K, and a free-stream temperature of 

105 K. This flow simulates the flow in a scramjet inlet. The 

associated complex flow field consists of two wedge shocks 

joined by an oblique shock, embedded internal shocks starting at 

the intersection points and terminating at the wedges, a triangular region defined by the corner shock, and two slip lines. The 

mesh for the calculation consisted of 32 cells in the streamwise 

direction and 64 x 64 cells in each cross-flow plane. In the 

accompanying figure, the density contours for this case are 

displayed and the basic structure of the flow is delineated. There 

is reasonably good agreement between the predicted surface 

pressures and experimental data.

Significance 

This project has provided a capability to compute high-speed 

flows efficiently with a multigrid method and central spatial 

differencing. Such an advancement is important with the current 

strong interest in high-speed flight vehicles. 

Future Plans 

We will demonstrate that the present three-dimensional flow 

solver can be used effectively for computing high-speed internal 

flows with chemistry. 

Density contours for Mach 3 flow in a symmetric corner. 

194
Low-Speed Aircraft Maneuvering Aerodynamics 

Tsze C. Tai, Principal Investigator 

Naval Surface Warfare Center 

Research Objective 

To investigate the deck-landing aerodynamics in an aircraft—ship 

interface environment and develop a capability for analyzing the 

effect of airwake, cross wind, wind-speed fluctuation, and gust 

on the lift characteristics of an aircraft during a landing maneuver. 

Approach 

The three-dimensional thin-layer Reynolds stress-averaged 

Navier—Stokes equations are solved subject to a low-speed, 

nonuniform free stream due to the airwake and/or cross wind. A 

zonal approach is adopted where the flow field is subdivided 

into smaller zones. This approach allows suitable grids for 

different zones to be generated with only a moderate memory 

requirement. The NASA Ames multi-zone ARC3D/CNS code and 

the NASA Langley multi-zone CFL3D code will be used as the 

basic programs. Steady flows with nonuniform free stream due to 

the airwake from the ship superstructure and/or cross wind will 

be considered. The ship airwake will be modeled separately and 

its flow field will be used as input to the computation for aircraft 

aerodynamics. 

Accomplishment Description 

A converged Navier—Stokes flow solution for a complete F-i 4A 

aircraft configuration has been obtained. The solution is based 

on a new grid consisting of six zones that significantly reduce 

the complexity of multi-zone computation. Other advantages 

include effectively modeling the fuselage juncture and the wing 

tip and easily extending to whole flow-field calculations subject 

to nonuniform free stream from the ship's airwake and/or cross 

wind. The particle trace of the flow field is shown in the 

accompanying figure. The effect of nonuniform free stream 

is evaluated by imposing an analytical defective function 

representing the nonuniform free stream and generating a 

more realistic airwake using a separate three-dimensional 

Navier—Stokes flow solver. The second approach, the ship 

airwake, is calculated by using the CFL3D code for two simple 

block structures on the ship. The more complex ship superstructure is then added to represent an aircraft—ship interface environment. A typical job run takes about 2-5 Cray-2 hours and 

12-15 megawords of memory. 

Significance 

The Navy has initiated a new research and development task to 

simulate analytical/computational technologies to improve the 

aircraft—ship interface capability. The simulation involves 

conditions of the ship airwake, ship motion, and interaction 

with aircraft or helicopters. 

Future Plans 

Calculations of the aircraft—ship interface aerodynamics with real 

ship airwake subject to cross wind have been planned. The 

fixed-wing aerodynamics technology developed in this work is 

being extended to the Navy's V-22 Osprey rotorcraft operation.

Publications 

1. Tai, T. C. and Taylor, B. "Grid Generation about an F-i 4A 

Aircraft Configuration for Navier—Stokes Flow Calculation." 

Presented at the Third International Conference on Numerical Grid Generation in Computational Fluid Dynamics and 

Related Fields, Barcelona, Spain, June 1991. 

2. Tai, T. C. and Walker, M. "On Three-Dimensional Flow 

Separation Criteria." AIAA Paper 91-1740, presented at AIAA 

22nd Fluid and Plasma Dynamics and Lasers Conference, 

Honolulu, HI, June 1991. 

3. Tai, T. C. "Flow Separation Patterns Over an F-i 4A Aircraft 

Wing." Journal of Aircraft 28, no. 12 (Dec. 1991): 818-827. 

E!. 

Navier—Stokes flow over an F- I 4A aircraft. 

- '1 •t-2.j lTit, ,jfl - - - - 

-.-

Navier—Stokes flow for the airwake of a ship.

195
Elements Nodes 

Mesh 1 100215 19048 

Mesh 2 212433 40322 

Mesh 3 863967 161608 

(a) Mesh 1 (c) Mesh 3 

(b) Mesh 2 (d) Pressure contours 

Unstructured Multigrid Euler Solver 

Rajiv Thareja, Principal Investigator 

Co-Investigators: J. Peraire, J. Peiro, and K. Morgan 

Lockheed Engineering and Sciences Company/Imperial College/University College, Wales 

Research Objective 

To develop a low-storage, computationally efficient algorithm for 

the solution of the compressible Euler equations on unstructured 

tetrahedral meshes. 

Approach 

An unstructured-mesh generator is used to create a sequence of 

refined meshes about complex aircraft shapes. Steady-state flow 

conditions are computed by advancing the transient form of the 

Euler equations using a multistage time-stepping scheme. A sidebased data structure is employed for the three-dimensional 

mesh, which reduces the memory demands. For a sequence of 

successively coarser nested meshes, multigrid solution algorithms accelerate convergence. The meshes are unstructured and 

not nested, thus the nodes on one grid do not coincide with the 

nodes on the next coarser or the next finer grid. The multigrid 

concept has been extended to these meshes. 

Accomplishment Description 

The multigrid scheme has been developed, tested, and applied 

to several complex realistic aircraft shapes. Appropriate intergrid 

transfer procedures between such unrelated meshes have been 

developed and different cycling strategies between meshes have 

been implemented. A higher-order form of the diffusion operator 

has also been incorporated. The computational performance of

the flow solver is dramatically improved in terms of lower 

memory requirements and improved rates of convergence via 

the multigrid scheme. The accompanying figures show three 

progressively refined surface meshes for Mach 0.85 flow at 

2 degrees angle of attack over a twin-engine Dassault Falcon. 

The finest mesh has 863,967 elements and 161,608 nodes. A 

three-stage time-stepping scheme and a double-W cycle was 

employed with one pre- and one post-smoothing step on each 

mesh. The computed steady-state pressure contours on an aircraft 

surface are shown in the last figure. Computations took about 

3 Cray Y-MP hours. The multigrid code requires 89 storage 

locations per node, which compares quite favorably to 63 per 

node for the flow solver with a single grid. 

Significance 

The scheme was used to solve flows about realistic complex 

aircraft bodies in the transonic-flow regime. A converged 

solution is obtained on large meshes in 2-3 Cray Y-MP hours. 

Future Plans 

We will extend the scheme to solve the Navier—Stokes equations, including the automatic generation of semi-structured 

meshes. 

(a)—(c) Three progressively refined surface meshes. (d) Computed steady-state pressure 

contours on an aircraft surface. 

196
Multi-Dimensional Simulations of Noctilucent Cloud Formation 

Gary E. Thomas, Principal Investigator 

Co-Investigator: Eric J. Jensen 

University of Colorado, Boulder/NASA Ames Research Center 

Research Objective 

Perform two-dimensional numerical simulations of noctilucent 

clouds (NLC) as they are affected by atmospheric gravity and 

tidal-wave perturbations. 

Approach 

The NASA Ames aerosol model is applied to the microphysical 

evolution of ice and dust particles in the mesosphere and a 

comprehensive set of physical and dynamical processes believed 

to be important for cloud evolution is included. Physical and 

optical properties are simulated for comparison to ground-based 

and satellite NLC data. 

Accomplishment Description 

The aerosol model was modified to accommodate the particular 

environment and specific properties of NLC particles. The first 

set of wave perturbation calculations designed to assess the gross 

properties of the wave—cloud effects is nearly finished. A typical 

job run requires 0.5 Cray-2 hours and 8 megawords of memory. 

Significance 

The first two-dimensional simulations of NLC were performed. 

It is a long-standing problem to explain the highly structured 

appearance of NLC in terms of wave action. It is unknown if the 

waves merely perturb the altitude of the cloud layer due to uplift 

and downlift in the wave crests and troughs. The adiabatic 

temperature changes and the divergence/convergence of winds 

may affect the microphysics, and therefore the physical and 

optical properties of the cloud particles. We have shown that the 

time-scale ratio of the wave compared with the time scale of ice 

formation may answer these unknowns. When this ratio is less 

than unity, the wave modification is primarily dynamic in 

nature; when it is somewhat greater than unity, microphysical 

processes become significant. Waves have a significant effect on 

eroding the overall cloud brightness. NLC are unstable in time 

periods greater than one day unless the air temperature is lower 

than expected from simple saturation requirements. 

Future Plans 

Wave erosion effects will be characterized in terms of a decay 

time. This will be useful in modeling studies of the cloud occurrence and brightness. The effects of the mean circulation on the 

occurrence of NLC and on its variation with latitude will be 

studied. The two-dimensional simulation will be applied to a 

more realistic situation of tidal/gravity-wave interactions using 

the mechanistic model of D. Fritts from the University of Colorado.

Publication 

Jensen, E. "Numerical Simulations of Wave Patterns in Noctilucent Clouds." Symposium on Noctilucent and Mesospheric 

Clouds, Middle Atmosphere Science Symposium, International 

Union of Geodesy and Geophysics Meeting, Vienna, Austria, 

Aug. 1991. 

Numerical simulation of wave-disturbed noctilucent cloud 

brightness; wavelength = 100 km, period = 1 hour. Brightness 

(albedo) shown in color; red = high, blue = low. Temperature 

contours (white lines) are in K.

197
40 

E 

C, -a 30 

20 

Stratospheric Ozone Destruction 

Owen B. Toon, Principal Investigator 

NASA Ames Research Center 

Research Objective 

To develop the capability to predict the particles in the stratosphere so that heterogeneous chemical reactions involving these 

particles can be investigated. 

Approach 

A microphysical cloud model, which treats heterogeneous 

nucleation, condensation of water and nitric acid vapors, 

evaporation, coagulation, and particle transport, has been used 

to simulate cloud properties, and a module to treat homogeneous nucleation has been developed. The model has been 

applied to investigate new particle formation in the polar regions 

and in the El Chichon and Pinatubo volcanic clouds. 

Accomplishment Description 

Observations show that new sulfate particles are forming in the 

polar winter stratosphere at altitudes above about 25 km. The 

origin of these particles is unknown and some researchers have 

implied that novel chemistry may be occurring. We have 

numerically investigated this process and found that the low 

temperatures in this region of the stratosphere lead to the 

nucleation of any sulfuric acid vapor that may be present. We 

are investigating whether the sulfuric acid vapor is the result of 

photochemical formation of new sulfuric Pinatubo volcanic 

clouds. Our goal is to understand how these particles evolve in 

space and time and the degree to which they may provide 

surfaces for heterogeneous reactions. 

Significance 

Ozone destruction has been observed at both poles and in the 

midlatitude stratosphere. It is clear that these losses of ozone are 

triggered by reactions that occur on the surfaces of particles in 

the stratosphere. In order to predict future ozone losses, we must 

have a predictive capability for stratospheric particles and we 

also need to determine the impact that supersonic aircraft might 

have on these particles. Aircraft contribute particles by injecting 

exhaust particles, by adding sulfate to the stratosphere, and by 

increasing the abundance of nitrogen oxides and water vapor. 

Successful simulations of stratospheric particles and comparisons 

with observations will contribute to the development of this 

predictive capability.

Future Plans 

The work on stratospheric ozone has been stopped. Work under 

this task has been redirected to tropospheric chemistry and a 

limited number of ozone simulations are now being done. 

Publications 

1. loon, 0. B. and Turco, R. P. "Polar Stratospheric Clouds and 

Ozone Depletion." Scientific American 264 (1991): 68-75. 

2. Hamill, P. and Toon, 0. B. "The Physics of Polar Stratospheric Clouds." Physics Today 14 (1991): 34-42. 

50

Day

80 

230 

240 

280 

310 

-

---- 

10 t I I • .';i 4.••.•.•. I 

1.0E-1 1.OE+0 1.0E+1 1.OE+2 1.0E+3 

Aerosol concentration (r> 0.01 urn) 

The concentration of aerosols in the polar stratosphere is 

computed as a function of time and altitude. Micrometeorites are 

present at high altitude and serve as nuclei for sulfuric acid 

condensation. Homogeneous nucleation occurs and the result is 

the formation of a new layer of particles at an altitude between 

20-30 km, just in and above the ozone hole. 

198
Radar Cross-Section Studies 

Ban H. Iran, Principal Investigator 

Rockwell International, North American Aircraft Division 

Research Objective 

To develop and verify a range of the most accurate and efficient 

radar signature prediction and optimization tools in advanced 

aircraft design. 

Approach 

Three codes were used to evaluate and optimize the electromagnetic scattering from structured airfoils, inlets, and flat plates 

coated and partially coated with layers of radar absorbing 

materials. They were the adaptive integral method code (AIM), 

the electromagnetic surface patch code (ESP), and the body of 

revolution methods of moments code (B0R). 

Accomplishment Description 

A study of a two-dimensional leading-edge configuration was 

conducted using the AIM code. An extensive study of long 

cylinders, cone cylinders, and missile-type configurations, both 

perfectly conducting and coated, was performed using the BoR 

code. The BoR-code study assessed the importance of and the 

means of controlling surface wave diffraction. A thorough study 

of flat-plate configurations was conducted and radar signatures 

of perfectly conducting plates, coated plates, and plates terminated with resistive sheets were computed using the AIM and 

ESP codes. When available, comparisons with measured data 

were made. Particular attention was focused on achieving high 

accuracy in the most difficult numerical case of scattering at 

grazing incidence. 

Significance 

The numerical study of radar cross-section performance of 

various test bodies will have significant impact on the cost and 

duration of current and future aircraft design cycles.

Future Plans 

A significant improvement of the aircraft design cycle can be 

achieved by the development and implementation of radarsignature optimization techniques. The radar-signature prediction codes will be interfaced with an optimizer, and multiparameter optimization runs yielding configurations with the 

most desirable material properties will be performed. 

Publication 

Iran, H. B. "Modern Radar Cross-Section Computation Techniques." IEEE AP-S International Symposium. Chicago, IL, 1992. 

20 r3.6 10.9 inch 

Inch 

(I) 

6 U 15.2 Inch I 0 deg 

—20 

—30 

U

—40 

—50 

—179.00 —119.30 —59.67 0 59.67 119.30 179.00 

Azimuth (deg) 

Back scattering from a perfectly conducting trapezodial plate 

illuminated by a transverse magnetic polarized grazing wave, 

f= 3.5 GHz.

S ESP code 

AIM code 

Experiment 

199
Turbulent Premixed Combustion 

Arnaud Trouvé, Principal Investigator 

Co-Investigator: Thierry Poinsot 

NASA Ames Research Center/Stanford University 

Research Objective 

To use direct numerical simulation (DNS) to increase the basic 

understanding of flame—turbulence interactions with primary 

interest in Reynolds-averaged models for turbulent combustion. 

Approach 

DNS applied to turbulent combustion solves the full Navier—Stokes 

equations and uses a reduced chemical scheme to model the 

combustion processes. In this study, the chemical model is a 

single-step irreversible finite-rate chemical reaction. The numerical 

code is a three-dimensional high-order finite-difference code. 

The code can treat non-period domains, and accounts for 

compressibility effects and thermal expansion of the reacting 

flow due to heat release. 

Accomplishment Description 

We completed (1) a flame—vortex interaction problem used to 

construct turbulent diagrams to predict the occurrence of flame 

quenching and to propose a new definition of the flamelet regime, 

(2) a model problem of curved flame elements to investigate the 

influence of curvature on the structure of the reaction zone, and 

(3) a premixed flame problem with a non-unity Lewis number 

embedded in isotropic turbulence to assess the effects of strain 

rate and flame curvature on the structure of the reaction zone. 

Recent simulations include premixed flames in isotropic 

turbulence and linear turbulent-shear flows to assess the effects 

of the turbulence structure on the flame response (see the 

accompanying figure), and flame ignition in isotropic turbulent 

flows to describe and help model the early growth of a flame 

kernel and its transition to turbulent flame propagation. Twodimensional computations required 10 Cray Y-MP hours and 

8 megawords of memory. Three-dimensional computations 

required 80 Cray Y-MP hours and 16 megawords of memory. 

Significance 

Flamelet models constitute the most common approach for 

premixed turbulent combustion. The validity of these models 

was investigated and most practical situations are likely to satisfy 

the basic assumptions used in the flamelet approach. The role of 

flame curvature was clarified and, while curvature determines 

the local characteristics of the reaction zone, global spaceaveraged features do not require curvature effects and are 

captured by predictions based on the stagnation point flame 

picture. 

Future Plans 

We will continue to use a more refined chemical model to 

capture important physical phenomena like chemical-kinetic 

extinction and pollutant formation. We will investigate the 

modeling of the flame surface density under various mean flowfield configurations to assess the effects of mean flow convection 

on the flame wrinkling.

Publications 

1. Poinsot, T.; Veynante, D.; and Candel, S. "Quenching 

Processes and Premixed Turbulent Combustion Diagrams." 

J. Fluid Mech. 228 (1990): 561-605. 

2. Meneveau, C. and Poinsot, T. "Stretching and Quenching of 

Flamelets in Premixed Turbulent Combustion." Combustion 

and Flame 86 (1990): 311-332. 

3. Poinsot, T. and Mungal, G. "A Study of the Laminar Flame 

Tip and Implications for Premixed Turbulent Combustion." 

Combustion Science and Technology 81(1992): 45-73. 

Direct numerical simulation of a premixed flame in threedimensional isotropic turbulence. Contours of the reaction 

rate and velocity field are shown. Fresh reactants are at the back 

of the plot; burnt products are at the front. Grid size = 1,283; 

turbulent Reynolds number = 20; Damkohler number = 1. 

200
0.000 0 025 0.050 0,075 

0.025 

00. 

-0.025 

Multigrid Solution of the Navier—Stokes Equations 

Eli Turkel, Principal Investigator 

ICASE/NASA Langley Research Center 

Research Objective 

To develop an efficient solver for the steady-state compressible 

Navier—Stokes equations at all speeds by utilizing multigrid 

techniques. 

Approach 

A three-dimensional central-difference code was used in 

generalized coordinates. An artificial viscosity were added and 

was constructed as a total-variation-diminishing scheme for a 

scalar one-dimensional equation. The equations are advanced in 

time using a multistage formula. In addition, local time-stepping, 

residual smoothing, and multigrid techniques were used to 

accelerate the convergence to a steady state. A standard FMGFAS multigrid was used for hypersonic flows. The code contains 

standard algebraic turbulence models that are used for comparisons with experiments. 

Accomplishment Description 

The code has been used to solve many problems in subsonic, 

transonic, supersonic, and hypersonic flow regimes. Multigrids 

have been used to accelerate the convergence rates in all the 

flow regimes. In particular, biconic and High-Speed Civil 

Transport configurations have been used and the results of the 

code have compared favorably with experimental results. Even 

for hypersonic flow with viscous and turbulent terms, the 

scheme converges rapidly with a typical residual reduced by 

three to four orders of magnitude within 300 Runge—Kutta

cycles. A medium size mesh is 129 x 97 x 25, and 250 cycles on 

the finest mesh requires about 2 Cray-2 hours and requires about 

20 megawords of memory. 

Significance 

A code has been developed to efficiently solve flow around 

aerodynamic bodies and is presently being used by several 

groups around the country. 

Future Plans 

The code will be generalized to include chemistry and geometric configurations. An incompressible version is also being 

developed. 

Publications 

1. Vatsa, V. N.; Turkel, E.; and Abolhasseni, J. S. "Extension of 

Multigrid Methodology to Supersonic/Hypersonic ThreeDimensional Viscous Flows." Presented at 5th Copper 

Mountain Conference on Multigrid Methods, 1991. 

2. Turkel, E.; Swanson, R. C.; Vatsa, V. N.; and White, J. A. 

"Multigrid for Hypersonic Viscous Two- and ThreeDimensional Flows." AIAA 10th CFD Conference. (1991): 

501-517. 

3. Jorgenson, P. and Turkel, E. "Central Difference TVD and 

TVB Schemes for Time-Dependent and Steady-State Problems." AIAA Paper 92-0053, 1992. 

P I 751.0 

721.0 

691,0 

661 0 I 631.0 

601.0 

5710 

541.0 

511.0 

481.0 

451.0 

421 0 

3910 

361 0 

331.0 

3010 

271 0 

241 0 

211 0 

1t1 0 

11 0 

121.0 

91,00 

61 00 

31 00 

I 00 

Turbulent flow about a blunt cone; Mach = 25.0.

201
Multi-Element High-Lift Concepts 

Walter 0. Valarezo, Principal Investigator 

Co-Investigators: Vincent D. Chin and Dimitri J. Mavriplis 

Douglas Aircraft Company/NASA Langley Research Center/ICASE 

Research Objective 

To calibrate a two-dimensional unstructured-grid Reynoldsaveraged Navier—Stokes method for multi-element high-lift 

configurations. 

Approach 

The method uses multigrid techniques to enhance computational 

efficiency and, due to its unstructured mesh approach, is well 

suited for the complex high-lift geometries of transport aircraft. 

A Baldwin—Lomax turbulence model is utilized. 

Accomplishment Description 

The present method has been applied to a three-element airfoil 

configured for landing. The leading-edge slat and trailing-edge 

flap are both deflected 30 degrees. We compared our predictions to available experimental data obtained at chord Reynolds 

numbers from 5 x 106 to 10 x 106. Studies were conducted to 

assess predicted performance increments for variations in flap 

deflection angle and flap positioning (including gap and 

overhang). Results obtained indicated good general agreement

with experiment, but changes in lift due to gap and overhang 

adjustments were inconsistent with experiment. The code 

successfully predicted separated-flow regions behind the slat and 

in the flapwell (see accompanying figure), but did not predict 

flow separation on the flap suction surface. A total of 175 Cray-2 

hours were used for this calibration. 

Significance 

The development of multi-element high-lift systems, which are 

simpler, but exhibit higher performance than previous generation 

designs, has been identified as a high priority in the transport 

aircraft industry. The timely development of advanced high-lift 

airfoils should be favorably impacted by the application of 

suitable computational methods. 

Future Plans 

Work will focus on the effect of more sophisticated turbulence 

models on the predictive capability of this method. The ability to 

numerically predict maximum lift remains a goal. 

Unstructured grid and Mach number contours for a transport airfoil. Higher than free stream is indicated by the color pink. 

202
Publication 

Atwood, C. A. and Van Dalsem, W. R. "Flow-Field Simulation 

about the SOFIA Airborne Observatory." AIAA Paper 92-0656, 

30th Aerospace Sciences Meeting and Exhibit, Reno, NV, 

Jan. 1992. 

180

E

ls$0 Quiet configuration 100 

CL 

0 

CO 120

3 4 5 

Cavity walls - 

Probe number

7 8 9 10

- Primary & sec.—

mirrors 

-j 

0. 

Co

Z 160 

Noisy configuration 25 

...... -k-- Present 

Experiment 

Simulation of the Unsteady Flow about Transonic Cavities 

W. R. Van Dalsem, Principal Investigator 

Co-Investigators: C. A. Atwood, S. P. Klotz, and D. Sondak 

NASA Ames Research Center/MCAT Institute

Research Objective 

To demonstrate and validate computational fluid dynamics 

capabilities as applied to the Stratospheric Observatory for 

Infrared Astronomy (SOFIA). Numerical evaluation of the fluid 

and optical fields, in addition to experimental results, will 

provide a means to evaluate configurations in a timely and cost 

effective way. 

Approach 

The use of the Pulliam–Chaussee diagonal algorithm within 

a three-dimensional Reynolds-averaged overset mesh 

Navier–Stokes code has shown that resonant frequencies and 

magnitudes have been accurately captured in comparison to 

experiment. These results were found through simulations of 

a steady-state solution of the flow field about the wind tunnel 

configuration, a strongly resonating solution of the cavity, and 

an acoustically quiet solution including the telescope geometry. 

The final simulation in the series required the use of 1.8 million 

cells distributed over 15 grids. The time-varying density field was 

used to determine the optical clarity of the shear layer. Efficient 

computation of this series of configurations was accomplished 

using the steady-state solution as the initial condition for the 

unsteady cases. For this problem the unsteadiness was limited to 

a domain immediately surrounding the cavity. This property was 

used to advantage by iterating only the grids that fell in this 

region. Computation required 16 megawords of memory and 

100 single-processor Cray Y-MP hours with the solid-state 

storage device. 

Accomplishment Description 

Simulations of complex flows over realistic geometries have 

demonstrated the effectiveness of numerical methods for this 

class of cavity problems. Agreement with experiment in frequency, magnitude, and spatial variation of the loads is generally within 5%. Computed optical distortion trends correctly, 

but underpredicts experiment. 

Significance 

The results of these simulations are impacting SOFIA design 

and unsteady telescope loads have been provided to the design 

group. 

Future Plans 

A sizable reduction in aircraft cost would be realized if the 

telescope were mounted aft of the wing. However, the proposed 

aft installation would incur potentially adverse effects on 

empennage loads and seeing capabilities. During the next year, 

we will address these issues.

Summary of 1991-92 NAS Operational Year Stratospheric 

Observatory for Infrared Astronomy results.

203
Powered-Lift Computational Fluid Dynamics Project 

W. R. Van Dalsem, Principal Investigator 

Co-Investigator: K. Chawla 

NASA Ames Research Center/MCAT Institute 

Research Objective 

To simulate a powered-lift configuration in descent to understand the differences in flow physics that result as a variation in 

lift loss and to use this capability to simulate a full configuration. 

Approach 

Using the simplified configuration of a delta wing with thrustreverser jets, we will simulate a powered-lift descent. We will 

solve the three-dimensional unsteady Navier—Stokes equations 

on overset/patched grid systems using OVERFLOW's diagonalized option and use the domain connectivity function to 

communicate solutions across grids and to descend the delta 

wing. 

Accomplishment Description 

Numerical simulations of the flow about a 60 degree delta 

planform wing flying at Mach 0.064, and a Reynolds number of 

1.2 million, equipped with two sonic thrust-reverser jets exiting 

from jet pipes, were performed at a height of about one wingspan above the ground. The converged solution was used to 

begin the descent of the delta wing from this height toward the 

final height of about one-quarter wing-span above the ground. 

Converged solutions were also obtained at heights of one-half 

and one-quarter wing-spans above the ground to compare 

against the dynamic case. 

Significance 

Analysis of solutions from the simulation of a powered-lift delta 

wing in descent will shed light on the mechanisms by which lift 

loss varies for the static and the dynamic cases. The capability to 

simulate accurate lift-loads will allow prediction of minimum 

hover height for a powered-lift aircraft. Also, the solution of this 

simplified configuration will point out the bottlenecks, such as 

slow input/output and lack of methods for visualization and 

analysis of large amounts of unsteady data, that must be resolved 

before full configuration simulations can be routinely performed. 

Future Plans 

We will finish the descent simulation and analyze the results to 

understand the differences in flow physics in the static and the 

dynamic cases. Knowledge gained in this work will be used to 

simulate a full-configuration powered-lift landing.

Publication 

Chawla, K. and Van Dalsem, W. R. "Numerical Simulation of 

STOL Operations using Thrust Reversers." AIAA Paper 92-4254, 

AIAA Aircraft Design Systems Meeting, Hilton Head, SC, 

Aug. 1992. 

SOFIA: Stratospheric Observatory For Infrared Astronomy 

Instantaneous Ma:- --

Re:4x106ft, 

0.85 - 

Configuratio n 25

Cr 

Configuration lOG 0 

A descending delta wing with thrust reverser jets. The ground is 

colored by the pressure coefficient (notice the ripple effect even 

before the jets have reached the ground); particle traces are 

colored by height. 

204
High-Reynolds-Number Viscous Flow over Aircraft Components 

Veer N. Vatsa, Principal Investigator 

Co-Investigator: Eli Turkel 

NASA Langley Research Center 

Research Objective 

To develop an efficient, accurate, and robust numerical procedure for computing high-Reynolds-number viscous flow over 

aircraft components. 

Approach 

A multistage Runge-Kutta time-stepping scheme with multigrid 

acceleration technique is extended to solve supersonic and 

hypersonic flows. A study is being conducted to assess the 

accuracy of predicted heat-transfer loads. 

Accomplishment Description 

A multigrid-based finite-volume numerical scheme developed 

for computing aerodynamic flows over aircraft components in 

the transonic-flow regimes has been extended to solve supersonic and hypersonic flows. The numerical code, TLNS3D, is 

one of the most efficient codes available today for solving 

viscous flows in the transonic-flow regime and is known for 

accurate prediction of aerodynamic loads. An attempt was made 

to calibrate the TLNS3D code for predicting heat transfer loads 

for high Mach-number flows. The flow over a modified shuttle 

orbiter at Mach 6 and 30 degrees angle of attack was used as a 

test case. Approximately 3 Cray Y-MP hours were required to get 

a converged solution for this configuration on a mesh consisting 

of approximately 82,000 nodes. As shown in the accompanying 

figure, the computed heat-transfer-coefficient distributions 

compare well with the experimental data and calculations from 

an upwind-based Navier-Stokes solver. 

Significance 

It appears that the present code can accurately predict the heat 

loads for supersonic (Mach 6) flows, resulting in an order-ofmagnitude reduction in computer time compared to nonmultigrid codes currently in use. 

Future Plans 

The method will be generalized to accommodate blockstructured grids so that flow over complex aircraft configurations 

can be computed.

Publication 

Vatsa, V. N. "Evaluation of a Multigrid-Based Navier-Stokes 

Solver for Aerothermodynamic Computations." AIAA 

Paper 92-4563, 1992. 

(M,,, = 6.0, a = 30 deg, Rel = 4.5 X 105)

2.0 o Experimental data 

- - - - - Navler-Stokes computations, 

upwind scheme (LAURA) 

1.5 Navler-Stokes computations, 

central-difference scheme (TLNS3D) 

X/L 0.3500 

rr 

C%1 E 

Cr

0 .1 .2 .3 .4 .5 .6

s/L 

Comparison of surface heat transfer distributions for a modified 

shuttle orbiter; M, = 6.0, a =30 degrees, Rej = 4.5 x io.

205
(a) Adapted unstructured 

grid for the shock 

Interaction problem

(b) Turbulent kinetic 

energy distribution

(c) Turbulent dissipation

rate distribution 

Two-Equation Turbulence-Model Implementation 

Gururaja R. Vemaganti, Principal Investigator 

Lockheed Engineering and Sciences Company 

Research Objective 

To implement a two-equation turbulence model in an unstructured grid solver and investigate the turbulent shear-layer effects 

in a shock—shock interference pattern developed in a hypersonic 

flow over a blunt body. 

Approach 

Severe aerodynamic heating caused by a shock—shock interference pattern in a hypersonic flow over a blunt body is a critical 

information needed to design an aerospace vehicle. Due to the 

flow complexities, unstructured meshes are preferred for this 

class of problems because of their mesh adaptation capabilities. 

In order to resolve the effects of free-stream turbulence on the 

flow field and the growth of a turbulent shear layer that is 

developed due to type Ill and type IV interactions, a twoequation turbulence model with built-in compressibility effects 

was implemented in the NASA Langley adaptive remeshing code 

Navier—Stokes solver (LARCNESS) algorithm, which is an 

unstructured grid solver. In this algorithm, the six governing 

equations (the four mean flow equations and two turbulence 

equations) are solved simultaneously based on a point-implicit 

upwind method.

Accomplishment Description 

The LARCNESS algorithm with the two-equation turbulence 

model is validated against analytical and numerical data for a 

flow over a flat plate and a free-shear-layer flow. The behavior 

of the turbulence variables near the shocks was predicted to be 

similar to that of other methods. A type III shock—shock interaction problem in a Mach 8.03 flow over a cylinder is currently 

being investigated. To perform three successive remeshings 

and obtain a converged solution on the last mesh takes about 

1 Cray-2 hour. 

Significance 

This research is performed as a part of a government work 

package related to the National Aero-Space Plane (NASP). Once 

the algorithm is validated against experimental data, it is used to 

predict the aerodynamic heating under NASP flight conditions. 

Future Plans 

Future plans include simulation of type III and type IV 

shock—shock interference patterns in comparison to the experiments to be performed at Mach 15 over a cylinder. 

Two-equation turbulence-model implementation in an unstructured grid solver. 

206
LA

Ia 

F-"^ 

Tactical Missile Aero-Propulsion Interaction 

Bill J. Walker, Principal Investigator 

Co-Investigators: C. D. Mikkelsen, K. D. Kennedy, P. F. Booth, and M. E. Vaughn, Jr. 

U.S. Army Missile Command 

Research Objective 

To develop the capability to model missile-plume flow fields, 

including (1) aerodynamic flow separation of the missile 

afterbody boundary layer due to the high pressures at the rocket 

nozzle exit plane; (2) the recirculation, mixing, and chemical 

reaction in the base region; (3) the mixing, shock structure, and 

after burning of the fuel-rich rocket exhaust in the near-plume 

region; and (4) the mixing in the far-plume region. 

Approach 

We are simultaneously solving the fluid-dynamic conservation 

equations, turbulence-model equations, two-phase gas—particle 

interaction equations, and the chemical-kinetics equations using 

proven numerical methods of computational fluid dynamics. 

Accomplishment Description 

PARCH, a three-dimensional Navier—Stokes flow-field solver 

with two-equation turbulence models, finite-rate chemistry, and 

a two-phase flow capability, has been developed and run against 

a set of high-quality, plume flow-field code-validation data. The 

accompanying figure shows the flow-field Mach number for an

axisymmetric, Mach 2.4 air jet into a Mach 1.4 co-flowing 

airstream. Overlaid on the flow field are radial profiles at an 

axial location two radii downstream from the nozzle exit plane 

that compare laser-Doppler-velocimeter measurements for axial 

velocity, radial velocity, and turbulence kinetic energy with 

PARCH Navier—Stokes model calculations and calculations for 

the spatial marching standardized plume flow-field model. This 

case required 20 Cray-2 hours and 32 megawords of memory. 

Significance 

The basic methodology now exists to numerically model and 

predict the complex chemically-reacting and two-phase flow 

fields for missiles with propulsive rocket exhaust plumes. This 

methodology, when fully validated, will provide invaluable 

missile design information that cannot be acquired in wind 

tunnels due to capability, safety, or cost. 

Future Plans 

Efforts will focus on validation of the model with experimental 

plume flow-field data for species and temperature. 

Flow-field Mach number for a Mach 2.4 air jet into a Mach 1.4 co-flowing airstream with comparisons 

of model calculations and laser-Doppler-velocimeter validation measurements.

ØA
Finite-Rate Chemistry Algorithms 

Robert W. Walters, Principal Investigator 

Co-Investigators: Bernard Grossman, Michael Applebaum, Andrew S. Godfrey, William D. McGrory, and David C. Slack 

Virginia Polytechnic Institute and State University 

Research Objective 

To develop and validate simulation tools capable of performing 

complete tip-to-tail analysis including internal and external flowfield calculations. 

Approach 

A three-dimensional implicit upwind Navier—Stokes code with 

generalized chemistry and vibrational nonequilibrium thermodynamics models is used to predict the flow field about hypersonic 

vehicles and in propulsion systems. 

Accomplishment Description 

We have been developing a new version of the General Aerodynamic Simulation Program (GASP) code with new features and 

improvements, particularly in memory and run time for threedimensional elliptic problems. As a separate task, we are 

developing a preconditioning strategy to extend the applicability 

of a compressible Navier—Stokes code to the incompressible 

regime while improving steady-state convergence across the 

Mach number range. We are extending the k-exact reconstruction process to handle discontinuities on unstructured grids. A 

typical calculation takes 5-10 hours and 10-30 megawords of 

memory. We recently submitted GASP 1.3a results for the Third 

JAN NAF Combustion Workshop problem. The results were of 

the turbulent flow over a rearward facing step with two normal 

injectors. On a two-zone grid with a total of 266,090 grid points, 

GASP 1.3a required 4.5 hours and 32 megawords of memory.

Our new version runs this same problem in 50 minutes and 

needs 6.5 megawords of memory. The accompanying figure 

shows particle traces in the recirculation and injection regions 

superimposed on pressure contours in the symmetry plane. 

Significance 

The need to accurately simulate a wide range of three-dimensional 

fluid-dynamic problems with complex physics, including turbulence and finite-rate chemistry models, is required for hypersonic 

vehicle design, propulsion system analysis, and a number of 

strategic defense problems. 

Future Plans 

We are investigating and coupling inverse design techniques 

for supersonic and hypersonic forebody design. In addition, we 

are developing preconditioning techniques for convergence 

acceleration and we are continuing our unstructured flow-solver 

effort. 

Publications 

1. Walters, R. W.; McGrory, W. D.; and Slack, D. C. "GASP 1.3." 

Third JAN NAF Workshop on Scramjet Combustion Modeling, 

Reno, NV, Jan. 1992. 

2. Godfrey, A.G.; Mitchell, C. R.; and Walters, R. W. "Practical 

Aspects of Spatially High Accurate Methods." AIAA 

Paper 92-0054, Jan. 1992. 

Particle traces superimposed on pressure contours in the symmetry plane. 

208
Hydrocarbon Scramjet 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 algorithms code is applied to existing 

combustor-flow test cases to validate the code and provide 

direction for future development of numeric and chemistr 

models. 

Accomplishment Description 

A Rocketdyne dual-mode hydrocarbon scramjet combustor was 

numerically investigated. The fuel—air ratio was increased with 

time from 0-1 by six steps with each one reaching a steadystate condition. The normal and tangential fuel injectors were 

distributed at various combustor stations. The problem was 

modeled as a two-dimensional quasi-steady-state flow with the 

fuel-injector opening adjusted for the correct mass-flow rate. A 

27,046-grid system was used and a global-reaction chemistry 

model was assumed for the oxidation of ethylene, silane, and 

hydrogen fuels used in the combustor. The problem took 

5.27 Cray Y-MP seconds per iteration and the solutions took 

2-4 hours to converge between two different flow conditions 

and required 3 megawords of memory. The predicted combustor 

Mach-number contours and the detailed flow fields adjacent to 

the backward-facing steps are shown in the accompanying 

figures. The Mach numbers for the combustor inlet and exit are 

1 and 2, respectively. The sudden changes of the Mach numbers 

are due to either the injection of the fuel or the change of the 

combustor geometry. The flow in the combustor is quite complicated and the pressure waves generated by the separation 

bubbles behind the top and bottom steps interact with each 

other and create interesting Mach contour patterns. The predicted 

wall static pressure was in good agreement with experimental 

data except for a small region where the three-dimensional flow 

effects were important. The predicted combustor thrusts for 

different fuel—air ratios were also in reasonable agreement with 

experimental data. The agreement between predictions and data 

were excellent for fuel—air ratio below 0.76. For a higher ratio, 

the predictions began to deviate from the data due to the threedimensional flow effects or the limitation of the present globalreaction chemistry model. 

Significance 

The design of an efficient hypersonic combustor is essential to 

the success of an air-breathing hypersonic vehicle. The complexity of the flow physics and the requirement to obtain efficient 

mixing and combustion require an accurate and robust computational tool. 

Future Plans 

This research will be extended to study realistic three-dimensional 

combustor flows with the effect of various hydrocarbon fueloxidation models.

Mach-number con tours. 

Velocity vectors colored with Mach number for the flow behind 

the backward-facing steps.

209
Flow Simulation in Fan and Core Compressor Stages 

Kurt F. Weber, Principal Investigator 

Co-Investigator: Dale W. Thoe 

General Motors Corporation, Allison Gas Turbine Division 

Research Objective 

To validate a Navier—Stokes code that simulates the flow through 

transonic fan and core compressor stages. The analysis must 

operate in a multistage environment and provide a detailed 

description of important phenomena, including the interaction 

of the tip-clearance flow and the blade-passage throughflow. 

Approach 

We applied a fully viscous three-dimensional finite-difference 

code to the solution of compressor and fan flow fields, applied 

domain decomposition, and used multiple grids and multiplegrid types. For a fan stage, the blade-passage flow fields are 

solved on body-conforming 0-grids, and the tip-clearance flow 

field is solved on a fine H-grid embedded in the clearance 

region. H-grids can be placed in the inter-blade-row region and 

upstream from the rotor. The equations are solved using the 

implicit approximate-factorization Beam—Warming algorithm. 

Flow-field data are interpolated between grids using the Chimera 

grid-embedding technique. 

Accomplishment Description 

A three-dimensional embedded-grid Navier—Stokes code, which 

utilizes blade-conforming 0-grids and embedded H-grids, has 

been applied to simulate transonic compressor flows, including 

the coupling of rotor and stator flows and tip-clearance modeling. The implicit solver uses Pulliam's pentadiagonal version of 

the Beam—Warming algorithm. The code accounts for multiple 

embedded grids and works with the output from the multiplegrid data management code PEGASUS. Fully viscous solutions 

with a very fine H-grid embedded in the clearance region have 

been calculated for several rotors, including NASA rotor 67. 

A fully viscous calculation for NASA stage 67 has been completed. The solutions show good agreement with laser-Doppleranemometer measurements. The finest grid solution for NASA 

stage 67 contained 437,999 points in the rotor 0-grid and 

58,890 points in the embedded H-grid. The accompanying 

figure shows a close-up view of a tip-clearance vortex formation 

near the leading edge of NASA rotor 67 at the peak-efficiency 

operating point. The blade surface is colored by pressure 

contours and the particle traces are released in the clearance 

region close to the blade tip. This solution used 21 megawords 

of memory and required 14.2 Cray-2 hours for 2,500 iterations. 

Significance 

The use of three-dimensional 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 significantly reduce the 

amount of testing required. Major performance improvements or 

significant cost reductions are unlikely unless the three-dimensional analyses focus on flow-field resolution and calculation 

accuracy.

Future Plans 

The existing code accounts for multiple-overset or embedded 

grids. This will be used to extend the application to simulate a 

turbofan-engine low-pressure compression-system flow field. 

Publication 

Weber, K. F. and Delaney, R. A. "Viscous Analysis of ThreeDimensional Turbomachinery Flows on Body-Conforming Grids 

using an Implicit Solver." ASME Paper 91 -GT-205, ASME 

International Gas Turbine Exposition, Orlando, FL, June 1991. 

Close-up view of the formation of a tip-clearance vortex. 

210
250 500 750 1000 1250 

z (In.) 

40 

35 

30 

25 

c.'1 

Winged Entry-Vehicle Computations 

Kenneth J. Weilmuenster, Principal Investigator 

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

NASA Langley Research Center 

Research Objective 

To develop a solution for the viscous thermal non-equilibrium 

flow-field computation about winged vehicles 

at hypersonic speeds. 

Approach 

The Halis shuttle model, which combines an accurate description of the lower surface of the shuttle orbiter with a simplified 

Ieeside geometry, was used as the configuration in this study. 

Two viscous volume grids having 145 axial planes, with 

81 points in each cross section and 41 and 81 points between 

the body and out boundary, were constructed. Each volume grid 

was divided into four blocks as shown in the accompanying 

figure. The procedure is to solve each block, independent of the 

other, using the solution on the downstream boundary of the 

previous block as the upstream boundary condition for the 

solution in the adjoining downstream block, thus reducing a 

very large job in terms of memory and time to a series of more 

manageable smaller jobs. The Langley Aerothermodynamics 

Upwind Relaxation Algorithm has been used for the flow-field 

computations. 

Accomplishment Description 

The viscous thermochemical nonequilibrium flow field about a 

winged entry vehicle at hypersonic speeds has been computed 

using a streamwise block by block marching technique. A 

comparison of computed windward centerline heating with data 

from three different flights, shown in the accompanying figure, 

indicates that this procedure yields excellent results. 

Volume grid divided into four blocks.

Significance 

By showing that a large problem such as this can be solved by 

breaking it into parts, the solution for flow over complex 

configurations becomes more realistic. A solution on the 

complete 145 x 81 x 81 grid used in the problem would require 

240 megawords of memory. A configuration with a complex 

leeside geometry might require 2-4 times that much memory 

and makes the blocking technique used here the only reasonable 

option. 

Future Plans 

This blocking technique will be applied to the complete shuttleorbiter geometry to access the vehicle's aerodynamics at hypersonic speeds. Particular attention will be given to identifying the 

flow characteristics which caused the unexpected aerodynamics 

encountered by the vehicle in flight at hypersonic speeds. 

Publication 

Weilmuenster, K. J. and Gnoffo, P. A. "Solutions, Strategies, and 

Heat Transfer Calculations for Three-Dimensional Configurations 

at Hypersonic Speeds." AIAA Paper 92-2921, July 1992. 

Comparison of computed windward centerline heating with data 

from three different flights.

211
Shuttle Leeside Temperature Study 

Kenneth J. Wei Imuenster, Principal Investigator 

Co-Investigator: William L. KIeb 

NASA Langley Research Center 

Research Objective 

To define the computational resources and solution techniques 

required to define the leeside flow over the shuttle orbiter. 

Approach 

Surface temperatures on the leeside of the Orbiter at hypersonic 

flight conditions have been measured by the Shuttle Infrared 

Leeside Temperature Study (SILTS). The SILTS data base has been 

used to access the validity of the computed results. 

Accomplishment Description 

Significant work was required to establish an acceptable volume 

grid for this computation. Software to define surface and grid 

quality and techniques to accelerate the convergence of existing 

elliptic solvers was developed to meet the requirements of this 

project. The Langley Aerothermodynamics Upwind Relaxation 

Algorithm code was used for the solution. The viscous thermochemical nonequilibrium solution incorporated variable wall 

temperature, a catalytic surface, and the dependence of emissivity on temperature. The accompanying figure shows a comparison of the computed leeside surface temperature (top half) and 

the measured surface temperatures (lower half) from STS-32 at 

M,0 = 24 and 40 degrees angle of attack. The SILTS data base is 

limited by the field of view of the infrared camera used in the 

experiment.

Significance 

The figure shows that the computed temperatures are in good 

agreement with those measured in flight. Current computational 

techniques are capable of adequately accessing the temperature 

and heating environment on the leeside of a winged entry 

vehicle at hypersonic speed and high angles of attack. 

Future Plans 

While the comparison with data on the upper surface of the 

wing is good, there are differences between data from thermocouples located on the side and top centerline of the vehicle. 

We will continue to resolve those differences. 

Publication 

Kleb, W. L. and Weilmuenster, K. J. "Characteristics of the 

Shuttle Orbiter Leeside Flow During a Reentry Condition." AIAA 

Paper 92-2951, July 1992. 

Comparison of the computed leeside surface temperature (top halt) and the measured surface 

temperatures (bottom half) from STS-32 at M,0 = 24 and 40 degrees angle of attack. 

212
Tropospheric Aerosols and Clouds 

Douglas L. Westphal, Principal Investigator 

Co-Investigator: Owen B. Toon 

NASA Ames Research Center 

Research Objective 

To create a general numerical model capable of simulating 

complicated interactions between dynamics, aerosol microphys 

ics, clouds, and radiative transfer in the Earth's atmosphere. 

Approach 

A model of atmospheric dynamics, aerosol and cloud microphysics, and radiative heat transfer is evaluated by simulating 

actual atmospheric phenomena and comparing the results with 

observations made during NASA-sponsored field programs. 

Sensitivity tests are done to study the importance of unmeasured 

processes or poorly known quantities. 

Accomplishment Description 

The model was applied to 10 different cases during the First 

International Satellite, Cloud, and Climatology Regional Field 

Experiment in Kansas. The installation of a local ethernet with 

Internet connectivity during the experiment allowed near-realtime simulations of the atmosphere and comparisons with the 

field measurements. Several different cloud parameterizations 

were tested. A typical one-day simulation on a grid of 

77 x 58 x 25 required 3,600 Cray Y-MP seconds and 

10 megawords of memory. 

Significance 

The model is used as a test bed for the different cloud and 

radiation parameterizations. The simulations are compared with 

instantaneous atmospheric measurements and the parameterization weaknesses are identified. The combined model is used to 

study the large-scale aspects of formation, maintenance, and 

dissipation of tropospheric clouds, to interpret aircraft data 

collected during field experiments, and to study the Earth's 

climate. 

Future Plans 

Comparisons between the model simulations and the field data 

will be made, the cloud and radiation parameterizations will be 

improved, and studies with a nested version of the model with 

60 km resolution for the coarse grid and 20 km in the nested 

grid will be initiated. The finer resolution of the model allows 

simulations on scales more consistent with the field and satellite 

observations and should yield a better representation of mesoscale (20-200 km) dynamics.

(c) 

West-to-east cross section of simulated cloud ice (white contours) and snow (shaded contours) at (a) four, (b) five, and 

(C) six hours of model simulation showing the development 

and eastward propagation of bands of precipitation.

213
Supersonic Chemically Reacting Turbulent Flow 

J. A. White, Principal Investigator 

Co-Investigators: J. P. Drummond, J. J. Korte, and M. H. Carpenter 

NASA Langley Research Center 

Research Objective 

To investigate the phenomenon that controls the generation of 

vortical structures in supersonic flow and to utilize that knowledge in optimizing the design of scramjet fuel injection configurations through mixing enhancement. 

Approach 

A three-dimensional parabolized Navier-Stokes (PNS) code that 

employed a near-third-order total-variation-diminishing (TVD) 

scheme was used to analyze flow around a candidate National 

Aero-Space Plane scramjet fuel-injection configuration to 

provide an improved understanding of the loss-generation 

mechanisms governing the vortex-generation phenomenon 

employed to enhance mixing. The PNS algorithm is based on a 

flux-difference-splitting approach and was used because of its 

superior robustness and shock capturing capability. 

Accomplishment Description 

Previously, a computer code was developed that solved the 

parabolized equations governing a multicomponent mixing flow 

utilizing a TVD upwind scheme. The code was validated against 

experiments involving reacting and non-reacting flows and then 

applied to analyze three-dimensional flows in scramjet combustors. Recent experimental research has been directed toward 

improving the efficiency of scramjet combustors through mixingenhancement configurations. A configuration under consideration is the swept-ramp fuel injector. It has been the subject of 

several experimental and numerical investigations. The vorticity 

induced by the swept-ramp configuration can significantly 

improve mixing and combustion in supersonic flows. The desire 

to increase mixing through the generation of streamwise vortices 

must be tempered by the inherent increase in total pressure loss. 

Recent work has been concerned with improving the understanding of the phenomenon governing vorticity and loss 

generation of the swept-ramp configuration. The accompanying 

figure shows the density contours at a symmetry and outflow 

plane with streak-line paths superimposed for a typical configuration. The streak-line paths delineate the vortex core and trace 

the origin of the core. 

Significance 

The numerical simulation provides an improved understanding 

of the complex interactions controlling mixing and loss generation in three-dimensional supersonic flows. This knowledge can 

be used to improve the design of fuel-injection configurations for 

scramjet combustors, thereby improving the overall combustor 

efficiency and contributing to vehicle closure.

Future Plans 

The injector-induced vortical structure will be coupled with the 

primary fuel-injection source interacting with a planar shock in 

order to investigate the interaction of the multiple vortical 

structures. 

Vortex structure and density contours of a scramjet swept-ramp 

fuel injector; red = high density, black = low density. 

214
Shock Interactions in Hypersonic Rarefied Flows 

Richard G. Wilmoth, Principal Investigator 

Co-Investigators: V. K. Dogra and J. N. Moss 

NASA Langley Research Center/ViGYAN, Inc. 

Research Objective 

To provide fundamental information on shock-shock, 

shock-boundary-layer, and shock-shear-layer interactions in 

hypersonic rarefied flows. 

Approach 

The direct simulation Monte Carlo (DSMC) method was used to 

model the shock layer of a sphere (1.6 meters in diameter) 

during Earth entry and the shear layers created by shock-shock 

interactions in hypersonic inlet cowl-lip flows. 

Accomplishment Description 

Calculations of the molecular velocity and energy distributions 

behind the bow shock of the sphere were completed for entry 

conditions at altitudes of 130-90 km. Initial calculations of 

mixing in hypersonic, rarefied shear layers were performed at 

conditions corresponding to those behind an oblique shock 

impinging on a bow shock. A block parallel version of the 

two-dimensional DSMC code was developed and allowed 

multitasking on the Cray-2 and Cray Y-MP through the use of 

Oak Ridge National Laboratory's parallel virtual machine 

message-passing software.

Significance 

The sphere results demonstrate the non-Maxwellian nature of 

the velocity distributions in transitional flow. The shear-flow 

simulations showed the mixing to be self-similar, but with 

considerable density and temperature variations introduced into 

the initially constant density shear layer. The block parallel 

version of the DSMC code will provide the capability to solve 

larger problems by taking advantage of the multiprocessing 

capabilities of the Crays. 

Future Plans 

Further calculations of shear-shear and shock-shear interactions 

at conditions corresponding to National Aero-Space Plane flight 

conditions are planned. Development of the parallel code will 

continue and applications to unsteady flows are planned. 

Velocity Vectors 

25

V25

: 

25 50 75 100 0 25 So 75 - 100 

xr,. VA. 

JDensity Temperature 

25 

0 

25

- 25 - 75 -

25

0 25 50 75 100

Direct simulation Monte Carlo model of hypersonic shear flow.

215
Supersonic Laminar Flow Control Concept 

Chung-Jin Woan, Principal Investigator 

Co-Investigators: Michael W. George and Philip B. Gingrich 

Rockwell International, North American Aircraft Division 

Research Objective 

To develop and demonstrate a computational fluid dynamics 

(CFD) ability in predicting and modeling transition due to 

attachment line, cross flow, and Tollmien—Schlichting instabilities in the supersonic flow regime. 

Approach 

A Navier—Stokes solver is used in conjunction with a threedimensional compressible stability analysis to predict the 

boundary-layer transition location. The methodology has been 

applied to an F-i 6xL aircraft with a wing suction glove. 

Accomplishment Description 

A coarse multiblock Euler grid with approximately 550,000 grid 

points was constructed from the existing Navier—Stokes grid. An 

Euler flow was calculated at Mach 1.6 and 2 degrees angle of 

attack using the USARG3 code and the NASA Langley CFL3D 

code. Excellent agreement in flow properties was noted for the 

two methods. The accompanying figure compares the two 

solutions in terms of pressure contours at the plane of symmetry. 

The computation used 85 Cray-2 hours and 35 megawords of 

memory for an average Navier—Stokes job run using approximately 1 million grid points, or 6 Cray-2 hours for an average 

Euler job run using approximately 550,000 grid points. 

Significance 

A numerical ability to predict transition and the extension of 

laminar flow control to supersonic and hypersonic regimes will 

benefit several national programs, including the High Speed 

Civil Transport and the National Aero-Space Plane. 

Future Plans 

Continued effort will be directed at validating the CFD methodology and the design approach through the revision of the 

numerical configuration model based on the measured F-16xL 

supersonic laminar-flow-control geometry and the expansion of 

computations aligned with the ongoing flight-test program.

Publication 

George, M. W. Status of the F-i 6xL Supersonic Laminar Flow 

Control: Numerical Design Validation." First Annual High-Speed 

Research Workshop, Williamsburg, VA, May 1991. 

Calculated pressures in the plane of symmetry of a modified 

F-l6xL at Mach 1.6 and 2 degrees angle of attack. The pressure 

ranges from 0.5-2.5 of free-stream pressure with blue representing the low value and red the high value. 

216
Nozzle Screech Phenomena 

Chung-)in Woan, Principal Investigator 

Co-Investigator: Michael W. George 

Rockwell International, North American Aircraft Division 

Research Objective 

To develop and validate a computational fluid dynamics 

capability in predicting and modeling nozzle acoustic noise 

generated by high-performance aircraft and to generate a better 

understanding of the physical phenomena of jet aeroacoustics. 

Approach 

A Navier—Stokes code based on a high-order accuracy upwing 

total-variation-diminishing scheme is used to calculate the timeaccurate jet-flow evolution and the predictions are correlated 

with the existing data base. 

Accomplishment Description 

A grid for a NASA experimental research convergent—divergent 

nozzle has been developed and the internal flow has been 

cal-culated and obtained for a laminar flow and two turbulent 

flows with zero- and two-equation turbulence models at 

exiI1'ambient = 1.55. Three external grids (coarse, fine, and very 

fine) were generated and coupled with the internal grid for the 

toroidal screech simulation. The solution, calculated as laminar

flow on the coarse grid, converged and became steady. Turbulent-flow calculation for the same jet flow on the very fine grid 

has been performed near the jet exit. The upper part of the 

accompanying figure shows the early evolution of the calculated 

jet near the nozzle exit in terms of the density contours. The 

lower part of the figure shows the Mach contours of the calculated laminar flow on the coarse grid and compares the calculated and measured static pressures along the axis of symmetry. 

Significance 

A numerical ability to predict jet aeroacoustics will provide a 

path to a better understanding of the jet aeroacoustic phenomena and will be applied to existing and future high-performance 

aircraft and space-transportation vehicles. 

Future Plans 

Continued effort will be directed at completion of toroidal and 

helical screech simulations and investigation of turbulence 

model effects on the accuracy of jet-flow predictions. 

Calculated jet flow into a static ambient air; the upper part of the figure shows the early evolution of the jet near the nozzle exit and 

the lower part of the figure shows the Mach number contours with blue representing the low value and red the high value.

217
Advanced Computational Materials Research 

J. P. Wolfe, Principal Investigator 

Co-Investigators: R. Averback, J. Kogut, and P. Wolynes 

University of Illinois, Urbana/Champaign 

Research Objective 

To develop and apply new algorithmic approaches to problems 

in the science of materials, particularly those related to quantum 

systems and statistical problems. 

Approach 

A variety of computational approaches are used. 

Accomplishment Description 

Molecular dynamics computer simulations of energetic clusters 

with metal surfaces were conducted. The project examined the 

impact of ion cluster beams with solid surfaces. The accompanying figure shows a resulting mechanism map for cluster—solid 

interactions, where the size, energy, and cohesive properties 

were variables. It appears that it is best to have soft clusters 

impact hard substrates. We simulated the Ising model on a twodimensional dynamically triangulated lattice and determined 

that it has a second-order phase transition. The critical indices 

were calculated by finite scaling methods (see the accompanying 

figure), showing that the liquid lattice dresses the critical indices 

in accord with expectations. The simulation algorithm used a 

dual-lattice formulation that was amenable to vectorization. 

Significance 

The simulations have analytical and predictive value for current 

problems in materials. 

Future Plans 

The preliminary studies of dipolar excitons on a disordered 

materials conducted in this study will be compared with analytic 

theories. 

Publications 

1. Hsieh, H. and Averback, R. S. "Molecular-Dynamics 

Simulations of the Impact of Energetic Cu Clusters on Cu and 

Ni Substrates." Materials Research Society Symposium 

Proceedings 206 (1991): 375-380. 

2. Hsieh, H.; Averback, R. S.; Sellers, H.; and Flynn, C. P. 

"Molecular-Dynamics Simulations of Collisions Between 

Energetic Clusters of Atoms and Metal Substrates." Phys. Rev. 

Bulletin 45, no. 8 (1992): 4417-4430. 

3. Catterall, S. M.; Kogut, J. B.; and Renken, R. L "Scaling 

Behavior of the Ising Model Coupled to Two-Dimensional 

Quantum Gravity." To be published in Phys. Rev. Bulletin.

Mechanism diagram 

Implantation 

10

radiation damage 

4 Reflection Penetration I dissociation Penetration

formation 

cl E/N i / \melting 

sub 

Ecoh 7 \J

wetting craters 

Xburled clusters 

globs 

.5 1.0 1.5 

CI sub E / Eh 

Mechanism diagram of the interaction between energetic 

clusters of atoms and metal substrates. 

—.2 

—.4 

.6 

—.8 

—1.0

5

In N 

Finite-size scaling of the magnetization at the critical coupling 

on N2 lattices. 

218
Computational Electromagnetics 

Alex C. Woo, Principal Investigator 

Co-Investigators: Michael J. Schuh and Michael P. Simon 

NASA Ames Research Center 

Research Objective 

To develop the ability to predict the radar cross section (RCS) of 

simple jet-engine inlet models. 

Approach 

RCS3D, a finite-volume time-domain program, was used to 

integrate Maxwell's equations through a variety of structured 

grids. Computations were performed with the scattered field, 

where the total field equaled the scattered field plus the incident 

field, and the incident field is obtained from an analytical 

expression. The root mean square (rms) difference between the 

computed signature and experimental data was used as an 

estimate of the calculation error. 

Accomplishment Description 

Factors influencing the accuracy of the predicted RCS signature 

were investigated, including grid density, the number of computational zones, skewness of the grid cells, and harmonic versus 

Gaussian pulse illumination. The impact on required memory, 

CPU time, rate of convergence, and overall accuracy were 

assessed. Calculations were usually run over an azimuthal range 

of 0-90 degrees, where 0 degrees is looking directly into the 

mouth of the inlet. The measure of error in the computed RCS is 

the rms difference over the range from 0-45 degrees. Although 

the range from 45-90 degrees is not used in the error criterion, it 

is of interest since it was the disagreement with experiment 

within this range that prompted the use of a second-order 

boundary condition at conducting surfaces. For an inlet with 

wavelength dimensions of 1.25 x 1.25 x 1.875, the grid that has 

yielded the best result to date has seven zones that fill and 

envelop the inlet with an outer-zone width of one wavelength 

and nonuniform spacing inside the inlet. The scattered electric 

field for this computation and the comparison of predicted and 

measured RCS signatures are shown in the accompanying 

figures. The preceding result refers to calculations using harmonic excitation. Thus far, the pulse method has not yielded 

more economical results for a particular frequency than harmonic excitations, but recent results indicate that the pulse 

method may be advantageous for wide-band calculations. 

Significance 

A first-order boundary condition at conducting surfaces is 

inadequate; a second order condition must be used to obtain 

acceptable results. For horizontal polarization, a rms difference 

between predicted and measured RCS of less than 1 dB is 

achievable for the inlet using 5.2 megawords of memory and 

1.4 Cray Y-MP hours. 

Future Plans 

Work is directed toward the use of Gaussian pulse-incident 

fields, computations with an inlet embedded in an infinite 

ground plane, and the RCS of inlets with cylindrical and more 

complex geometries.

Scattered electric field of a square-aperture inlet for 30 degrees 

angle of incidence. 

20 

15 

10 

0 

C.)

0 

-5 

-10 

-15

) 20 40 60 80 

Azimuth (deg) 

Comparison of computed and experimental radar cross sections. 

219
Numerical Simulation of the PAMPA Jet Trainer 

Perry A. Wooden, Principal Investigator 

Co-Investigator: Ronald D. Lowe 

LTV, Aircraft Division 

Research Objective 

To investigate three-dimensional Euler and Navier—Stokes 

equations on an aircraft at high angle of attack and to highlight 

each application toward understanding the simulation and 

identifying areas of improvement to obtain detailed flow-field 

information of an airplane's aerodynamics. 

Approach 

The PAMPA jet trainer was simulated at a high angle of attack 

and transonic Mach number using the PARC3D code. The 

multiblock grid was generated using the GRIDGEN codes. 

Contiguous block interfacing was maintained to eliminate linear 

interpolation and PLOT313 was used to display the result. 

Accomplishment Description 

A numerical experiment was performed on the PAMPA jet 

trainer at a free-stream Mach number of 0.700 and 20 degrees 

angle of attack. A multiblock grid containing 70 blocks 

was generated using the GRIDGEN procedure. Only the wing 

and fuselage were modeled with the inlet faired to simulate the 

wind tunnel model. The total number of grid points for the 

Euler solution was 850,000 with the largest block containing 

58,000 points. The simulation required 2,500 iterations at a 

Courant number of 4 for each block to converge six orders of 

magnitude. The computations required 20 Cray-2 hours and 

4 megawords of memory. Results showed the computed lift 

coefficient differed by 7% and the pitching-moment coefficient 

was within 15% of the wind tunnel values. 

Significance 

High-angle-of-attack simulation capability of full configuration 

computational fluid dynamics (CFD) has increased significantly 

in recent years. As grid generation efficiency has increased, 

computer memory and speed increases have allowed CFD to be 

used as an effective tool in aircraft analysis. The influence of fullconfiguration PAMPA analysis provides the split in load distributions between wing and fuselage components for structural 

integrity over isolated component analysis. This has reduced 

risk, wind tunnel test time, and flight test time.

Future Plans 

A Navier—Stokes solution for the PAMPA jet trainer will be 

computed to improve prediction of the load distribution at high 

angle of attack. Target areas include increased code efficiency, 

reduced number of blocks to increase input/output performance, 

and unstructured grid solvers. 

Mach number color contours from a PARC3D Euler solution at 

Mach 0.700 and 20 degrees angle of attack. 

220
Numerical Simulation of Submarine Propulsion 

Cheng-1 Yang, Principal Investigator 

David Taylor Research Center 

Research Objective 

To simulate the high Reynolds-number flow over a sail—hull 

configuration. The flow pattern information will then be used for 

propulsor design. Flow phenomena of particular interest are the 

tip and horseshoe vortices originated from the sail top and the 

sail—hull juncture regions. Model tests are concurrently conducted at the towing tank at David Taylor Model Basin. Experiment data will be used for validation. 

Approach 

Viscous flow over a sail—hull combination is numerically 

simulated. Computations are performed with a modified FMC1 

code originally developed at NASA Langley. The method is 

based on an implicit high-resolution scheme for the threedimensional incompressible Navier—Stokes equations. The 

inviscid flux is discretized with total-variation-diminishing-like 

flux-differencing splitting and the viscous flux is discretized with 

central differencing. The computational grid is generated with a 

transfinite interpolation technique. 

Accomplishment Description 

In this simulation, the Reynolds number is 15 million based on 

the hull length. The vehicle is at a straight flight condition. A 

Baldwin—Lomax turbulence model is modified to account for the 

larger boundary interaction near the sail—hull juncture region. 

Typical flow patterns are shown in the accompanying figure. The 

flow patterns on the starboard and the port side are symmetrical, 

so only those on the starboard side are shown. The development 

of the hull boundary layer, the tip, and the horseshoe vortices 

are shown. Similar patterns were found in the towing-tank model 

test. The computational domain is composed of two blocks with 

an overlapped region to facilitate data communication between 

the blocks. A total of 400,000 grid points are used. Computation 

takes 5 megawords of memory and 2 Cray Y-MP hours. 

Significance 

The sail and other appendages may alter the flow pattern 

considerably and can produce adverse effects on propulsor 

performance. An accurate prediction of the inflow pattern 

enables the designer to create a propulsor which will meet all 

vehicle operating requirements.

Future Plans 

The present method will be used to study the viscous flow 

around a compound propulsor and will be integrated in the 

propulsor design cycle. The method will be used to study the 

detailed flow pattern on a propeller-blade surface and tip vortex. 

o.m I

a

Axial velocity contours. 

Cross-flow plane particle traces.

221 
Steady and Unsteady Flows at High Angle of Attack 

David T. Yeh, Principal Investigator 

Co-Investigator: Hiroshi Ide 

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 is used for highangle-of-attack flow simulations. 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 scheme is incorporated to achieve higher order spatial accuracy while maintaining 

numerical stability. The X-31 enhanced fighter maneuverability 

demonstrator aircraft is the configuration used in the study and 

was chosen because of its inherently designed high-angle-ofattack characteristics and capabilities, the existence of a wind 

tunnel data base, and the ongoing flight-test program. 

Accomplishment Description 

A blocked-grid approach was exploited to model the X-31 flow 

field. It applied both aligned and non-aligned blocks to capture 

the complexity of the geometry and associated flow physics. 

To identify the flow characteristics of the X-31 at high angle 

of attack, baseline computations were carried out at 15 and 

30 degrees angle of attack corresponding to pre- and post-stall 

flow regimes, respectively. The baseline calculations were made 

for all control surfaces at zero deflection. Converged turbulentflow solutions were obtained under the free-stream conditions of 

Mach 0.3 and Re, 1 x 1 06/foot. Numerical results were also 

obtained at 30 degrees angle of attack with the canard deflected 

under the same free-stream conditions to illustrate the effects of 

the movable canard on aerodynamic control in the post-stall 

regime. Numerical results compare favorably with wind tunnel 

measurements and flight-test data. A typical computation of the 

complete X-31 aircraft requires 25 megawords of memory and 

40 Cray Y-MP hours. 

Significance 

The numerical study parallels the ongoing X-31 flight-test 

program and establishes and extends computational fluid 

dynamics capabilities for analysis of high-angle-of-attack flow 

phenomena. This study will have significant impact on current 

and future military aircraft design.

Future Plans 

We will expand the study to include structural interface and 

thrust-vectoring paddles. 

Publication 

Yeh, D. T.; George, M. W.; Clever, W. C.; Tam, C. K.; and 

Woan, C. J. "Numerical Study of the X-31 High-Angle-of-Attack 

Flow Characteristics." AIAA Paper 91-1630, presented at the 

AIAA 22nd Fluid Dynamics, Plasma Dynamics, and Lasers 

Conference, Honolulu, HI, June 1991. 

Surface pressure distribution on an X-31 aircraft accompanied by 

a surface grid and some particle traces emanating from the wingroot leading-edge strake at 30 degrees angle of attack, Mach 0.3, 

and Re,, 1 x 1061foot. 

222
Volcanic Aerosol Clouds in the Stratosphere 

Richard E. Young, Principal Investigator 

Co-Investigators: Howard Houben and Owen B. Toon 

NASA Ames Research Center/Space Physics Research Institute 

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. 

Approach 

A three-dimensional spectral primitive-equation model is used to 

compute wind and temperature fields in the stratosphere. This 

model is coupled with a three-dimensional aerosol transport and 

microphysical model which computes the dispersion of the 

volcanic aerosol cloud using the computed winds from the 

circulation model and accounting for such processes as sedimentation and coagulation of the aerosol particles. 

Accomplishment Description 

Passive tracer simulations for both the El Chichon and 

Mt. Pinatubo volcanic aerosol clouds have been conducted 

for the six months following each eruption. Each simulation 

required about 12 Cray Y-MP hours and 16 megawords of 

memory. The results have been compared to the extensive 

satellite, airborne, and ground-based data sets which exist for 

each aerosol cloud. Most of the aerosol mass for the eruptions 

resided in the stratosphere above 20 km altitude. The El Chichon 

aerosol cloud stayed between the equator and 30°N for six 

months, while the Mt. Pinatubo cloud spread rapidly southward, 

reaching 20°S within weeks after the eruption. The model 

simulations reproduce these observed characteristics. The 

accompanying figure shows the Mt. Pinatubo volcanic gas and 

aerosol cloud after it had circumnavigated the globe one time. 

Initially, the cloud was a column between 20-25 km over the 

Philippine Islands, but it spread due to wind shear. 

Significance 

Major NASA programs study the climatology of the stratosphere. 

Unique opportunities to better understand stratospheric transport 

and aerosol microphysical processes occurred when the 

El Chichon volcano erupted in April 1982 and Mt. Pinatubo 

erupted in June 1991. By simulating the behavior of the volcanic 

clouds we can assess and improve theoretical understanding of 

transport and aerosol processes in the stratosphere.

Future Plans 

Multi-year simulations of the undisturbed stratosphere will be 

carried out. The results will be used to conduct passive-tracer 

simulations of the El Chichon and Mt. Pinatubo volcanic clouds. 

Once the passive-tracer computations are complete, fully 

interactive tracer simulations will be done. 

Publication 

Young, R. E.; Houben, H.; and Toon, 0. B. "Simulation of the 

Dispersion of the El Chichon and Mt. Pinatubo Stratospheric 

Aerosol Clouds During the First Year after the Eruptions." AGU 

Chapman Conference on Volcanoes, Climate, and Global 

Change, Hilo, HI, March 1992. 

The Mt. Pinatubo volcanic gas and aerosol cloud after it had 

circumnavigated the globe one time.

223
Single Expansion Ramp Nozzle with External Burning 

Shaye Yungster, Principal Investigator 

ICOMP/NASA Lewis Research Center 

Research Objective 

To develop computational fluid dynamics (CFD) prediction 

techniques to support external burning and nozzle performance 

experiments aimed at reducing the transonic drag of National 

Aero-Space Plane (NASP) vehicles. 

Approach 

The Reynolds-averaged Navier-Stokes equations with finiterate chemistry are solved using two- and three-dimensional 

multiblock fully implicit codes based on the LU-SSOR factorization scheme. The spatial discretization is based on a second 

order total-variation-diminishing scheme for the two-dimensional 

code, and central differences with second- and fourth-order 

artificial damping for the three-dimensional code. Several finiterate chemistry models for hydrogen-oxygen combustion are 

being used, ranging from a one-step irreversible chemical 

reaction to a nine species 18-step reaction mechanism. The 

Baldwin-Lomax algebraic turbulence model is used. 

Accomplishment Description 

The effects of external combustion on NASP-like nozzles at 

transonic speeds were numerically investigated. A computational 

model for the external fuel injection and combustion processes 

was developed, and results were obtained for complete nozzle 

configurations with hot and cold exhaust flow, and with or 

without external burning. A typical two-dimensional turbulent 

calculation with a one-step combustion model is shown in the 

accompanying figure. The combustion occurs in a layer between 

the external stream and the internal nozzle plume. Proper 

handling of the downstream boundary condition was important 

due to the large subsonic flow region created by the external 

burning process. A multiblock grid containing a total of 

25,000 mesh points was used requiring about 5 megawords of 

memory and 4 Cray-2 hours. At transonic speeds and without 

external burning, three-dimensional effects become important 

due to nozzle over expansion. Three-dimensional calculations 

of the nozzle and external flow (without external burning) 

have been performed utilizing four blocks, each having 

140 x 61 x 10 grid points. These computations require about 

20 megawords of memory and 12 Cray-2 hours. 

Significance 

NASP-like nozzles are highly over-expanded at transonic speeds 

and this results in large drag increases. One possible 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.

Future Plans 

The present model will be enhanced by including complete 

reaction mechanisms and by detailed modeling of the external 

fuel-injection process. Simulations are planned for large-scale 

flame-holding and ignition experiments and hot and cold nozzle 

tests of current and advanced NASP configurations. 

o Experiment 

- External burning 

- - No external burning 

-.4 -.2 0 .2 .4 .6 .8 1.0

x/L

- - 

Mach number contours and upper surface pressure distribution

for a hot exhaust nozzle with external burning; = 1.2, 

NPR = 6, and F/A = 0.021. 

C 

2

i 

224
Transition Modeling 

T. A. Zang, Principal Investigator 

Co-Investigators: S. Dinavahi, U. Piomelli, R. Garg, and K. Ball 

NASA Langley Research Center/University of Maryland/University of Texas 

Research Objective 

To develop effective models for the process by which flows 

transition from smooth laminar states to chaotic turbulent states. 

Approach 

Perform direct numerical simulations (DNS) and large-eddy 

simulations (LES) of simple, low-speed flows under controlled 

conditions in the transitional and turbulent regimes. These are 

used for investigation of basic transition physics as well as for 

assessing the utility of various LES, Reynolds-averaged, and 

dynamical system approaches. 

Accomplishment Description 

The DNS of fundamental channel flow transition at Reynolds 

number 8,000, begun in 1989, was continued into the fully 

turbulent regime. This data base was analyzed to furnish the 

Reynolds-stress budgets for the fully turbulent regime, but more 

interestingly, for the transitional regime. In addition, the proper 

orthogonal decomposition was performed for the turbulent state. 

In a separate investigation, DNS was applied to tuned and 

detuned boundary-layer transition. The latter type of transition is 

more closely related to natural transition than the former. The 

results indicated that a small amount of detuning, even if only in 

the form of background noise, of an otherwise tuned transition 

problem can have a dramatic effect upon experimental diagnostics of high-shear layers. This is illustrated in the accompanying 

figure, which displays the mean velocity profiles in the symmetry 

plane for three transition simulations. The initial conditions for 

each included the same primary wave, but they differed in the 

initial three-dimensional component: a tuned subharmonic, a 

pure detuned mode, and a tuned subharmonic plus a small,

detuned, "environmental," component. A strong inflectional 

profile (indicative of high-shear layers) appeared only for the first 

case. Detuning led to a much milder average inflection (closer to 

experimental measurements), although closer examination 

revealed the presence of intense local inflections. Typical runs 

took 50-100 Cray-2 hours and 40 megawords of main memory. 

Significance 

The DNS of boundary-layer transition may explain why certain 

transition experiments have not produced evidence of high-shear 

layers. 

Future Plans 

The later stages of detuned transition will be investigated. LES 

methodology will be more thoroughly tested and then extended 

to spatial and compressible transition. 

Publications 

1. Zang, T. A. "Aspects of Laminar Breakdown in BoundaryLayer Transition." ICASE/LaRC Workshop on Transition and 

Turbulence, ed. M. Y. Hussaini. New York: Springer, 1992. 

2. Dinavahi, S. P. G. and Zang, T. A. "Reynolds Stress Budget in 

a Transitional Channel Flow." ICASE/LaRC Workshop on 

Transition and Turbulence, ed. M. Y. Hussaini. New York: 

Springer, 1992. 

3. Garg, R. P. and Piomelli, U. "Large-Eddy Simulation of 

Detuned Transition in Plane Channel Flow." 4th International 

Symposium on Computational Fluid Dynamics. Davis, CA, 

Sept. 1991. 

Detuned Detuned 

Tuned (forcibly) (environmentally) 

- Transitional 

Blasius

0 

0 .5 1.0 .5 1.0 .5 1.0 

'CU> 'CU> <U> 

Mean velocity profiles in the wall-normal direction from three related simulations of boundary-layer transition. The wall-normal 

coordinate is scaled by the local displacement thickness *• The laminar (Blausius) profile is included for reference.

225
Index by Research Sites

Page 

A.T.M., Inc. 

Jeffrey N. Cuzzi Particle-Gas Dynamics in the Protoplanetary Nebula............................................................................36 

(with NASA Ames Research Center and University of California, Santa Cruz) 

Adaptive Research Corporation 

R.R. Chamberlain Aerothermal Analysis of Hypersonic Defense Interceptors ....................................................................20 

Advanced Projects Research, Inc. 

Joseph W. Humphrey Hypersonic Scramjet and Detonation Flows .........................................................................................83 

(with University of Arizona and NASA Langley Research Center) 

Darrell W. Pepper Hypersonic Scramjet Flow ..................................................................................................................151 

Amtec Engineering, Inc. 

Scott T. Imlay Aero-Assisted Orbital Transfer Vehicle Flow Fields ...............................................................................86 

Analytical Services and Materials, Inc. 

William B. Compton, Ill Three-Dimensional Afterbody Flow with Jet Exhaust.............................................................................31 

(with NASA Langley Research Center) 

Charles R. McClinton High Mach-Number Mixing in a Shock Tunnel Environment ..............................................................129 

(with NASA Langley Research Center) 

Charles R. McClinton National Aero-Space Plane Configuration Trade Studies .....................................................................130 

(with NASA Langley Research Center) 

Charles R. McClinton Scramjet Inlet Interaction ....................................................................................................................1 31 

(with NASA Langley Research Center) 

Charles R. McClinton Scramjet Engine Design Optimization ................................................................................................132 

(with NASA Langley Research Center) 

S.Paul Pao Afterbody Aerodynamics with Canted Pitch-Vectoring Twin Nozzles .................................................149 

(with NASA Langley Research Center) 

C. D. Pruett Laminar Breakdown of High-Speed Boundary-Layer Flow ..................................................................158 

(with NASA Langley Research Center) 

R. Clayton Rogers Hypervelocity Mixing and Combustion in Pulse Facility Flows ...........................................................165 

(with NASA Langley Research Center and University of Missouri, Rolla) 

Aurora Science, Inc. 

C. S. Lin Electron-Beam Injections from the Space Shuttle ................................................................................118 

(with Southwest Research Institute) 

Bell Helicopter Textron, Inc. 

J. C. Narramore Rotorcraft Drag Prediction ..................................................................................................................144 

PRECEDING PAGE BLANK NOT flLMED

227
Page 

Berkeley Research Associates 

Stephen H. Brecht Global-Venus/Solar-Wind Interaction ...................................................................................................14 

Boeing Commercial Airplane Group 

Philippe R. Spalart complex Turbulent Boundary Layers ..................................................................................................186 

Boeing Company, The 

H. C. Chen Euler Analysis of Turboprop and Turbofan Integration ..........................................................................24 

(with ViGYAN, Inc. and NASA Langley Research Center) 

California Polytechnic State University, San Luis Obispo 

Karlin R. Roth Validation of a Short Takeoff and Vertical Landing Model ..................................................................170 

(with NASA Ames Research Center) 

Carnegie Mellon University 

Wei J. Chyu Airframe and Inlet Aerodynamics..........................................................................................................29 

(with NASA Ames Research Center and Cornell University) 

Tom I-P. Shih Wankel Engine Flow Fields.................................................................................................................179 

(with NASA Lewis Research Center) 

Case Western Reserve University 

Russell G. DeAnna Transition over a Rough Surface ...........................................................................................................42 

(with U.S. Army Propulsion Directorate, AVSCOM) 

Isaac Greber Numerical Calculation of a Three-Dimensional Separated Flow ........................................................... 63 

Center for Turbulence Research 

Parviz Moin Shock-Wave/Turbulent-Boundary-Layer Interaction ............................................................................ 140 

(with Stanford University and NASA Ames Research Center) 

Cornell University 

Wei I. Chyu Airframe and Inlet Aerodynamics.......................................................................................................... 29 

(with NASA Ames Research Center and Carnegie Mellon University) 

David Taylor Research Center 

Chao-Ho Sung Computational Fluid Dynamics for Naval Applications ...................................................................... 193 

Cheng-I Yang Numerical Simulation of Submarine Propulsion.................................................................................. 220 

Douglas Aircraft Company 

Lee T. Chen Advanced Transonic Wing Concepts .................................................................................................... 25 

Gregory A. Intemann Wing—Nacelle--Pylon Installations......................................................................................................... 87 

(with McDonnell Aircraft Company)

Walter 0. Valarezo Multi-Element High-Lift Concepts .......................................................................................................202 

(with NASA Langley Research Center and ICASE) 

228 
Page 

Eloret Institute 

G. S. Deiwert Self-Adaptive Grid Code Applied to Complex Three-Dimensional Flows ..............................................45 

(with NASA Ames Research Center and Sterling Software) 

Larry L. Erickson Flow Solver for Euler Equations on Unstructured Tetrahedral Meshes ...................................................53 

(with NASA Ames Research Center) 

General Dynamics, Convair Division 

Jack H. Kennedy Aeronautical Vehicle Radar Cross Sections ...........................................................................................96 

General Dynamics, Electric Boat Division 

Donald W. Davis Hydrodynamic Performance Evaluation ................................................................................................40 

General Dynamics, Fort Worth Division 

Steve L. Karman, Jr. Unstructured Grid-Generation/Flow-Solver Calibration ........................................................................93 

Brian R. Smith Computational Fluid Dynamics Methods for Highly Maneuverable Aircraft........................................181 

General Dynamics, Space Systems Division 

Thanh T. Phan Rocket-Base Flow-Field Simulations ...................................................................................................153 

General Electric Aircraft Engines 

Robert F. Bergholz Three-Dimensional Exhaust-Nozzle Flow Fields for High-Speed Civil Transport ...................................10 

Joseph L. Garrett Integrated Hypersonic-Propulsion Flow Paths .......................................................................................59 

(with Pratt & Whitney) 

Sundares V. Subramanian Viscous Reacting-Flow Applications to Scramjet Propulsion ...............................................................191 

General Motors Corporation, Allison Gas Turbine Division 

Andrew J. Crook Endwall and Casing Treatment Flow in a Transonic Fan Rotor..............................................................33 

Edward J. Hall Unsteady Counterrotation of Ducted Propfans ......................................................................................69 

Jeff L. Hansen Incorporation of a Three-Dimensional Multistage Viscous Code into a Compressor 

DesignSystem ......................................................................................................................................74 

(with NASA Lewis Research Center) 

Kurt F. Weber Flow Simulation in Fan and Core Compressor Stages..........................................................................210 

General Motors Technical Center 

Budugur Lakshminarayana Steady and Unsteady Turbomachinery Flow Fields .............................................................................108 

(with Pennsylvania State University) 

GEOLAB 

Robert E. Smith Grid Generation for Aerodynamic Configurations ............................................................................... 184 

(with NASA Langley Research Center)

229
Page 

Georgia Institute of Technology 

James G. Brasseur Local Passive Scalar Dispersion in a Turbulent Boundary Layer............................................................13 

(with Pennsylvania State University) 

High Technology Corporation 

M. R. Malik Supersonic Boundary-Layer Transition on a Cone at Incidence ......................... .................................. 123 

(with NASA Langley Research Center) 

M. R. Malik Compressible Boundary-Layer Transition............................................................................................ 124 

(with NASA Langley Research Center) 

ICASE 

M. G. Macaraeg Ignition and Structure of a Diffusion Flame with Vortex Interaction .................................................... 1 21 

(with NASA Langley Research Center and Old Dominion University) 

Dimitri J. Mavriplis Multigrid Solution of the Euler Equations ............................................................................................ 128 

(with NASA Langley Research Center) 

Eli Turkel Multigrid Solution of the Navier—Stokes Equations .............................................................................. 201 

(with NASA Langley Research Center) 

Walter 0. Valarezo Multi-Element High-Lift Concepts ....................................................................................................... 202 

(with Douglas Aircraft Company and NASA Langley Research Center) 

ICOMP 

Shaye Yungster Single Expansion Ramp Nozzle with External Burning ........................................................................ 224 

(with NASA Lewis Research Center)

Imperial College 

Rajiv Thareja Unstructured Multigrid Euler Solver ....................................................................................................196 

(with Lockheed Engineering and Sciences Company and University College, Wales) 

Iowa State University 

Richard H. Pletcher Three-Dimensional Liquid-Sloshing Flows ..........................................................................................154 

JAI Associates, Inc. 

Samuel P. Shanks Computational Fluid Dynamics of Store Separation ............................................................................178 

Jason Associates 

Thomas T. Huang Submarine-Appendage Design and Turbulence Modeling ...................................................................81 

(with Naval Surface Warfare Center; Scientific Research Associates, Inc.; 

and MCAT Institute) 

Jet Propulsion Laboratory 

David Halpern Satellite Data Assimilation and Ocean General Circulation Models ......................................................70 

(with University of California, Los Angeles) 

David Halpern Atmospheric General Circulation Model Sensitivity to Sea Surface Temperature Fields.........................71 

(with University of California, Los Angeles) 

230 
Page 

Christian L. Keppenne Orographically Forced Oscillations of the Martian Atmosphere ............................................................97 

Korea Advanced Institute of Science and Technology 

Carolyn R. Kaplan Unsteady Ethylene Jet-Diffusion Flames ................................................................................................92 

(with Naval Research Laboratory) 

Lockheed Engineering and Sciences Company 

Lawrence D. Huebner Generic National Aero-Space Plane Fuselage Configuration Study .......................................................82 

(with NASA Langley Research Center) 

Kenneth M. Jones Analysis of High-Speed Civil Transport Configurations .........................................................................88 

(with NASA Langley Research Center and ViGYAN, Inc.) 

Elizabeth M. Lee Subsonic/Transonic Flutter Boundaries ...............................................................................................112 

(with NASA Langley Research Center) 

Ajay K. Pandey Thermo-Viscoplastic Analysis of ari Engine-Cowl Leading Edge..........................................................148 

(with NASA Langley Research Center) 

Ramadas K. Prabhu Hypersonic-Flow Computations using Adaptive Unstructured Meshes................................................157 

(with NASA Langley Research Center) 

Rajiv Thareja Unstructured Multigrid Euler Solver .................................................................................................... 196 

(with Imperial College and University College, Wales) 

Gururaja R. Vemaganti Two-Equation Turbulence-Model Implementation ..............................................................................206 

LTV, Aircraft Division 

Perry A. Wooden Numerical Simulation of the PAMPA Jet Trainer .................................................................................220 

Martin Marietta Astronautics Group 

C. T. Edquist Aerobraking Studies of Three-Dimensional Nonequilibrium Viscous Flow............................................49 

MCAT Institute 

Thomas A. Edwards Aerodynamic Optimization of Supersonic Transport Aircraft ................................................................50 

(with NASA Ames Research Center)

Thomas T. Huang Submarine-Appendage Design and Turbulence Modeling ....................................................................81 

(with Naval Surface Warfare Center; Jason Associates; and Scientific Research 

Associates, Inc.) 

Scott L. Lawrence Integrated Hypersonic Vehicle Flow-Field Analysis ............................................................................110 

(with NASA Ames Research Center) 

W. R. Van Dalsem Simulation of the Unsteady Flow about Transonic Cavities .................................................................203 

(with NASA Ames Research Center) 

W. R. Van Dalsem Powered-Lift Computational Fluid Dynamics Project..........................................................................204 

(with NASA Ames Research Center)

231 
Page 

McDonnell Aircraft Company 

Shreekant Agrawal Numerical Calculation of Vortical Flows ................................................................................................3 

Alan B. Cain Simulation and Modeling of Turbulence and Flow Acoustics................................................................16 

(with NASA Ames Research Center) 

Raymond R. Cosner Analyses of F/A-i 8E/F Upgrades ...........................................................................................................32 

David W. Halt Transonic Analysis on Unstructured Grids ............................................................................................72 

(with McDonnell Douglas Research Laboratories) 

Gregory A. Intemann Wing—Nacelle--Pylon Installations.........................................................................................................87 

(with Douglas Aircraft Company) 

Linda D. Kral Turbulence Modeling for Three-Dimensional Flow Fields...................................................................103 

(with McDonnell Douglas Research Laboratories) 

McDonnell Douglas Corporation 

C. C. Lee Numerical Investigation of Vehicles Similar to the National Aero-Space Plane ...................................111 

McDonnell Douglas Research Laboratories 

Ramesh K. Agarwal Numerical Solution of Three-Dimensional Maxwell's Equations .............................................................1 

Ramesh K. Agarwal Helicopter Fuselage, Rotor, and Rotor/Body Interaction Flow-Field Calculations ....................................2 

William W. Bower Short Takeoff and Vertical Landing Aircraft Thermal/Acoustic Loads.....................................................12 

(with Nielsen Engineering and Research, Inc.) 

Jerry E. Deese Flow about Almost Complete Aircraft and Hypersonic Configurations ..................................................43 

(with NASA Ames Research Center) 

David W. Halt Transonic Analysis on Unstructured Grids ............................................................................................72 

(with McDonnell Aircraft Company) 

Linda D. Kral Receptivity, Transition, and Turbulence Phenomena .......................................................................... 102 

Linda D. Kral Turbulence Modeling for Three-Dimensional Flow Fields................................................................... 103 

(with McDonnell Aircraft Company) 

David L. Marcum Flow-Feld Calculations using Unstructured Grids ............................................................................... 126 

(with Mississippi State University) 

Mississippi State University 

David L. Marcum Flow-Field Calculations using Unstructured Grids .............................................................................. 126 

(with McDonnell Douglas Research Laboratories) 

NASA Ames Research Center 

Jeffrey R. Barnes Nonlinear Baroclinic Instability ............................................................................................................... 6 

(with Oregon State University)

232 
Page 

Alan B. Cain Simulation and Modeling of Turbulence and Flow Acoustics................................................................16 

(with McDonnell Aircraft Company) 

Frank Caradonna Prediction of Advanced Rotor Performance ..........................................................................................19 

(with U.S. Army Aeroflightdynamics Directorate, AVSCOM) 

Dean R. Chapman Boost-Phase Detection Study ................................................................................................................21 

(with Stanford University) 

Denny Chaussee Transonic and Supersonic Flow Past Aircraft Configurations .................................................................22 

C. L. Chen Numerical Analysis of Three-Dimensional Separated Juncture Flows....................................................23 

(with Rockwell International Science Center) 

Wei J. Chyu Airframe and Inlet Aerodynamics..........................................................................................................29 

(with Cornell University and Carnegie Mellon University) 

Jeffrey N. Cuzzi Particle-Gas Dynamics in the Protoplanetary Nebula............................................................................36 

(with A.T.M., Inc. and University of California, Santa Cruz) 

Jerry E. Deese Flow about Almost Complete Aircraft and Hypersonic Configurations ..................................................43 

(with McDonnell Douglas Research Laboratories) 

G. S. Deiwert Self-Adaptive Grid Code Applied to Complex Three-Dimensional Flows ..............................................45 

(with Sterling Software and Eloret Institute) 

Thomas A. Edwards Aerodynamic Optimization of Supersonic Transport Aircraft ................................................................50 

(with MCAT Institute) 

Larry L. Erickson Flow Solver for Euler Equations on Unstructured Tetrahedral Meshes ....................................................53 

(with Eloret Institute) 

William J. Feiereisen Discrete Particle Simulation of Compressible Flow ...............................................................................54 

(with Stanford University) 

William J. Feiereisen Aeroassist Flight Experiment Flow Simulation ....................................................................................... 55 

Karen L. Gundy-Burlet Rotor—Stator Interaction in Turbomachines ........................................................................................... 66 

Guru P. Guruswamy Fluid and Structure Integration for Aerospace Applications ................................................................... 67 

Dochan Kwak High-Lift Aerodynamic Flow............................................................................................................... 105 

Dochan Kwak Space Shuttle Main Engine Flow......................................................................................................... 106 

Scott L. Lawrence Integrated Hypersonic Vehicle Flow-Field Analysis ............................................................................ 110 

(with MCAT Institute)

Robert D. MacElroy Computer Simulation of Water—Membrane Interfaces .........................................................................122 

Nagi N. Mansour Turbulence in Compressible Fluids .....................................................................................................125 

Fred W. Martin, Jr. Space Shuttle Flow Field.....................................................................................................................127 

(with NASA Johnson Space Center)

233 
Page 

W. J. McCroskey Aerodynamics and Acoustics of Rotorcraft .......................................................................................... 1 33 

(with U.S. Army Aeroflightdynamics Directorate, AVSCOM) 

W. J. McCroskey High-Performance Rotor-Blade Tips.................................................................................................... 134 

(with U.S. Army Aeroflightdynamics Directorate, AVSCOM) 

ParvizMoin Turbulent Flow Over Riblets ............................................................................................................... 137 

(with Stanford University) 

Parviz Moin Turbulence Over a Backward-Facing Step .......................................................................................... 138 

Parviz Moin Aerodynamic Sound Generation ......................................................................................................... 139 

(with Stanford University) 

Parviz Moin Shock-Wave/Turbulent-Boundary-Layer Interaction ............................................................................ 140 

(with Stanford University and Center for Turbulence Research) 

Robert D. Moser Wall-Bounded Turbulent Flows ..........................................................................................................141 

Grant Palmer Space Transportation Vehicle Aerothermodynamics ........................................................................... 147 

Leonhard Pfister Wave-Induced Transports ................................................................................................................... 152 

James B. Pollack Martian Atmosphere General Circulation ........................................................................................... 155 

Michael M. Rogers Mixing and Reacting in Plane Mixing Layers....................................................................................... 164 

(with Stanford University and University of Wisconsin, Madison) 

William C. Rose Hypersonic Inlet Flow Fields............................................................................................................... 166 

William C. Rose National Aero-Space Plane Inlet Flow Fields ....................................................................................... 168 

James C. Ross Trapped-Vortex Flows on Highly Swept Wing Configurations ............................................................. 169 

Karlin R. Roth Validation of a Short Takeoff and Vertical Landing Model .................................................................. 170 

(with California Polytechnic State University, San Luis Obispo) 

Stephen M. Ruffin Coupled Rotation—Vibration—Dissociation Processes in Hypersonic Flows .......................................... 171 

Lewis B. Schiff High Alpha Technology Program F-i 8 Aerodynamics ........................................................................ 174 

Lewis B. Schiff High Angle-of-Attack Vortex Flow Aerodynamics ............................................................................... 175 

Merritt H. Smith YAV-813 Aircraft Flow Simulation........................................................................................................ 183

Bruce F. Smith Numerical. Experiments in the Formation and Evolution of Galaxies ...................................................182 

(with University of Chicago) 

K. Snyder National Aero-Space Plane High-Speed Combustion Experiment........................................................185 

(with NASP Joint Program Office) 

Michael W. Stortz Harrier YAV-813 Wing Aerodynamics .................................................................................................. 190 

Gary E. Thomas Multi-Dimensional Simulations of Noctilucent Cloud Formation ........................................................197 

(with University of Colorado, Boulder) 

234 
Page 

Owen B. Toon Stratospheric Ozone Destruction ........................................................................................................198 

Arnaud Trouvé Turbulent Premixed Combustion ........................................................................................................200 

(with Stanford University) 

W. R. Van Dalsem Simulation of the Unsteady Flow about Transonic Cavities .................................................................203 

(with MCAT Institute) 

W. R. Van Dalsem Powered-Lift Computational Fluid Dynamics Project ..........................................................................204 

(with MCAT Institute) 

Douglas L. Westphal Tropospheric Aerosols and Clouds .....................................................................................................213 

Alex C. Woo Computational Electromagnetics.........................................................................................................219 

Richard E. Young Volcanic Aerosol Clouds in the Stratosphere ....................................................................................... 223 

(with Space Physics Research Institute) 

NASA Johnson Space Center 

Fred W. Martin, Jr. Space Shuttle Flow Field.....................................................................................................................127 

(with NASA Ames Research Center) 

NASA Langley Research Center 

Jassim A. AI-Saadi Transonic Reynolds Number Effects on Aircraft Configurations ............................................................... 4 

(with ViGYAN, Inc.) 

John T. Batina Euler Flutter Analysis of a Complex Aircraft Configuration ......................................................................7 

(with Purdue University) 

Richard L. Campbell Automated Transonic Wing Design ...................................................................................................... 1 7 

H. C. Chen Euler Analysis of Turboprop and Turbofan Integration .......................................................................... 24 

(with The Boeing Company and ViGYAN, Inc.) 

Charles E. Cockrell, Jr. Generic National Aero-Space Plane Forebody—Inlet Integration ............................................................ 30 

William B. Compton, Ill Three-Dimensional Afterbody Flow with Jet Exhaust............................................................................. 31 

(with Analytical Services and Materials, Inc.) 

Virendra K. Dogra Rarefied Hypersonic Condition Wake Structures................................................................................... 47 

(with ViGYAN, Inc.)

J. Philip Drummond Fuel—Air Mixing Enhancement by Jet—Shock Interactions ......................................................................48 

Dean R. Eklund Transpiration Cooling for Scramjet Combustor Flow Fields ...................................................................52 

(with National Research Council) 

Michael C. Fischer Fl 6xL Supersonic Laminar Low-Control Experiment .............................................................................56 

Neal T. Frink Analysis of Slender-Wing Geometries using Unstructured Grids ...........................................................57 

Thomas B. Gatski Compressible Turbulent Flows..............................................................................................................60 

235 
Page 

Peter A. Gnoffo 

W. L. Grose 

H. Harris Hamilton 

Julius E. Harris 

Lin C. Hartung 

Lawrence D. Huebner 

Joseph W. Humphrey 

Kenneth M. Jones 

Ajay Kumar 

Elizabeth M. Lee 

James M. Luckring 

M. G. Macaraeg 

M. R. Malik 

M. R. Malik 

Dimitri J. Mavriplis 

Charles R. McClinton 

Charles R. McClinton 

Charles R. McClinton 

Charles R. McClinton 

James N. Moss 

Hypersonic Flows in Chemical and Thermal Nonequilibrium...............................................................61 

Three-Dimensional Atmospheric Simulation Model .............................................................................. 65 

Radiative Structure in Aerobrake Shock Layers ..................................................................................... 73 

(with ViGYAN, Inc.) 

Swept-Wing Leading-Edge Transition.................................................................................................... 75 

(with ViGYAN, Inc.) 

Radiation Transport around Axisymmetric Blunt-Body Vehicles............................................................ 76 

Generic National Aero-Space Plane Fuselage Configuration Study ....................................................... 82 

(with Lockheed Engineering and Sciences Company) 

Hypersonic Scramjet and Detonation Flows ......................................................................................... 83 

(with Advanced Projects Research, Inc. and University of Arizona) 

Analysis of High-Speed Civil Transport Configurations ......................................................................... 88 

(with ViGYAN, Inc. and Lockheed Engineering and Sciences Company) 

Translating-Strut Scramjet Inlet ........................................................................................................... 104 

Subsonic/Transonic Flutter Boundaries ............................................................................................... 112 

(with Lockheed Engineering and Sciences Company) 

Advanced Computational Fluid Dynamics Applications for Complex Configurations.......................... 120 

Ignition and Structure of a Diffusion Flame with Vortex Interaction .................................................... 1.21 

(with Old Dominion University and ICASE) 

Supersonic Boundary-Layer Transition on a Cone at Incidence........................................................... 123 

(with High Technology Corporation) 

Compressible Boundary-Layer Transition ............................................................................................ 124 

(with High Technology Corporation) 

Multigrid Solution of the Euler Equations ............................................................................................ 128 

(with ICASE) 

High Mach-Number Mixing in a Shock Tunnel Environment .............................................................. 129 

(with Analytical Services and Materials, Inc.) 

National Aero-Space Plane Configuration Trade Studies ..................................................................... 1 30 

(with Analytical Services and Materials, Inc.) - 

ScramjetInlet Interaction .................................................................................................................... 131 

(with Analytical Services and Materials, Inc.) 

Scramjet Engine Design Optimization................................................................................................. 132 

(with Analytical Services and Materials, Inc.) 

Hypersonic Rarefied Flow about a Compression Ramp....................................................................... 142

236
Page 

Ahmed K. Noor Thermomechanical Buckling and Post-Buckling of Multilayered Composite Panels 

withCutouts ....................................................................................................................................... 146 

(with University of Virginia) 

Ajay K. Pandey Thermo-Viscoplastic Analysis of an Engine-Cowl Leading Edge .......................................................... 148 

(with Lockheed Engineering and Sciences Company) 

S. Paul Pao Afterbody Aerodynamics with Canted Pitch-Vectoring Twin Nozzles ................................................. 149 

(with Analytical Services and Materials, Inc.) 

Ramadas K. Prabhu Hypersonic-Flow Computations using Adaptive Unstructured Meshes................................................ 157 

(with Lockheed Engineering and Sciences Company) 

C. D. Pruett Laminar Breakdown of High-Speed Boundary-Layer Flow .................................................................. 158 

(with Analytical Services and Materials, Inc.) 

R. Ramakrishnan Shock Impingement in Hypersonic Engine Inlets................................................................................. 159 

R. Clayton Rogers Hypervelocity Mixing and Combustion in Pulse Facility Flows ........................................................... 165 

(with University of Missouri, Rolla and Analytical Services and Materials, Inc.) 

Christopher L. Rumsey High-Alpha Flow Fields ...................................................................................................................... 1 72 

B. A. Singer Formation and Growth of Hairpin Vortices ......................................................................................... 180 

Robert E. Smith Grid Generation for Aerodynamic Configurations ............................................................................... 184 

(with GEOLAB)

Olaf 0. Storaasli Parallel-Vector "Out-of-Core" Equation Solver....................................................................................189 

R. C. Swanson Simulation of Scramjet Flow Fields .....................................................................................................194 

Eli Turkel Multigrid Solution of the Navier—Stokes Equations ..............................................................................201 

(with ICASE) 

Walter 0. Valarezo Multi-Element High-Lift Concepts .......................................................................................................202 

(with Douglas Aircraft Company and ICASE) 

Veer N. Vatsa High-Reynolds-Number Viscous Flow over Aircraft Components .......................................................205 

Kenneth J. Weilmuenster Winged Entry-Vehicle Computations ..................................................................................................211 

Kenneth J. Weilmuenster Shuttle Leeside Temperature Study .....................................................................................................212 

J. A. White Supersonic Chemically Reacting Turbulent Flow ................................................................................214 

Richard G. Wilmoth Shock Interactions in Hypersonic Rarefied Flows ................................................................................. 215 

(with ViGYAN, Inc.) 

T. A. Zang Transition Modeling ............................................................................................................................. 225 

(with University of Maryland and University of Texas)

237 
Page 

NASA Lewis Research Center 

Rodr,ckv. Chima Prediction of Turbine Endwall Heat Transfer.........................................................................................27 

Chuen-Yen Chow Compressible Taylor—Couette Flow.......................................................................................................28 

(with University of Colorado, Boulder) 

Robert G. Deissler Sensitivity of Turbulence to Initial Conditions ....................................................................................... 44 

Jeff L. Hansen Incorporation of a Three-Dimensional Multistage Viscous Code into a Compressor 

DesignSystem ...................................................................................................................................... 74 

(with General MotOrs Corporation, Allison Gas Turbine Division) 

James D. Holdeman Hot-Gas Ingestion by a Short Takeoff and Vertical Landing Aircraft in Ground Proximity ..................... 78 

(with University of Illinois, Urbana/Champaign) 

Osama A. Kandil Simulation and Control of Slender Wing Rock ...................................................................................... 91 

(with Old Dominion University) 

Suk C. Kim Low-Thrust Chemical Rockets............................................................................................................... 98 

(with Sverdrup Technology, Inc.) 

S. W. Kim Fluid Flow of jets in Cross Flow ............................................................................................................ 99 

Kevin R. Kirtley Multistage Turbomachinery Flows ...................................................................................................... 100 

(with Sverdrup Technology, Inc.) 

Sang Soo Lee Resonant-Triad Interaction in an Adverse Pressure-Gradient Boundary Layer...................................... 113 

(with Sverdrup Technology, Inc.) 

William C. Rose National Aero-Space Plane Inlet Boundary-Layer Control ................................................................... 167 

Jose M. Sanz Aerodynamic Inverse Design and Analysis for a Full Engine ............................................................... 173 

T. I-P. Shih Wankel Engine Flow Fields .................................................................................................................. 179 

(with Carnegie Mellon University) 

Shaye Yungster Single Expansion Ramp Nozzle with External Burning ........................................................................ 224 

(with ICOMP)

NASA Marshall Space Flight Center 

Edward J. Reske Complex Three-Dimensional Flows in the Advanced Solid-Rocket Motor ..........................................160 

Eric T. Stewart Space Station Freedom Internal Flow Analysis ....................................................................................188 

(with Sverdrup Technology, Inc.) 

NASP Joint Program Office 

K. Snyder National Aero-Space Plane High-Speed Combustion Experiment........................................................185 

(with NASA Ames Research Center) 

National Research Council 

Dean R. Eklund Transpiration Cooling for Scramjet Combustor Flow Fields ...................................................................52 

(with NASA Langley Research Center) 

238 
Page 

Naval Research Laboratory 

Russell B. Dahlburg Dynamical Modeling of the Solar Atmosphere......................................................................................37 

Fernando F. Grinstein Spatially Evolving Reactive Jets .............................................................................................................64 

Fazle Hussain Chemically Reacting Free-Shear Flows .................................................................................................85 

(with University of Houston) 

K. Kailasanath Structure and Dynamics of Multidimensional Flames............................................................................89 

K. Kailasanath Unsteady Nozzle Flow Fields ...............................................................................................................90 

Carolyn R. Kaplan Unsteady Ethylene Jet-Diffusion Flames ................................................................................................92 

(with Korea Advanced Institute of Science and Technology) 

Naval Surface Warfare Center 

Carl T. Hsieh High Angle-of-Attack Missile Aerodynamics .........................................................................................80 

Thomas T. Huang Submarine-Appendage Design and Turbulence Modeling ....................................................................81 

(with Jason Associates; Scientific Research Associates, Inc.; and MCAT Institute)

Yu-Tai Lee Unsteady Flows for Naval Applications .............................................................................................. 114 

Tsze C. Tai Low-Speed Aircraft Maneuvering Aerodynamics................................................................................. 195 

Nielsen Engineering and Research, Inc. 

William W. Bower Short Takeoff and Vertical Landing Aircraft Thermal/Acoustic Loads..................................................... 12 

(with McDonnell Douglas Research Laboratories) 

Robert E. Childs Turbulence-Model Development for Impinging Jet Flows ..................................................................... 26 

David Nixon Controlling Combustion in Free-Shear Layers ..................................................................................... 145 

North Carolina State University 

H. A. Hassan New Approach for Transitional-Flow Modeling .................................................................................... 77 

Northrop Corporation 

K. Appa Hydrodynamic Ram Structural Response Analyses.................................................................................. 5 

Richard D. Crouse Fighter Acoustic-Load Predictions......................................................................................................... 34 

Old Dominion University 

Oktay Baysal Aerodynamic Shape Optimization .......................................................................................................... 8 

A. 0. Demuren Complex Three-Dimensional Turbulent Flows ...................................................................................... 46 

Osama A. Kandil Simulation and Control of Slender Wing Rock ...................................................................................... 91 

(with NASA Lewis Research Center)

239 
Page 

M. G. Macaraeg Ignition and Structure of a Diffusion Flame with Vortex Interaction ....................................................121 

(with NASA Langley Research Center and [CASE) 

Oregon State University 

Jeffrey R. Barnes Nonlinear Baroclinic Instability ..............................................................................................................6 

(with NASA Ames Research Center) 

Pennsylvania State University 

James G. Brasseur Local Passive Scalar Dispersion in a Turbulent Boundary Layer............................................................13 

(with Georgia Institute of Technology) 

Budugur Lakshminarayana Steady and Unsteady Turbomachinery Flow Fields .............................................................................108 

(with General Motors Technical Center) 

Budugur Lakshminarayana Steady and Unsteady Viscous Flow ....................................................................................................109 

Charles L. Merkle Coupled Navier—Stokes Maxwell Analysis for Microwave Propulsion .................................................136 

Pratt & Whitney 

Joseph L. Garrett Numerical Simulation of Integrated Hypersonic Propulsion Flow Paths ................................................59 

(with General Electric Aircraft Engines) 

Princeton University 

George E. Karniadakis Turbulent Flow over Riblet-Mounted Surfaces ....................................................................................... 94 

G. E. Karniadakis Turbulent Flow over a Backward-Facing Step .......................................................................................95 

Purdue University 

John T. Batina Euler Flutter Analysis of a Complex Aircraft Configuration ......................................................................7 

(with NASA Langley Research Center) 

QUEST Integrated, Inc. 

Suresh Menon Active Control of Ramjet Combustion Instability ................................................................................ 135 

Rockwell International Science Center 

Marek K. Bleszynski Electromagnetic Scattering from Large Three-Dimensional Targets .......................................................11 

C. L. Chen Numerical Analysis of Three-Dimensional Separated Juncture Flows....................................................23 

(with NASA Ames Research Center) 

Yeu-chuan Hsia Full Navier—Stokes Analysis of a Three-Dimensional Scramjet Inlet ......................................................79 

(with Rockwell International, Rocketdyne Division) 

Vijaya Shankar Computational Fluid Dynamics Approach to Computational Electromagnetics ...................................1 77 

Rockwell International, North American Aircraft Division 

Ban H. Iran Radar Cross-Section Studies................................................................................................................199 

Jong H. Wang Hydrocarbon Scramjet Combustor Flows............................................................................................209 

240
Page 

Chung-Jin Woan Supersonic Laminar Flow Control Concept.........................................................................................216 

Chung-Jin Woan Nozzle Screech Phenomena ...............................................................................................................217 

David T. Yeh Steady and Unsteady Flows at High Angle of Attack ...........................................................................222 

Rockwell International, Rocketdyne Division 

Yeu-chuan Hsia Full Navier—Stokes Analysis of a Three-Dimensional Scramjet Inlet ......................................................79 

(with Rockwell International Science Center) 

S. J. Lin Turbopump Rotor/Stator Flows ...........................................................................................................119 

Sandia National Laboratory 

Brian J. Cantwell Simulation of a Time-Developing Incompressible Plane Wake .............................................................18 

(with Stanford University) 

Science Applications International Corporation 

Sanford M. Dash Simulation of Turbulent Jets for Aeroacoustic Applications ...................................................................38 

Sanford M. Dash Three-Dimensional High-Speed Plume/Propulsive Flow-Field Analysis ................................................39 

Scientific Research Associates, Inc. 

Richard C. Buggeln Hot-Gas-Manifold Flow Simulation ......................................................................................................15 

Thomas T. Huang Submarine-Appendage Design and Turbulence Modeling ....................................................................81 

(with Naval Surface Warfare Center; Jason Associates; and MCAT Institute) 

Southern Methodist University 

Fazle Hussain coherent-Structure Interactions with Turbulence ..................................................................................84 

(with University of Houston) 

Southwest Research Institute 

C. S. Lin Electron-Beam Injections from the Space Shuttle ................................................................................118 

(with Aurora Science, Inc.) 

Space Physics Research Institute 

Richard E. Young Volcanic Aerosol Clouds in the Stratosphere.......................................................................................223 

(with NASA Ames Research Center) 

Turbulent Boundary Layer with Heat Transfer.........................................................................................9 

Simulation of a Time-Developing Incompressible Plane Wake .............................................................18 

(with Sandia National Laboratory) 

Boost-Phase Detection Study ................................................................................................................21 

(with NASA Ames Research Center) 

Discrete Particle Simulation of Compressible Flow ...............................................................................54 

(with NASA Ames Research Center)

Stanford University 

Diane M. Bell 

Brian J. Cantwell 

Dean R. Chapman 

William Feiereisen

241
Page 

Sanjiva K. Lele Skewed Compressible Mixing Layers ..................................................................................................11 5 

Sanjiva K. Lele Turbulence Effects on Stagnation-Point Flow ......................................................................................116 

Parviz Moin Turbulent Flow Over Riblets ...............................................................................................................137 

(with NASA Ames Research Center) 

Parviz Moin Aerodynamic Sound Generation .........................................................................................................139 

(with NASA Ames Research Center) 

Parviz Moin Shock-WavFurbulent-Boundary-Layer Interaction ........................................................................... .140 

(with NASA Ames Research Center and Center for Turbulence Research) 

Michael M. Rogers Mixing and Reacting in Plane Mixing Layers.......................................................................................164 

(with NASA Ames Research Center and University of Wisconsin, Madison) 

Arnaud Trouvé Turbulent Premixed Combustion ........................................................................................................200 

(with NASA Ames Research Center) 

Stanford University School of Medicine 

Dennis Sullivan Microwave Hyperthermia Computer Modeling...................................................................................192 

Sterling Software 

G.S. Deiwert Self-Adaptive Grid Code Applied to Complex Three-Dimensional Flows ..............................................45 

(with NASA Ames Research Center and Eloret Institute) 

Sverdrup Technology, Inc. 

Suk C. Kim Low-Thrust Chemical Rockets...............................................................................................................98 

(with NASA Lewis Research Center) 

Kevin R. Kirtley Multistage Turbomachinery Flows ......................................................................................................100 

(with NASA Lewis Research Center) 

H.T. Lai Hydrogen—Air Reacting Flow Fields in Drag Reduction External Combustion .....................................107 

Sang Soo Lee Resonant-Triad Interaction in an Adverse Pressure-Gradient Boundary Layer......................................113 

(with NASA Lewis Research Center) 

David L. Rigby Leading-Edge Heat Transfer in a Flow with Spanwise Variations.........................................................161 

Eric T. Stewart Space Station Freedom Internal Flow Analysis ....................................................................................188 

(with NASA Marshall Space Flight Center) 

Teledyne Brown Engineering 

R. P. Roger Jet Interaction Aero-Optic Effects on Hypersonic Interceptors .............................................................163 

TRW, Ballistic Missiles Division 

M. I. Cruz Flexible Aerobrake Aerothermodynamic Study .....................................................................................35 

(with TRW, Federal Systems Division) 

242
Page 

TRW, Federal Systems Division 

M. I. Cruz Flexible Aerobrake Aerothermodynamic Study .....................................................................................35 

(with TRW, Ballistic Missiles Division) 

U.S. Army Aeroflightdynamics Directorate, AVSCOM 

Frank Caradonna Prediction of Advanced Rotor Performance ..........................................................................................19 

(with NASA Ames Research Center) 

W. J. McCroskey Aerodynamics and Acoustics of Rotorcraft..........................................................................................133 

(with NASA Ames Research Center) 

W. J. McCroskey High-Performance Rotor-Blade Tips....................................................................................................134 

(with NASA Ames Research Center) 

U.S. Army Propulsion Directorate, AVSCOM

Russell G. DeAnna Transition over a Rough Surface ...........................................................................................................42 

(with Case Western Reserve University) 

U.S. Army Missile Command 

Bill J. Walker Tactical Missile Aero-Propulsion Interaction ....................................................................................... 207 

United Technologies Research Center 

Roger L. Davis Turbine "Hot Spot" Alleviation using Film Cooling ...............................................................................41 

T. Alan Egolf Rotary-Wing Airload Performance Prediction .......................................................................................51 

Jon S. Mounts Transonic-Shock/Boundary-Layer Interaction to Alleviate Separation.................................................. 143 

Greg D. Power Three-Dimensional Scramjet Combustors ........................................................................................... 156 

University College, Wales 

Rajiv Thareja Unstructured Multigrid Euler Solver.................................................................................................... 196 

(with Lockheed Engineering and Sciences Company and Imperial College) 

University of Arizona 

Joseph W. Humphrey Hypersonic Scramjet and Detonation Flows ......................................................................................... 83 

(with Advanced Projects Research, Inc. and NASA Langley Research Center) 

Universityof California, Davis 

W. Kollmann Turbulent Boundary Layers with Suction ............................................................................................. 101 

University of California, Los Angeles - 

David Halpern Satellite Data Assimilation and Ocean General Circulation Models ...................................................... 70 

(with Jet Propulsion Laboratory) 

David Halpern Atmospheric General Circulation Model Sensitivity to Sea Surface Temperature Fields......................... 71 

(with Jet Propulsion Laboratory)

243 
Page 

University of California, Santa Cruz 

Jeffrey N. Cuzzi Particle-Gas Dynamics in the Protoplanetary Nebula............................................................................36 

(with NASA Ames Research Center and A.T.M., Inc.) 

University of Chicago 

Bruce F. Smith Numerical Experiments in the Formation and Evolution of Galaxies ...................................................182 

(with NASA Ames Research Center) 

University of Colorado, Boulder 

Chuen-Yen Chow Compressible Taylor—Couette Flow.......................................................................................................28 

(with NASA Lewis Research Center) 

Karl E. Gustafson Airfoil Lift and Thrust Generation in Hover Mode .................................................................................68 

Gary E. Thomas Multi-Dimensional Simulations of Noctilucent Cloud Formation ........................................................197 

(with NASA Ames Research Center) 

University of Houston 

Fazle Hussain Coherent-Structure Interactions with Turbulence ..................................................................................84 

(with Southern Methodist University) 

Fazle Hussain Chemically Reacting Free-Shear Flows .................................................................................................85 

(with Naval Research Laboratory) 

University of Illinois, Urbana/Champaign 

James D. Holdeman Hot-Gas Ingestion by a Short Takeoff and Vertical Landing Aircraft in Ground Proximity .....................78 

(with NASA Lewis Research Center) 

J. P. Wolfe Advanced Computational Materials Research .....................................................................................218 

University of Iowa 

V. C. Patel Hydrodynamics of Self-Propelled Bodies ............................................................................................150 

Fred Stern - Free-Surface Effects on Boundary Layers and Wakes...........................................................................187 

University of Maryland 

T. A. Zang Transition Modeling............................................................................................................................225 

(with NASA Langley Research Center and University of Texas) 

University of Missouri, Rolla 

R. Clayton Rogers Hypervelocity Mixing and Combustion in Pulse Facility Flows ............................................................ 165 

(with NASA Langley Research Center and Analytical Services and Materials, Inc.) 

University of Texas 

T. A. Zang Transition Modeling ............................................................................................................................ 225

(with NASA Langley Research Center and University of Maryland) 

244
Page 

University of Virginia 

Ahmed K. Noor Thermomechanical Buckling and Post-Buckling of Multilayered Composite Panels 

withCutouts ....................................................................................................................................... 146 

(with NASA Langley Research Center) 

University of Wisconsin, Madison 

Michael M. Rogers Mixing and Reacting in Plane Mixing Layers....................................................................................... 164 

(with NASA Ames Research Center and Stanford University) 

VIGYAN, Inc. 

Jassim A. Al-Saadi Transonic Reynolds Number Effects on Aircraft Configurations............................................................... 4 

(with NASA Langley Research Center) 

H. C. Chen Euler Analysis of Turboprop and Turbofan Integration .......................................................................... 24 

(with The Boeing Company and NASA Langley Research Center) 

Virendra K. Dogra Rarefied Hypersonic Condition Wake Structures................................................................................... 47 

(with NASA Langley Research Center) 

H. Harris Hamilton Analysis of Radiative Structure in Aerobrake Shock Layers.................................................................... 73 

(with NASA Langley Research Center) 

Julius E. Harris Swept-Wing Leading-Edge Transition.................................................................................................... 75 

(with NASA Langley Research Center) 

Kenneth M. Jones Analysis of High-Speed Civil Transport Configurations ......................................................................... 88 

(with NASA Langley Research Center and Lockheed Engineering and Sciences Company) 

Richard G. Wilmoth Shock Interactions in Hypersonic Rarefied Flows................................................................................ 215 

(with NASA Langley Research Center) 

Virginia Polytechnic Institute and State University 

Robert W. Walters Finite-Rate Chemistry Algorithms........................................................................................................ 208 

Vista Research, Inc. 

Charles L. Rino Scattering from Ocean Surfaces and Near-Surface Objects ................................................................. 162

VRA, Inc. 

Clark H. Lewis Three Dimensional Parabolized and Full Navier—Stokes Techniques ..................................................117 

WL/FIMM, Wright Patterson AFB 

Datta Gaitonde Turbulent Flow Past a Complete Hypersonic Reentry Configuration .....................................................58 

Raymond Gordnier Unsteady Delta-Wing Flow...................................................................................................................62 

Balu Sekar Viscous Non-Reacting Flows in High-Speed Combustors....................................................................176 

U.S. GOVERNMENT PRINTING OFFICE: 1993-785-088 245 
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