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On the Cover: The cover images are from three NASTechnicaI

Summaries: Forrester T. Johnson's summary, "MultipointAerodynamic Design," is on page 10; John Rakiewicz's summary, "Titan

Entry Flow Field of the Cassini-Huygens Probe," is on page 27;

and Richard E. Young's summary, "Simulation of Volcanic Aerosol

Clouds," is on page 99.

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 that form the supercomputer solutions to the studies contained in this report and on the

cover would be virtually incomprehensible.
NAS

Technical

Summaries

March 1993—

February 1994

NASA Ames Research Center
Number of Cray Projects by Discipline

n Administrative q Hypersonics

Aerodynamics i Propulsion

Rotocraft

. Turbulence

Other

Astrophysics

Atmospheric Sciences

Chemistry

• Computer Science

Electrical Engineering

Eleclromagnetics

Fluid Mechanics

• Hydrodynamics

• Hydrogeo logy

Life Sciences

• Material Sciences

Structural Mechanic

Other

Number of Parallel Projects by Discipline

zs n Administrative q Hypersonics

X

21

7 X2111

r

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16

15 [

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n Turbulence

FI Other

• Astrophysics

Atmospheric Sciences

Chemistry

• Computer Science

Electromagnetics

Fluid Mechanics

Hydrodynamics

Life Sciences

Material Sciences

Other Fluids

Physics

Structural Mechanics

Other

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 computing resources to solve the most

demanding computational fluid dynamics problems and serves as

a pathfinder in integrating leading-edge supercomputing technologies. The benefits of numerical simulation of aeroscience problems are tremendous savings in both time and money—savings

that are critical to the viability and competitive edge of the

American aircraft industry.

During the 1993-94 Operational Year, NAS high-speed processor

resources included a Cray Y-MP and a Cray C-90. The Cray Y-MP

had eight processors, 256 megawords of central memory, a peak

speed of 4 billion floating point operations per second (GFLOPS)

and a sustained speed of 1 GFLOPS. The Cray C-90 had 16 processors, 1,024 megawords of main memory, a peak speed of

16 GFLOPS and a sustained speed of 6 GFLOPS.

NAS Parallel resources included an Intel iPSC/860, an Intel

Paragon, and a Thinking Machine CM-5. The Intel iPSC/860 had

128 nodes with 8 megabytes of memory per node and a peak

speed of 5 GFLOPS. The Intel Paragon had 208 nodes with

32 megabytes of memory per node, a peak speed of 15 GFLOPS,

Number of Cray Projects by Category

120

118 v

rn

100

M 87 88

rn

80

71

t0

j 60 51

c0

p 40

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O 3 3 1

ay A °° P . ate yZa A °°

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sac a^5 `a

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2

Oa

Number of Parallel Projects by Category

and a sustained speed of 1.2 GFLOPS. The Thinking Machine

CM-5 had 128 nodes with 32 megabytes of memory per node, a

peak speed of 16 GFLOPS and a sustained speed of 1.4 GFLOPS.

The 1993-94 Operational Year concluded with 448 high-speed

processor projects and 95 parallel projects representing NASA,

the Department of Defense, other Government agencies, private

industry, and universities. This document provides a glimpse at

some of the significant scientific results for the year.

The content of the NAS Technical Summaries for the 1993-94

Operational Year contains selected scientific results computed

using NAS high-speed processor (HSP) and parallel (PAR)

resources. An icon at the top of each technical summary page

indicates the resource used— 'c' represents HSP, and represents PAR. This edition is organized by research disciplines and

alphabetically within each discipline.

An online Mosaic version of this book is accessible from the NAS

home page at URL http://www.nas.nasa.gov/home.html . Additional copies of the NAS Technical Summaries Report for this year

and for previous years are available on request. Contact the

NAS Documentation Center at NASA Ames Research Center,

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

to doc-center@nas.nasa.gov or telephone (415) 604-4632.
Page intentionally left blank 
Table of Contents

Principal

Investigator NAS Summary Page

Aeronautics: Basic Research

Ramesh K. Agarwal Unstructured Grid Finite-Element Solutions ................................................................ 3

Co-investigators: David W. Halt and David L. Marcum

McDonnell Douglas Corporation/Mississippi State University

Timothy J. Barth Parallel Implementation ofan Unstructured Mesh Navier—Stokes Solver ..................... 4

Co-investigator: Samuel W. Linton

NASA Ames Research Center/Sterling Software Systems

Farhad Ghaffari Applied Computational Fluid Dynamics ..................................................................... 5

Co-investigators: Brent L. Bates and James M. Luckring

NASA Langley Research CenterNiGYAN, Inc.

Philip B. Gingrich Design-by-Optimization Method ................................................................................ 6

Rockwell International, North American Aircraft Division

Fernando F. Grinstein Simulation of Spatially Evolving Jets ............................................................................ 7

Naval Research Laboratory

Guru P. Guruswamy Aeroelasticity of Wing—Body—Control Configurations .................................................. 8

Co-investigators: Chansup Byun, Eugene Tu, and Shigeru Obayashi

NASA Ames Research Center/MCAT Institute

Guru P. Guruswamy Wing—Body Aeroelasticity ........................................................................................... 9

Co-investigator: Chansup Byun

NASA Ames Research Center/MCAT Institute

Forrester T. Johnson Multipoint Aerodynamic Design ................................................................................. 10

Co-investigators: Michael B. Bieterman, Craig L. Hilmes, William P. Huffman,

Robin G. Melvin, and David P. Young

The Boeing Company

StevenB. Kern Vortex-Flow Control ................................................................................................... 11

Naval Air Warfare Center

Sanjiva K. Lele Skewed Compressible Mixing Layers .......................................................................... 12

Co-investigator: Ganyu Lu

Stanford University

Frank Marconi Comparison of TLNS3D Computations with Test Data ................................................. 13

Co-investigator: M. Siclari

Grumman Corporate Research Center

Dimitri J. Mavriplis Navier—Stokes Computations on Unstructured Grids ................................................... 14

ICASF

Pradeep Raj CFL3D Code Evaluation .............................................................................................. 15

Co-investigators: Tom Kinard and Richard Semmes

Lockheed Aeronautical Systems Company

v
Principal

Investigator NAS Summary Page

Yehia Rizk Prediction of the Unsteady F-18 Flow Field ................................................................ 16

Co-investigator: Scott Murman

NASA Ames Research Center

David M. Schuster Analysis of High-Speed Civil Transport Transient Response ......................................... 17

Lockheed Engineering and Sciences Company

James L. Thomas Complete F-18 Hybrid Solution .................................................................................. 18

Co-investigators: Robert T. Biedron, Christopher L. Rumsey, W. Kyle Anderson,

and Daryl L. Bonhaus

NASA Langley Research Center

Chung-Jin Woan Flow Simulation about Attack Aircraft ........................................................................ 19

Co-investigator: John D. Duino

Rockwell International, North American Aircraft Division

David T. Yeh Multiple-Body Aerodynamics ..................................................................................... 20

Rockwell International, North American Aircraft Division

Aeronautics: Hypersonics, Subsonics, Supersonics, Transonics

Hypersonics

Andrew Anagnost Time-Accurate Simulation of Hypervelocity Wakes .................................................... 23

Stanford University

Neal M. Chaderjian Effects of Roll Angle on Vortex Breakdown ................................................................. 24

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

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

Robert R. Chamberlain Jet Interaction Effects on Aerodynamic Control of Missiles .......................................... 25

Co-investigators: Anthony Dang and Ken Xiques

Adaptive Research Corporation

Hassan A. Hassan Transition Model for High-Speed Flow ....................................................................... 26

Co-investigator: Eric S. Warren

North Carolina State University

John Rakiewicz Titan Entry Flow Field of the Cassini—Huygens Probe .................................................. 27

Co-investigators: Scott Ward, Darren Fricker, and David Wyn-Roberts

Jet Propulsion Laboratory/McDonnell Douglas Aerospace/European Space Agency

S. H. Konrad Zhu Simulation of Rarefied )et Interaction ......................................................................... 28

Co-investigator: Leonardo Dagum

Rockwell International, Rocketdyne Division/NASA Ames Research Center

Vi
Principal

Investigator NAS Summary Page

Subsonic High-Lift Analysis ..........................................................................

Co-investigators: Victor Lessard, Simha Dodbele, Kevin Kjerstad, Steve Yaros,

and Mamad Takallu

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

Sciences Company

Investigation of Forebody Flow Control .........................................................

Co-investigators: Ken Gee and Roxana Agosta

NASA Ames Research Center/MCAT Institute

Evaluation of Navier—Stokes Codes for Transport Configuration Analysis .......

Co-investigators: H. V. Cao and S. R. Allmaras

Boeing Commercial Airplane Group

Supersonic Laminar Flow Control Experiment

Co-investigator: Chandra S. Vemuru

NASA Lewis Research Center/AS&M, Inc.

Unstructured Grid/Flow Solver Calibration ..................................................

Co-investigators: Tracy J. Welterlen, Elizabeth A. Fuchs, and Doug Howlett

Lockheed Engineering and Sciences Company

Laminar Flow Supersonic Wind Tunnel ........................................................

Co-investigator: G. H. Klopfer

NASA Ames Research Center/MCAT Institute

Subsonics

Kenneth M. Jones

Lewis B. Schiff

N. J. Yu

Supersonics

Michael C. Fischer

Steve L. Karman, Jr.

E. Tu

29

30

31

32

33

34

Transonics

Jassim A. AI-Saadi

Christopher A. Atwood

Transonic Flows Around Complex Geometries .........................

Co-investigators: Peter A. Cavallo and William D. Smith

NASA Langley Research Center/George Washington University

Simulation of Coupled Fluid and Flightdynamic Systems .........

Overset Methods, Inc.

............................. 35

36

Raymond R. Cosner F/A-18 Aerodynamic Assessment .............................................................. 37

Co-investigators: Robert H. Bush, William W. Romer, and Paul G. Willhite

McDonnell Douglas Corporation

Jerry E. Deese Launch Vehicle Flow-Field Simulation ...................................................... 38

Co-investigator: Ramesh K. Agarwal

McDonnell Douglas Corporation

Daniel F. Dominik Space Shuttle Flow Fields with Plume Simulation ..................................... 39

Co-investigators: K. Rajagopal, C. 011ing, K. Mani, S. Vuong, J. Wisneski,

G. Hock and J. Sikora

Rockwell International, Space Systems Division

V11
Principal

Investigator NAS Summary Page

W. Kelly Londenberg Reynolds Number Effects on a Delta Wing ......................................................... 40

ViGYAN, Inc.

Christopher L. Reed High-Angle-of-Attack Aerodynamics .................................................................. 41

Lockheed Engineering and Sciences Company

Arvin Shmilovich Computational Studies of Advanced Wings at High Incidence ............................ 42

Co-investigators: K. C. Chang and R. A. Pelkman

McDonnell Douglas Aerospace

Aeronautics: Propulsion

Alan B. Cain Simulation of Supersonic Jet Screech ........................................................................

Co-investigators: William W. Bower and Yutaka Ikeda

McDonnell Douglas Corporation

Rodrick V. Chima Multiblock Solver for Turbomachinery Flows ............................................................

NASA Lewis Research Center

Wei J. Chyu Airframe/Inlet Aerodynamics ....................................................................................

Co-investigator: Tom I-P. Shih

NASA Ames Research Center/Carnegie Mellon University

Frederik J. de Jong Turbine Blade Tip Clearance Flows ...........................................................................

Co-investigator: Tony Chan

Scientific Research Associates, Inc.

W. W. Follett Vortex Induced Mixing Behind Ramp Injectors .........................................................

Co-investigators: J. V. Madison and S. Palaniswamy

Rockwell International, Rocketdyne Division/Rockwell International Science Center

45

46

47

48

49

Karen L. Gundy-Burlet Unsteady Turbomachinery Computations .............................................

NASA Ames Research Center

Chunill Hah Unsteady Three-Dimensional Flow in a Transonic Compressor .............

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

NASA Lewis Research Center/ICOMP/Wright Patterson Air Force Base

Edward J. Hall Transonic Core Compressor Rotor/Stator Interaction ............................

Co-investigators: Kurt F. Weber, K. V. Rao, and R. A. Delaney

Allison Engine Company

Edward J. Hall Advanced Ducted Propfan Analysis Code Certification ........................

Co-investigator: G. Scott McNulty

Allison Engine Company

Edward J. Hall Fan Rotor/Endwall Treatment Interaction .............................................

Co-investigators: Andrew J. Crook and Kathy P. Nardini

Allison Engine Company

...................... 50

...................... 51

.......... I ... ........ 52

...................... 53

...................... 54

viii
Principal

Investigator NAS Summary Page

Buclugur Lakshminarayana Three-Dimensional Steady and Unsteady Turbulent Flows in Turbomachines .............. 55

Co-investigators: S. Fan, J-F. Gallardo, Y H. Ho, J. Luo, and W. S. Yu

Pennsylvania State University

C. L. Merkle Computational Propulsion Applications ...................................................................... 56

Co-investigators: S. Venkateswaran, J-Z. Feng, H-H. Tsuei, and J. M. Grenda

Pennsylvania State University

Edward J. Reske Complex Three-Dimensional Flows in the Advanced Solid Rocket Motor .................... 57

Co-investigators: Lee A. Kania and James W. Zuercher

NASA Marshall Space Flight Center

William C. Rose High-Speed Inlet Design and Analysis ........................................................................ 58

Rose Engineering and Research, Inc.

Balu Sekar Three-Dimensional Mixing Flows in High-Speed Combustors ..................................... 59

Wright Patterson Air Force Base

Toni I-P. Shih Flow in Turbine-Blade Cooling Passages ...................................................................... 60

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

Carnegie Mellon University/NASA Lewis Research Center

Arvin Shmilovich Analysis of Nacelle/Pylon/Wing Installations .............................................................. 61

Co-investigator: Lie-Mine Gea

McDonnell Douglas Aerospace

Clifford E. Smith Jets-in-Cross-Flow Mixing ........................................................................................... 62

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

CFD Research Corporation/NASA Lewis Research Center

Robert H. Sues Parallel Multidisciplinary Stochastic Optimization ...................................................... 63

Co-investigator: Graham S. Rhodes

Applied Research Associates, Inc.

R. C. Swanson Simulation of Scramjet Flow Fields ............................................................................. 64

Co-investigators: J. Korte and E. Turkel

NASA Langley Research Center/[CASE

long H. Wang Hydrocarbon Scramjet Combustor .............................................................................. 65

Rockwell International, North American Aircraft Division

Scott Ward Reusable Launch Vehicle Flow Simulations ................................................................. 66

Co-investigator: Darren Fricker

McDonnell Douglas Aerospace

Kurt F. Weber Chimera Domain Decomposition Applied to Turbomachinery Flow ............................ 67

Co-investigator: Dale W. Thoe

General Motors Corporation, Allison Gas Turbine Division

ix
Principal

Investigator NAS Summary Page

Shaye Yungster National Aero-Space Plane Nozzles with External Burning ......................................... 68

Co-investigator: Charles J. Trefny

ICOMP/NASA Lewis Research Center

Aeronautics: Rotorcraft

W. J. McCroskey Calculations of High Performance Rotorcraft .............................................................. 71

Co-investigators: E. P. N. Duque, S. Ko, J. Ahmad, and A. C. B. Dimanlig

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

Sterling Software Systems/University of California, Davis

Roger C. Strawn Solution-Adaptive Computations of Helicopter Acoustics ........................................... 72

Co-investigators: Rupak Biswas and Michael Garceau

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

RIACS/Stanford University

Aeronautics: Turbulence

David E. Ashpis Transition in a Highly Disturbed Environment ............................................................ 75

Co-investigator: Philippe R. Spalart

NASA Lewis Research Center/Boeing Commercial Airplane Group

Brian J. Cantwell Incompressible Plane Wake Computation Comparisons ............................................. 76

Co-investigators: Nagi N. Mansour, Rolf Sondergaard, and Eric Monsen

Stanford University/NASA Ames Research Center

Haecheon Choi Numerical Simulation of Riblets ................................................................................. 77

Co-investigators: Parviz Moin and John Kim

Stanford University/NASA Ames Research Center

George E. Karniadakis Direct and Large-Eddy Simulation of Wake Flows ...................................................... 78

Co-investigators: D. Newman and R. D. Henderson

Brown University

George E. Karniadakis Simulation of Complex-Geometry Wall-Bounded Flows ............................................. 79

Co-investigators: Catherine H. Crawford and Ronald D. Henderson

Brown University

Linda D. Kral Turbulence Modeling for Three-Dimensional Flow Fields ........................................... 80

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

McDonnell Douglas Corporation

Thomas S. Lund Large-Eddy Simulation of Jet Combustors ................................................................... 81

Co-investigators: Parviz Moin and Knut Akselvoll

Stanford University/NASA Ames Research Center

Nateri K. Madavan Boundary-Layer Transition on a Heated Flat Plate ....................................................... 82

MCAT Institute

X
Principal

Investigator NAS Summary Page

Parviz Moin Direct Numerical Simulation of Shock/Turbulence Interaction .................................... 83

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

Stanford University

Parviz Moin Large-Eddy Simulation of Separated External Flows .................................................... 84

Co-investigators: Thomas S. Lund, Patrick Beaudan, and Haecheon Choi

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

Parviz Moin Large-Eddy Simulation of Complex Flows ................................................................... 85

Co-investigators: Thomas S. Lund and Hans J. Kaltenbach

Stanford University/NASA Ames Research Center

Steven A. Orszag Unsteady Turbulence Model Simulations .................................................................... 86

Co-investigator: William S. Flannery

Cambridge Hydrodynamics, Inc.

Richard H. Pletcher Three-Dimensional Liquid Sloshing Flow Simulation .................................................. 87

Co-investigators: S. Babu, F. J. Kelecy, and W-P. Wang

Iowa State University

Man Mohan Rai Direct Simulation of Turbulent/Transitional Airfoil Flow .............................................. 88

NASA Langley Research Center

Michael M. Rogers Self-Similar Turbulent Plane Wakes ............................................................................. 89

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

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

Bart A. Singer Origins of Near-Wall Turbulence ................................................................................ 90

Co-investigator: Ronald D. Joslin

High Technology Corporation/NASA Langley Research Center

General: Aeroacoustics, Astronautics, Astronomy, Atmospheric Science,

Chemistry, Computer Science, Electromagnetics, Fluid Mechanics, Life Science,

Magnetospheric Physics, Reactive Flow, Space Science

Aeroacoustic

Lyle N. Long Hybrid Massively Parallel Aeroacoustics Scheme ..................................................... 93

Pennsylvania State University

Astronautics

Creon Levit Evolution of Planetary Rings ................ ... 94

NASA Ames Research Center

Astronomy

James Caristi Simulating the Cosmic X-Ray Experiment ................................................................... 95

Co-investigators: Brian Ramsey, Martin Weisskopf, and Jeff Youngen

Valparaiso University/NASA Marshall Space Flight Center

xi
Principal

Investigator NAS Summary Page

Bruce F. Smith Studies of Galaxy Formation and Evolution ................................................................ 96

Co-investigators: Richard A. Gerber, Richard H. Miller, and

Thomas Y. Steiman-Cameron

NASA Ames Research Center/University of Chicago

Atmospheric Science

Christian L. Keppenne Intraseasonal Atmospheric Oscillations ...................................................................... 97

Jet Propulsion Laboratory

James B. Pollack Dynamics of the Martian Atmosphere ........................................................................ 98

Co-investigator: Robert Haberle

NASA Ames Research Center

Richard E. Young Simulation of Volcanic Aerosol Clouds ....................................................................... 99

Co-investigators: O. B. Toon and H. Houben

NASA Ames Research Center

Chemistry

Sang-Wook Kim Chemically Reacting Compressible Shear Layer ......................................................... 100

University of Toledo

Computer Science

Jerry Yan Automated Instrumentation and Monitoring System ................................................... 101

Co-investigators: Pankaj Mehra, Sekhar Sarukkai, Cathy Schulbach,

Melisa Schmidt, and Brian Vanvoorst

Recom Technologies/NASA Ames Research Center

Electromagnetics

Ramesh K. Agarwal Computational Fluid Dynamics Solution for Maxwell's Equations .............................. 102

Co-investigators: Dau-Sing Wang and Mark R. Axe

McDonnell Douglas Corporation

Vijaya Shankar Computational Fluid Dynamics Approach to Computational Electromagnetics ........... 103

Co-investigators: William Hall, Chris Rowell, and Alireza Mohammadian

Rockwell International Science Center

John L. Volakis Electromagnetic Scattering by Airborne Structures ...................................................... 104

Co-investigator: Arindam Chatterjee

University of Michigan

Alex C. Woo Computational Radar Cross Sections .......................................................................... 105

Co-investigators: Michael Simon and Michael Schuh

NASA Ames Research Center/Sterling Software Systems

Fluid Mechanics

Joseph D. Baum Arbitrary Lagrangian—Eulerian Three-Dimensional Simulations ................................... 106

Co-investigator: Rainald Lohner

xii
Science Applications International Corporation/George Mason University

Principal

Investigator NAS Summary Page

Fred Stern Unsteady Viscous Marine Propulsor Hydrodynamics .................................................. 107

Co-investigators: E. Paterson and B. Chen

University of Iowa

Fred Stern Free-Surface Effects on Boundary Layers and Wakes ................................................... 108

Co-investigators: Y. Tahara and J. Choi

University of Iowa

Matthew E. Thomas Impeller Computational Fluid Dynamics Modeling ..................................................... 109

Co-investigators: Mahesh M. Athavale and Mark L. Ratcliff

CFD Research Corporation

Life Science

Jeffrey R. Hammersley Modeling of Pulmonary Fluid Dynamics ..................................................................... 110

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

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

Andrew Pohorille Structure and Functions of Membranes ....................................................................... 111

Co-investigator: Michael A. Wilson

University of California, San Francisco

Muriel D. Ross Finite Volume Neuronal Modeling .............................................................................. 112

Co-investigators: David G. Doshay, Samuel W. Linton, and Timothy J. Barth

NASA Ames Research Center/Sterling Software Systems

Magnetospheric Physics

Stephen H. Brecht Solar Wind Interaction with Mars ................................................................................ 113

Co-investigator: John R. Ferrante

Berkeley Research Associates

Reactive Flow

K. Kailasanath Burner-Stabilized Flames in Microgravity .................................................................... 114

Co-investigator: G. Patnaik

Naval Research Laboratory/Berkeley Research Associates

Space Science

Jeffrey N. Cuzzi Protoplanetary Nebula Particle-Gas Dynamics ............................................................ 115

Co-investigators: Anthony R. Dobrovolskis, Robert C. Hogan,

Joelle M. Champney, and Jennifer M. Dacles-Mariani

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

xiii
=TQ its

^Ti 1►7 lA \ ^ i
Page intentionally left blank 
Unstructured Grid Finite-Element Solutions

Ramesh K. Agarwal, Principal Investigator

Co-investigators: David W. Halt and David L. Marcum

McDonnell Douglas Corporation/Mississippi State University

I'C^

Research Objective

To investigate the viability of unstructured grids in computing

three-dimensional finite-element Euler and Navier-Stokes steady

and unsteady flows.

Approach

The Euler or 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 Lax-Wendroff scheme or an explicit multistage

Runge-Kutta scheme. Turbulence effects can be modeled using

a Baldwin-Lomax, Baldwin-Barth, or two-equation k-e model.

The incremental-point-insertion Delaunay method is used for grid

generation and solution adaptation.

Accomplishment Description

An unstructured grid was generated for the Delta launch vehicle

with nine attached boosters. The objective was to allow every

third booster to separate symmetrically from the main body after

the booster fuel was burned. The grid for a 120-degree section

was generated with 88,325 grid points and 460,722 tetrahedra.

The initial steady state solution is computed at a free-stream

Mach number of 2.313 and 0 degrees angle of attack. The force

and moments were integrated on each free booster with initial

trajectories set and the unsteady flow solution was initiated. The

unstructured grid was distorted as the free boosters separated

from the main body. A continual cycle of unsteady flow analysis,

integration of body forces and moments, calculation of the trajectory, and movement of the grid was repeated until the level of grid

distortion required maintenance. This case required 15 megawords of memory and 9 Cray Y-MP microseconds per iteration for

each node.

Significance

An unsteady Euler analysis of the Delta launch vehicle with

separating boosters was demonstrated.

Future Plans

To calculate the viscous unsteady flow fields about multiple

bodies in relative motion, such as a launch vehicle with separating boosters or an aircraft releasing a store.

Keywords

Computational fluid dynamics, Viscous flow, Finite element

method

Publication

Marcum, D. L.; and Weatherill, N. P.: Finite Element Calculations

of Inviscid and Viscous Flow Fields about Launch Vehicle Configurations. Eighth International Conference on Finite Elements in

Fluids, Barcelona, Spain, Sept. 1993.

Surface grids for a launch vehicle with nine boosters.

r+

Density contours for a launch vehicle with nine boosters.

Aeronautics/Basic Research
Parallel Implementation of an Unstructured Mesh Navier—Stokes Solver

Timothy J. Barth, Principal Investigator

Co-investigator: Samuel W. Linton

NASA Ames Research Center/Sterling Software Systems

Research Objective

To develop a parallel implementation of an existing unstructured

mesh Navier—Stokes solver.

Approach

A parallel implementation of an existing Navier—Stokes solver for

the Cray computer was modified for the CM-5. The code utilizes

a static mesh partitioning combined with a message passing protocol. In the baseline Cray code, the Navier—Stokes equations are

discretized on arbitrary meshes using a finite-volume approach

with upwind dissipation added to stabilize the scheme. A singleequation turbulence model is included to simulate the effects of

turbulence on the mean flow. The algorithm exhibits secondorder spatial accuracy on sufficiently smooth meshes. An implicit

time-stepping strategy is used that approaches Newton's method

for large time steps. This strategy requires the solution of a

sequence of large, sparse matrix equations that are solved

using a generalized minimum residual method with incomplete

lower—upper preconditioning. The parallel implementation

retains the discretization and Newton-like properties of the

uniprocessor code.

Accomplishment Description

The figure shows typical results for subsonic flow over a high-lift

airplane configuration. The subdivisions induced by the volume

partitioning are shown on the body surface and cross plane.

Significance

The use of unstructured meshes permits the calculation of

aerodynamic flow over complex geometries. Moreover, the

improved efficiency and salability of the parallel algorithm permits the calculation of these flows using modest resources.

Future Plans

A shape optimization capability will be incorporated into the

flow solver using an adjoint equation approach.

Keywords

Unstructured mesh, Navier—Stokes

Mach number Gouraud shading of surface and cross plane for subsonic flow calculation.

Aeronautics/Basic Research
Applied Computational Fluid Dynamics

Farhad Ghaffari, Principal Investigator

Co-investigators: Brent L. Bates and James M. Luckring

NASA Langley Research Center/ViGYAN, Inc.

Research Objective

To assess the vortical-flow prediction capability of the state-ofthe-art unstructured-grid methodology for a complex generic

fighter wind tunnel model over a wide range of flow conditions.

Approach

The analytical description of a generic fighter wind tunnel model,

known as modular transonic vortex interaction (MTVI), was used

to generate the geometrical database from which the computational surface grid definition was derived for the entire geometry.

The MTV] configuration incorporated a 60-degree sharp-edged

cropped delta wing with segmented leading-edge flaps, chineshaped fuselage, and twin vertical tails positioned on the aftinboard region of the wings. The complete geometry was represented by about 28,000 surface triangles. The flow-field grid,

generated using an advancing front method, consisted of about

825,000 tetrahedral cells. The grid spacing, distribution, and the

extent of the far-field boundaries were consistent with an earlier

calibration study performed on the isolated MTV] fuselage configuration. Since the subject vortical flow (emanating from sharp

edges) is presumed to be mostly governed by inviscid flow phenomena, the present computational effort was performed based

on an Euler formulation.

Accomplishment Description

Computational results were obtained for the MTVI configuration

for a wide range of flow conditions using the flow solver USM3D.

A typical solution obtained at 22 degrees angle of attack and a

Mach number of 0.4 and superimposed over the computational

surface grid is shown in the figure. The figure illustrates the totalpressure contours at several cross-flow planes along with the particle tracings within the core region for both primary vortex systems (one emanating from the chine forebody and the other from

the wing leading-edge flaps). The figure clearly shows the geometric complexity, in particular the deflected wing leading-edge

flaps (two inboard segments deflected at 30 degrees), and the predicted vortical-flow structures. This solution, initiated from the

free-stream flow conditions, was advanced for 2,000 iterations

during which the total residuals dropped about 1.5 orders of magnitude. The solution required approximately 150 megawords

of memory and used about 10 hours of computational time on

Cray Y-MP. Additional analysis has shown that the computed surface pressure coefficients, forces, and moments agree reasonably

well with the measured experimental wind tunnel data.

Significance

The present unstructured-grid methodology provides reasonable

flow predictions. The versatility of the methodology in grid generation for complex geometries makes it an efficient technique for

routine analysis applications.

Future Plans

Computations will be extended to parametrically investigate the

aerodynamic effects for various geometrical modifications such

as the fuselage chine shapes, tail arrangements, and wing

leading-edge flap deflections.

Keywords

Euler computations, Fighter configuration, High angle of attack,

Vortex flows

Publication

Ghaffari, F.: On the Vortical-Flow Prediction Capability of an

Unstructured-Grid Euler Solver. AIAA Paper 94-0163, Jan. 1994.

L00

0.90

0.80

- 0.70

Total-pressure contours in various cross-flow planes and vortex core particle traces; a = 22 degrees and free-stream Mach number = 0.4.

Aeronautics/Basic Research
Design-by-Optimization Method

Philip B. Gingrich, Principal Investigator

Rockwell International, North American Aircraft Division

r^

Research Objective

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

optimization design method applicable to complete aircraft

configurations.

Approach

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

solver is coupled with a gradient-based, constrained optimization

technique. Geometry perturbation functions, with an associated

design variable, operate on the surface grid. The optimizer determines the design variable set, which minimizes a composite function formed with the objectives and constraints. The objective

may be any aerodynamic variable such as drag or lift-to-drag

(UD) ratio. Constraint functions consist of both aerodynamic and

geometric variables. To simulate an inverse procedure, the objective is the root-mean-square difference between current and specified target pressure distribution.

Accomplishment Description

A 3-D design-by-optimization method for application to

complete aircraft configurations is expanded to increase geometric flexibility and to reduce computational time. The method was

originally developed as an inverse design tool to derive a shape

for a specified pressure distribution. Recent efforts have focused

on deriving a wing shape for minimum drag in nonlinear flow

regimes. The ONERA M6 wing, simulated with a coarse Euler

grid, was used as a test case. An Euler analysis of the baseline

configuration at Mach = 0.84 and 3.06 degrees angle of attack is

shown in the figure. Design studies were conducted to minimize

drag with the total lift constrained. Shape functions were applied

alternately to the upper surface and the mean camber plane with

various spanwise distributions. A typical optimization result is

shown in the figure. The L/D ratio, including estimated skin friction drag, increased from 11.24 to 12.35. An optimization for a

200,000 point grid required 1-2 Cray Y-MPhours and 20 megawords of memory.

Significance

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

and numerical optimization techniques will reduce design cycle

time for subsonic and supersonic aircraft.

Future Plans

Inclusion of analytically derived sensitivity derivatives will be

investigated as a means of reducing computational time and

reducing potential noise arising from finite-difference gradient

computation.

Keywords

Aerodynamics, Computational fluid dynamics

Upper SUrldce isobars on an ONERA A46 wing with UD =1 1.24 (left side); lift-constrained numerical optimization

with UD =12.35 (right side) at Mach = 0.84 and 3.06 degrees angle of attack.

Aeronautics/Basic Research
Simulation of Spatially Evolving Jets

Fernando F. Grinstein, Principal Investigator

Naval Research Laboratory

Research Objective

To characterize the underlying fluid-mechanical processes

leading to the deformation and breakdown of vortex rings in rectangular jets, their dependence on aspect ratio, and their role in

entrainment and transition to turbulence.

Approach

Monotonic algorithms on structured grids were used to develop a

three-dimensional numerical simulation of rectangular jets developing in space and time. The model uses fourth-order fluxcorrected transport algorithms and appropriate inflow and outflow boundary conditions for convective transport, finite-rate

chemistry, and temperature- and species-dependent diffusive

transport.

Accomplishment Description

The figures show simulation results of an isolated vortex ring with

an aspect ratio of 4:1 as it is convected away from the nozzle.

The evolution of the ring is visualized in terms of instantaneous

iso-surfaces of the vorticity magnitude. The self-deformation of

the ring first follows a pattern similar to that observed for the

2:1 jet, leading to a rectangular-shaped ring with its axis rotated

90 degrees relative to its initial cross section. As the ring moves

farther downstream and away from the nozzle, the centers of its

longer portions approach each other, vortex reconnection takes

place, and the ring eventually bifurcates into two smaller vortex

rings that are linked by very thin threads of vorticity. Vortex-ring

bifurcation was suggested by laboratory experiments with

4:1 rectangular jets. The figures reveal that the detailed vorticity

dynamics in this vortex "fission" process involve self-deformation

and axis-switching followed by vorticity redistribution leading to

the bridging of antiparallel portions of the vortex ring and the

forming of threads linking the new vortex rings. This simulation

used 23 megawords of memory and about 64 Cray Y-MP hours.

Significance

The simulations provide insight into the physics of large-scale

coherent structures in rectangular jets and the mechanisms affecting entrainment and the transition to turbulence. The effort

advances the state of the art of turbulent jet simulations.

Future Plans

The study of reactive jets will continue and the flame extinction

phenomena in jets will be investigated.

Keywords

Fluid mechanics, Vortex, Turbulence

Iso-surfaces of the vorticity magnitude equal to 40 percent of its

initial peak value. Flow evolves spatially from bottom to top in

each frame. Time increases from left to right in each row, and from

the bottom row to the top row.

Iso-surfaces of the vorticity magnitude from a different perspective

and for 15 percent of the peak vorticity value.

Aeronautics/Basic Research
Aeroelasticity of Wing—Body—Control Configurations

Guru P. Guruswamy, Principal Investigator

Co-investigators: Chansup Byun, EugeneTu, and Shigeru Obayashi

NASA Ames Research Center/MCAT Institute

Research Objective

To compute the aeroelasticity of wing—body—control

configurations using the Navier—Stokes flow equations directly

coupled with finite-element structural equations.

Approach

The finite-difference based three-dimensional Euler/

Navier—Stokes equations of motion coupled with the finiteelement plate/shell structural equations of motion were solved

by using a time-accurate numerical integration scheme with

configuration-adaptive dynamic grids. An accurate and robust

domain decomposition method suitable for aeroelasticity associated with strong fluid/structural interactions was used.

Accomplishment Description

The computational domain was divided into fluid and structural

domains. The information between domains were communicated

using a robust grid-point (fluid) to element (structural) approach.

A typical unsteady computation of a wing—body—control

configuration with 1 million grid points required about

35 Cray C-90 hours and 33 megawords of memory.

Significance

This research has a major impact on the Advanced Subsonic Civil

Transport (ASCT), High Speed Civil Transport (HSCT), and High

Performance Computing and Communications programs. It can

also be used for Department of Defense projects.

Future Plans

This technology will be extended to wing—body—controlempennage configurations for the ASCT and HSCT programs.

Keywords

Aeroelasticity, Controls

Publications

1. Guruswamy, G. P.; and Byun, C.: Fluid-Structural Interactions

Using Navier—Stokes Flow Equations with Shell Finite-Element

Structures. AIAA Paper 93-3087, 24th Fluid Dynamics Conference, Orlando, Fla., July 1993.

2. Obayashi, S.; Chiu, I.; and Guruswamy, G. P.: Navier—Stokes

Computations on Full-Span Wing—Body Configurations with

Oscillating Control Surfaces. AIAA Paper 93-3687, AIAA

Atmospheric Flight Mechanics Conference, Monterey, Calif.,

August 1993.

Flow over an HSCT--type wing—body—control configuration with oscillating control surfaces.

Aeronautics/Basic Research
Wing—Body Aeroelasticity

Guru P. Guruswamy, Principal Investigator

Co-investigator: Chansup Byun

NASA Ames Research Center/MCAT Institute

Research Objective

To perform aeroelastic computations on realistic configurations

using the Intel iPSC/860 parallel computer.

Approach

The finite-difference-based three-dimensional Euler equations of

motion coupled with the finite-element plate/shell structural

equations of motion are solved by using a time-accurate numerical integration scheme with configuration adaptive dynamic

grids. A domain decomposition method suitable for parallel computers is used. In this method, the computational domain is

divided into fluids and structural domains. The information

between domains is communicated using a robust grid-point

(fluid) to element (structural) approach. The fluids domain is

solved in parallel with the structural domain in a separate group

(cube) of processors. Within each cube, computations are performed in parallel.

Accomplishment Description

A new version of the computer code ENSAERO that solves the

Euler equations simultaneously with the finite-element structural

equations of motion has been developed on the Intel iPSC/860

computer. The flow is modeled using H-O type grid topology

using 250,000 grid points. The structures of the wing and body

are modeled using plate and shell elements. This resulted in a

structural matrix equation with 1,641 degrees of freedom. The

fluids part of ENSAERO, including the moving grid, is computed

in a cube of 32 processors and the finite-element structures are

computed in a cube of 8 processors. A typical aeroelastic

response (shown in the figure) required about 20 central processing unit hours. The same computation requires 8 Cray C-90

hours. The Intel iPSC/860 computer performed at an efficiency of

about 10 percent for this practical problem.

Significance

The successful implementation of ENSAERO with finite-element

structures on the Intel iPSC/860 is a major step in the development of general purpose computer codes to solve fluid/structures

interaction problems on parallel computers.

Future Plans

The structures solver will be improved because direct solvers do

not scale well for larger problems. This technology will be

extended to model more complex configurations using zonal

grids for fluids and substructures for structures. Navier—Stokes

capability with Baldwin—Lomax turbulence modeling, moving

control-surface, and tail modeling capabilities will be added to

ENSAERO. Transonic aeroelastic results will be computed for

advanced subsonic civil transport configurations. Detailed finiteelement capabilities including active controls and optimization

disciplines will be implemented.

Keywords

Aeroelasticity, Finite elements, Euler equations

Publication

Byun, C.; and Guruswamy, G. P.: Wing—Body Aeroelasticity

Using Finite- Difference Fluid/Finite - Element Structural Equations

on Parallel Computers. AIAA Paper 94-1487, Apr. 1994.

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Time (sec)

Aeroelastic response of the Boeing 1807 high-speed civil

transport model in pitch-up motion.

Aeronautics/Basic Research
Multipoint Aerodynamic Design

Forrester T. Johnson, Principal Investigator

Co-investigators: Michael B. Bieterman, Craig L. Hilmes, William P. Huffman, Robin G. Melvin, and David P. Young

The Boeing Company

r^

Research Objective

To demonstrate and evaluate a multipoint aerodynamic design

capability for three-dimensional (3-D) aircraft configurations on

Cray supercomputers. Interim objectives supporting these goals

are the application and evaluation of recently developed techniques for improved aerodynamic analysis and single-point

design of complex configurations.

Approach

The TRANAIR code, a full-potential/boundary-layer coupled

code that uses solution adaptive local refinement of Cartesian

grids to resolve flow fields about complex aircraft configurations,

is used in this project. To solve the discrete system of flow equations, a Newton method is used that incorporates incomplete

sparse matrix factorizations. A single-point design optimization

methodology has been added to TRANAIR to enable treatment of

many realistic constraints and objective functions. In this project,

improvements that were made to TRANAIR's design analysis

capabilities are applied, evaluated, and extended to multipoint

design problems. Applications of particular interest are those

involving high-speed civil transport (HSCT) configurations.

Accomplishment Description

One recently developed improvement that was tested on the

Cray C-90 is second-order upwinding, which can result in substantially better accuracy in supersonic flow regions. This method

is especially promising for HSCT problems requiring high resolution of flow features near the aircraft or for sonic boom, wind

tunnels, and other simulations where accurate capture of shocks

in the field away from the aircraft is of interest. A recently developed integral boundary layer coupling capability was evaluated.

This capability allows the modeling of viscous effects and enables

these effects to be incorporated into design objective functions

and constraints. Drag accuracy, the use of drag minimization in

single-point design, and recently implemented treatment of manufacturing, structural, and flow constraints were evaluated. This

work contributed to the development of improved aerodynamic

optimization formulations for HSCT configurations. A multipoint

design optimization methodology was developed and applied to

two-dimensional (2-D) problems in a workstation environment.

Typical computer requirements for 3-D configurations in this

project have been 2-4 Cray C-90 hours and 64-128 megawords

of memory.

Significance

Competitive commercial transport aircraft designs require stateof-the-art aerodynamic performance under nominal cruise conditions and acceptable performance under "off-design" conditions.

The design cycle time of such aircraft may be significantly shortened if computational fluid dynamics design tools that achieve

some degree of optimality for multiple flow conditions are available early in the design cycle. Such cycle time reduction may be

especially valuable for future aircraft such as the HSCT, where

high performance in both supersonic and transonic flight regions

would be desirable.

Future Plans

Evaluation of analysis and single-point design method improvements that are necessary for efficient and accurate multipoint

design will continue. The multipoint design methodology tested

for 2-D problems will be implemented and evaluated for 3-D aircraft configurations in the multiple-CPU Cray C-90 environment.

Keywords

TRANAIR, Adaptive grid, Boundary layer coupling

Mach contours of a TRANAIR solution of a high-speed civil transport problem; yellow = high and dark green = low.

Cross section of the inboard wing and nacelle on the 7.2-millionpoint computational grid.

10 Aeronautics/Basic Research
Vortex-Flow Control

Steven B. Kern, Principal Investigator

Naval Air Warfare Center

r

k^

Research Objective

To investigate unique vortex-flow control concepts on a common

configuration and to apply the most promising of these concepts

to a complete aircraft configuration such as the F/A-18E.

Approach

A combined computational and experimental approach is used

for this research. Applying the OVERFLOW and CFL3D codes,

the three-dimensional Reynolds-averaged Navier—Stokes

equations are used to simulate the vortical flow fields resulting

from three unique vortex-flow control concepts: leading-edge

extension/wing junction fillets, spanwise blowing to control the

secondary vortex, and windward side modifications. These concepts are being investigated on a single delta wing and a cropped

double delta wing. The most promising concepts are tested experimentally and, once validated, can be applied to a realistic fighter

configuration.

Accomplishment Description

To date, a study of the effects of spanwise tangential blowing from

a slot on the upper surface of a delta wing and a study on the

effects of grid refinement and embedded grids on a cropped

double delta wing configuration have been completed. Each

run required an average of 32 megawords of memory and

7 Cray C-90 hours. Spanwise tangential slot blowing at the point

of boundary layer separation was effective in completely eliminating the secondary separation vortex on the delta wing. The

double delta wing with fillets was tested in the NASA Langley

7- by 10-Foot Transonic Wind Tunnel, 8-Foot Transonic Wind

Tunnel, and Basic Aerodynamic Research Tunnel. The figure

shows the computational surface pressure coefficients compared

Computational surface pressure coefficients compared to pressure

sensitive paint and surface pressure taps at (a) 25, (b) 75, and (c) 90

percent chord stations.

to pressure sensitive paint and surface pressure taps at 25, 75,

and 90 percent chord stations.

Significance

The control of vortices has the potential to increase lift and rolling

moments on demand, which will enhance aircraft maneuverability, agility, and controllability.

Future Plans

Vortex-flow control concepts on the complete F/A-18E aircraft at

high angle of attack will be computed.

Keywords

Vortex flows, Fighter configuration, High angle of attack

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Aeronautics/Basic Research 11
Skewed Compressible Mixing Layers

Sanjiva K. Lele, Principal Investigator

Co- investigator: Ganyu Lu

Stanford University

,_ -a

Research Objective

To establish the potential for enhanced mixing because of

free-stream skewing.

Approach

Three-dimensional (3-D) unsteady Navier–Stokes equations are

solved for skewed mixing-layer flows. The linear-stability analysis

of the skewed 3-D flow is used to prescribe inflow disturbances.

The unsteady dynamics of the large-scale structures and the overall statistical evolution of the flow is studied.

Accomplishment Description

Direct numerical simulations were conducted to study a particular subclass of skewed compressible mixing layers—mixing layers

between streams with equal velocity magnitude, but skewed in

the opposite directions. The inflow and initial conditions were

prescribed by the linear stability eigenfunctions for the most

unstable waves. Nonlinear spatial evolution of the vortex

structures were then simulated. For these flows, a stationary

spanwise wave was the most unstable wave for low-convective

Mach numbers and was used as the inflow disturbance in the

direct numerical simulation. The grid size of the simulation was

251 x 189 x 65. The first figure shows the vortex lines and cuts of

a passive scalar field for a skewed mixing layer with a 90-degree

skewing angle. The free-stream Mach number is 0.6, which keeps

the stabilizing compressibility effect at a small level. Near the

inflow, the disturbance rolls up and forms a pattern of streamwise

vortices. Further downstream, the straight concentrated vortex

tubes suddenly spread out. The vortex breakdown can enhance

the mixing between the two streams, which is shown quantitatively by the spreading of momentum thickness for different

skewing angles. A typical simulation requires approximately

80 Cray C-90 hours, 15 megawords of central memory, and

110 megawords of disk scratch space.

Significance

Skewing the free streams offers a simple method for mixing

enhancement. 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.

Keywords

Vortex breakdown, Mixing enhancement

Publications

1 . Lu, G.; and Lele, S. K.: Inviscid Instability of Skewed Compressible Mixing Layers. J. Fluid Mech., vol. 249, Apr. 1993,

pp. 441 -463.

2. Lu, G.; and Lele, S. K.: Spatial Growth of Disturbances in a

Skewed Compressible Mixing Layer. AIAA Paper 93-0214,

1993.

3. Lu, G.; and Lele, S. K.: Vortex Breakdown in Skewed Compressible Mixing Layers. AIAA Paper 94-0821, 1994.

Vortex lines and cuts ut the passive scalar field of the skewed mixing layer with equal, but oppositely, skewed streams (red = lowerstream fluid, blue = upper-stream fluid, green = mixed fluid).

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Downstream distance

Momentum thickness for different skewing angles as a function of

downstream distance.

12 Aeronautics/Basic Research
Comparison of TLNS3D Computations with Test Data

Frank Marconi, Principal Investigator

Co-investigator: M. Siclari

Grumman Corporate Research Center M

Research Objective

To validate five NASA-developed codes for use in aircraft design.

Approach

Five codes were selected for evaluation by NASA and the

Multidisciplinary Analysis and Design Industrial Consortium

(MADIC). This project concentrated on the evaluation of the thinlayer Navier—Stokes TLNS3D code. The code was designed for

transonic external-flow computations. The application of the

code to a transonic transport wing with a simple axisymmetric

fuselage was studied. The configuration was realistic, but

it excluded geometric complexities that cannot currently

be resolved.

Accomplishment Description

TheTLNS3D code was exercised to certify its applicability as a

design tool, thus the ability of the code to predict aerodynamic

forces and moments accurately was a major focus. To assure a

detailed certification of the code, test data were selected that

included significant surface pressure data and force and moment

data. The flight conditions tested were typical of cruise. The configuration matched that tested in the Aircraft Research Association

(UK) 9-ft by 8-ft wind tunnel. Significant surface pressure data

and forces and moments were measured for a Mach test matrix a

containing 14 points. The TLNS3D code was found to be user

friendly and robust. The grid resolution required to capture the

wing surface pressure distribution accurately was surprisingly fine

and resulted in larger than expected running times. Although the

experimental data were intended for computational fluid dynamics validation and were well documented, an unaccounted for

aeroelastic effect caused a 10 percent error in lift. The investigation concentrated on the effects of free-stream conditions (Mach

and angle of attack), numerical dissipation (matrix and scalar),

and turbulence model and grid resolution. This code assessment

was a snapshot in time; the code was exercised essentially in its

current form (no major changes were made in the code during the

evaluation).

Significance

The accuracy of the TLNS3D code was enhanced and the matrix

form of the dissipation was made more robust.

Future Plans

Code improvements will be identified.

Keywords

Aircraft design, External flow

Publication

Marconi, F.; Siclari, M.; Chow, R.; and Carpenter, G.: Comparison

of TLNS3D Computations with Test Data for a Transport Wing/

Simple Body Configuration. AIAA Paper 94-2237, 25th AIAA

Fluid Dynamics Conference, Colorado Springs, Colo., June 1994.

Wing/simple body surface pressure (A/L = 0.8, a = —1.46 degrees).

Wing/simple body hanslucent supersonic zone (iVL = 0.8,

(Y = 2.02 degrees).

Aeronautics/Basic Research 13
Navier—Stokes Computations on Unstructured Grids

Dimitri J. Mavriplis, Principal Investigator

ICASE

r^

7

Research Objective

To develop an accurate and efficient method for computing

steady state turbulent viscous flows about complex threedimensional (3-D) configurations.

Approach

The steady state Reynolds-averaged Navier—Stokes equations are

solved on an unstructured mesh using a Galerkin finite-element

discretization strategy. An edge-based data structure is employed

to minimize memory overheads. An unstructured multigrid

approach is employed to accelerate convergence to steady state.

A combination of microtasking and autotasking has been

employed to parallelize the solver on the shared memory

architecture of the Cray Y-MP.

Accomplishment Description

The low overheads incurred by the present algorithm as well as

the parallel implementation enable the solution of flows on very

fine grids in rapid turnaround time. The turbulent flow over

a high-lift wing geometry has been computed on a grid of

1.8 million points, or 10 million cells, using 320 megawords

of memory in 1.5 hours of wall clock time on the 16-processor

Cray Y-MP.

Significance

The ability to compute highly resolved flows over complex

geometries is of extreme importance to the aircraft and propulsion industries. In particular, 3-D flows over high-lift geometries

have proved particularly difficult to simulate in the past.

Future Plans

Improved multigrid unstructured mesh algorithms are being

investigated. The inclusion of adaptive meshing techniques

and computations over more complicated geometries are

also planned.

Keywords

High lift, Turbulent flows, Complex geometries

Publication

Mavriplis, D. J.: A Three-Dimensional Multigrid Reynolds

Averaged Navier—Stokes Solver for Unstructured Meshes. AIAA

Paper 94-1878, June 1994.

Illustration of fine mesh employed for 3-D high-lift configuration.

Computed Mach contours (symmetry plane) and density contours

(wing surface) for 3-D flow over high-lift wing geometry.

14 Aeronautics/Basic Research
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CFL3D Code Evaluation

Pradeep Raj, Principal Investigator

Co-investigators: Tom Kinard and Richard Semmes

Lockheed Aeronautical Systems Company

Research Objective

To evaluate the capabilities of the CFL3D code for reliably

predicting design and performance data in a timely and easy

manner with established error measures over a clearly defined

envelope of operating conditions. This research is in support of a

joint project between NASA and the Multidisciplinary Analysis

and Design Industrial Consortium on certifying NASA computational fluid dynamics (CFD) codes for use in multidisciplinary

analysis, design, and optimization of aircraft configurations.

Approach

The approach to assess the accuracy, reliability, efficiency,

responsiveness, and ease of use of CFL3D was to (1) apply the

code to a set of test cases, (2) compare the computed results with

measured data and solutions from other codes, and (3) conduct

sensitivity studies for grid size, turbulence models, and boundary

conditions. Data were generated for three configurations: a

Wing C, an ogive-cylinder (axisymmetric) boat-tail, and a twodimensional C-D nozzle.

Accomplishment Description

The Wing C configuration was analyzed for two flow conditions:

Mach number = 0.82 and angle of attack = 5 degrees, and Mach

number = 0.84 and angle of attack = 4.62 degrees. Both flows

were at Reynolds number = 10 million. Grid sensitivity was investigated using two single-block C-O topology grids; one grid had

578,641 nodes and the other had 1,356,225 nodes. Solutions on

both grids agreed well. The pressure distribution on the upper

surface and plane of symmetry are shown in the first figure for the

first Wing C configuration. The second figure shows sensitivity of

surface pressures (at a wing station halfway between the root and

the tip) to four turbulence models. The boat-tail configuration was

analyzed as an axisymmetric body at Mach number = 1.2,

Reynolds number = 0.13 million, and a nozzle pressure ratio

of 12.06. A three-block grid with 17,135 nodes and the

Baldwin—Lomax turbulence model were used.

Significance

This work will accelerate the acceptance and integration of NASA

CFD codes into the industrial design environment by providing

industry an assessment of NASA technology and by providing

NASA with future code development requirements.

Future Plans

Additional sensitivity studies involving grids, turbulence models,

and Reynolds numbers are planned.

Keywords

Computational fluid dynamics, Aircraft configurations

Pressure distribution on the upper surface and plane of symmetry

for a Wing C configuration. Mach number = 0.84, angle of

attack= 4.62, Reynolds number= 10 million.

Wing C sensitivity of surface pressures to four turbulence models.

Aeronautics/Basic Research 15
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Prediction of the Unsteady F-18 Flow Field

Yehia Rizk, Principal Investigator

Co-investigator: Scott Murman

NASA Ames Research Center

F^3

Research Objective

To predict the unsteady flow field around the F-18 aircraft at high

angle of attack using computational fluid dynamics (CFD),

including a study of the effects of simulating more details of the

engine inlet and refining the grid on the vortex burst location. The

effect of grid refinement on the tail buffet and the structural

response of the tail with and without a leading-edge extension

(LEX) fence is examined.

Approach

A Chimera scheme is used to discretize the computational space

around the complete F-18 aircraft including the inlet duct, splitter

plate, and boundary layer vent (first figure). A time-accurate flow

solver (F3D) is used to simulate the unsteady flow field by solving

the Reynolds-averaged Navier–Stokes equations. The structural

response of the flexible vertical tail is computed using modal

analysis. The fine mesh calculations include about 2 million grid

points. A coarse mesh of about 1 million grid points is constructed as a subset of the fine mesh for the grid sensitivity study.

Accomplishment Description

Fine grid calculation with more details of the engine inlet components improves prediction of the vortex burst location (second figure). The results indicate a strong coupling of the external and

inlet flows, especially at the maximum power engine setting. The

time-accurate fine-mesh tail buffet computations show the same

trend observed in the coarse mesh calculations—tail buffet

occurs because the airload dominant frequency coincides with

the tail's first natural frequency of bending. Also, installing the

LEX fence introduces an unsteady vortex interaction that results in

reducing the airload dominant frequency (third figure). Coarse

grid computations typically take about 50 Cray C-90 hours and

require about 12 megawords of central memory for each case.

Fine grid computations typically require twice as much time and

memory as the coarse grid computations.

Significance

This study provides a better understanding of the F-18 flow field

and the phenomenon of the vortex/tail interaction. Understanding

of the buffet phenomenon will lead to more effective means to

control or alleviate buffet and prevent premature tail structural

fatigue for the F-18 and other twin-tail aircraft.

Future Plans

The highly unsteady flow field around the F-18 aircraft at angles

of attack between 45 and 60 degrees will be computed. The

results will be compared with the flight data obtained from

the F-18 High-Angle-of-Attack Research Vehicle with thrust

vectoring.

Keywords

Tail buffet, High angle of attack, Engine inlet

Details of the grid near the engine inlet.

600 • Flight data

• C CFD - coarse grid

• o CFD - fine grid

500 • •

400 •

• • o

x

300

•

•

200 I- o

100'

10 20 30 40 50

a

Vortex burst location comparison.

0 10 20 30 40 50

Nondimensional time

Effect of the LEX fence on vertical tail loading for the fine grid.

16 Aeronautics/Basic Research
Analysis of High-Speed Civil

David M. Schuster, Principal Investigator

Lockheed Engineering and Sciences Company

Transport Transient Response

Research Objective

To analyze the aerodynamic and structural response of future

high-speed civil transport (HSCT) aircraft by coupling computational fluid dynamics (CFD) methods with computational structural mechanics (CSM) algorithms.

Approach

An existing HSCT aircraft concept, the NASA Langley Mach 2.4

HSCT, was analyzed. The Euler/Navier—Stokes three-dimensional

Aeroelastic (ENS3DAE) method was used to perform a quasisteady analysis of the HSCT performing a specified maneuver by

choosing flight conditions at discrete points in the maneuver,

assuming the flow was steady at these points, and performing a

steady, rigid aerodynamic analysis at each condition. Aerodynamic loads were computed at each point in the trajectory, and

these data were transferred to a dynamic, finite-element CSM

method to compute the structural response of the vehicle

throughout the maneuver. Coupling of the CFD and CSM methods involved transfer of loads from the aerodynamic grid to the

finite-element structural model through a conservative loadbeaming technique and the transfer of surface deflections

from the structural model back to the aerodynamic grid

using interpolation.

Accomplishment Description

A loose coupling of the CFD and CSM methods was accomplished and aerodynamic loads on the HSCT concept vehicle

were computed. These loads were successfully transferred to the

corresponding finite-element structural model and structural

deflections were computed. The figure shows contours of the vertical component of the aerodynamic load and the respective

aerodynamic and structural models on which these loads were

applied. The aerodynamic and structural models were independent and a structured single-zone aerodynamic grid was used to

compute the loads, while a grid composed of triangles served as

the structural finite-element model for computing deflections.

Slight differences were seen in the loading pattern because the

values displayed were absolute loads and were dependent on the

local panel areas. Differences in grid refinement and surface

point distribution affect the local load values.

Significance

This research provides a method for performing structural analysis

using detailed aerodynamic surface loads as input. Transient

loads caused by aircraft maneuvers or gusts and computed by the

CFD analysis are used by the CSM method to predict structural

deformations and stresses. These data are important to the design

of efficient, lightweight aircraft structures.

Future Plans

Strong coupling of the CFD and CSM methods to calculate

dynamic aeroelastic loads, structural deformation, and stress will

be performed.

Keywords

Computational fluid dynamics, Computational structural

mechanics, High-speed civil transport

Contours of the computed vertical load component on the aerodynamic (left) and structural (right) models; Mach number = 2.4,

angle of attack = 2.0 degrees. Magenta represents a large load into the page and blue represents a large load out of the page.

Aeronautics/Basic Research 17
Complete F-1 S Hybrid Solution

James L. Thomas, Principal Investigator

Co-investigators: Robert T. Biedron, Christopher L

NASA Langley Research Center

Rumsey, W. Kyle Anderson, and Daryl L. Bonhaus

Research Objective

To perform unsteady three-dimensional (3-D) Reynolds-averaged

Navier-Stokes computations for aerodynamic configurations at

high angles of attack in support of the High Alpha Technology

Program.

Approach

Two algorithms have been merged using a hybrid zonal

technique. The CFL3D code solves the compressible 3-D

Reynolds-averaged Navier-Stokes equations using a structuredgrid finite-volume scheme. The UNS3D code solves the Euler

equations using an implicit unstructured-grid finite-volume

scheme. To solve the linearized system, the GMRES code is used.

Both CFL3D and UNS3D employ flux-difference splitting for

treating the convective terms. The viscous terms in CFL3D are

centrally differenced.

Accomplishment Description

Computations were performed for several high-alpha flows past a

complete F-18 configuration. The solutions were obtained by

using a hybrid zonal technique, and they were a preliminary

study of the tail buffet problem on the F-18 aircraft. For this

hybrid approach, the flow field was divided into two primary

zones: a structured zone over the forebody/leading-edge extension (LEX) where the Navier-Stokes equations were solved in

order to model vortex roll-up, and an unstructured zone over the

aft part of the aircraft where the Euler equations were solved

to model the convection of the forebody/LEX vortices. Earlier

calculations for a simple delta wing indicated little difference

between the vortex path obtained from a similar hybrid zonal

approach as compared to the vortex path obtained using a single

Navier-Stokes zone. CFL3D was used to compute the solution in

the structured zone and UNS3D was used to compute the solution in the unstructured zone; communication between the zones

was established using a modified Chimera grid scheme. The

accompanying figures show the solution obtained for 30 degrees

angle of attack. The vortex bursts between the canopy and the

wing leading-edge/fuselage juncture. This burst position qualitatively agrees with flight-test data. The comparison between the

computed results and flight-test data for the LEX surface pressures

is also shown. Quantitative agreement between the computation

and the flight test was obtained. The computation was performed

using a grid with approximately 325,000 grid points in the structured section and approximately 90,000 nodes in the unstructured section. The computation took 12 Cray Y-MP hours and

required storage of 64 megawords.

Significance

The hybrid structured/unstructured approach relieves many of the

grid-generation difficulties associated with using a structured grid

over the entire configuration, but it retains the ability to track the

forebody/LEX vortices for tail-buffet studies.

Future Plans

The scheme will be modified to perform time-accurate

computations of the F-18 tail-buffet problem.

Keywords

High angle of attack, Finite-volume scheme, Tail buffet

Vortex core streamlines of F-18 hybrid solution at 30 degrees

angle of attack.

-Cp 10 -

0

-1

3 0 0

FS 253

2

-Cp 1

0

3 FS 357

2 0 0

0

0 0

-Cp 1 0 0

COD 

c0

o ° 0

-1 0 .25 .50 .75 1.00

Y/b

Comparison between the computed surface pressures with those

from flight at three stations on the LEX and at 30 degrees angle

of attack.

18 Aeronautics/Basic Research
Flow Simulation about Attack Aircraft

Chung-Jin Wean, Principal Investigator

Co-investigator: John D. Duino

Rockwell International, North American Aircraft Division

A

Research Objective

To demonstrate a computational fluid dynamics (CFD) capability

of computing the complex flows about a highly integrated aerodynamics/propulsion/low-observable vehicle, and to obtain a

better understanding of the physics of this class of flow phenomena. The effects of wind tunnel supporting sting on the vehicle

aerodynamic characteristics will also be investigated.

Approach

A Navier—Stokes solver was used to simulate the flow about a

highly integrated aerodynamics/propulsion/low-observable attack

experimental vehicle and correlate the prediction with the wind

tunnel test data.

Accomplishment Description

A Navier—Stokes solution was obtained for a high-performance/

low-observable inlet/duct with a bell mouth at a static sea level

condition simulating a wind tunnel test condition. The computational grid consists of four blocks of approximately one million

grid points. The solution took 43 Cray C-90 hours and required

28 megawords of memory. Also completed was an Euler flow

simulation for a wind tunnel model of an advanced aircraft configuration with sting at Mach number 0.85 and 3.3 degrees angle

of attack. Four blocked grids were used with 714,580 total grid

points. The computation used 16 Cray-2 hours and 20 megawords of memory for an average Euler run. The predicted force

and drag divergence Mach number compared favorably with the

wind tunnel data.

Significance

The CFD combined with advanced computer graphics can

provide aerodynamic configuration designers with details of flow

around complex geometry and vivid flow visualization for easy,

quick, and detailed analysis. Such detailed flow information and

visualization are generally difficult or too expensive to obtain in

experiments. The calculated results assisted in determining the

instrumentation locations on the tested wind tunnel model.

Future Plans

Effort will be made to obtain a Navier—Stokes solution for the

aircraft model to investigate the effects of the wind tunnel sup

porting sting on the vehicle aerodynamic characteristics.

Keywords

Aerodynamics, Computational fluid dynamics, Propulsion

Pressure contours on a high-performance/low-observable inlet/

duct with a bell mouth (the middle section is not shown); black =

low pressure and white = high pressure.

Mach contours on a wind tunnel model of a fighter with sting;

blue = low pressure and red = high pressure.

Aeronautics/Basic Research 19
Multiple-Body Aerodynamics

David T. Yeh, Principal Investigator

Rockwell International, North American Aircraft Division

r

Research Objective

To assess the feasibility and technical challenges facing the

numerical methods development for simulating the crew escape

process at low-supersonic, low-altitude conditions, and to

develop a better understanding of the aerodynamic and dynamic

aspects of the ejection system.

Approach

The existing overset grid techniques were used and requisite

improvements were made to model the pilot and seat in and out

of an aircraft configuration. The overset grid approach was

chosen because it provides the geometric and computational

freedom to simulate multiple-body separation processes with the

versatility to modify design alternatives with minimal regridding

effort. The OVERFLOW code was used to simulate the flow

phenomena. This code solves the three-dimensional Reynoldsaveraged Navier—Stokes equations through a finite-difference,

central-differencing scheme for structured overlapping grids. The

F-1 6A aircraft was the primary focus of this study because of the

existing wind tunnel results database.

Accomplishment Description

The future trends in designing ejection seats point to high-speed,

low-altitude flight conditions. A test case was conducted to assess

the numerical capability in predicting the flow field around an

ejection seat in the vicinity of the F-1 6A aircraft. Because the flow

condition of interest is in the supersonic regime, only the forebody geometry was included in the calculation. Overlapping volume grids were generated that consisted of the F-1 6A forebody, a

cavity representing the seating compartment, and a generic seat.

The first figure shows the centerline grid arrangement. An additional grid over the cavity was generated to provide the necessary

communication between the cavity and the other grids. A flow

solver intergrid connectivity file was generated to obtain the interface boundary conditions among the overlapping grids. A

Reynolds-averaged Navier—Stokes solution was obtained for a

static case where the seat was fixed at a prescribed location.

The second figure shows the surface grid and associated pressure

contours and some particle traces emanating from the seat.

A typical computation required 16 megawords of memory and

12 Cray Y-MP hours.

Significance

Wind tunnel tests are used extensively to analyze the ejection

devices that can sufficiently reduce wind blast pressure on the

pilot and correct the stabilization problems associated with highspeed ejection. The high cost involved in testing limits the number of concepts and conditions that can be evaluated with limited

knowledge of their effectiveness prior to testing. In spite of significant advances in computational fluid dynamics analysis, continuing effort is required to bring promising procedures and

techniques to maturity and to facilitate applications in design.

Future Plans

Sixth-degree-of-freedom rigid body dynamic analysis and

moving/unsteady grid coupling for multiple-body separation

applications will be investigated.

Keywords

Overset grids, Crew escape

Centerline overlapping grids arrangement.

Flow field about an ejection seat in the vicinity of an F-16A aircraft; Mach number = 1.5 and Reynolds number = 1 million.

20 Aeronautics/Basic Research
:l Cie

II.YI

1IU ikil
Page intentionally left blank 
ing the computational work load. Also, unsteady behavior of a

Y o

hypervelocity base flows can be predicted computationally,

which will help in the design of thermal protection systems.

Future Plans

U

Research will focus on the impact of vehicle geometry and freestream Reynolds numbers on computed unsteadiness. The geom- N _5

etry used in this study was not fully representative of a flight a

vehicle. The free-stream conditions were based on ballistic range

experiments with known levels of unsteady behavior. Simulations

will be performed to determine if the same behavior observed in

a ballistic range is observed in flight.

—10

Time-Accurate Simulation of Hypervelocity Wakes

Andrew Anagnost, Principal Investigator

Stanford University

c

Research Objective

To modify traditional implicit algorithms for solving blunt body

flow fields so that they map efficiently to the message passing

paradigm, and to apply this capability to the time-accurate

simulation of hypervelocity base flows.

Approach

The block implicit Gauss—Seidel line relaxation technique is

employed to solve the Navier—Stokes equations for an unsteady

laminar flow over a blunt body at hypervelocity. The technique

accounts for the physical propagation of information across the

computational domain. If information is only allowed to transverse a block of processors in a single time step, then only the

processors within that block need to communicate in a global,

coupled fashion. By limiting communication, large performance

gains are possible and the stability limitations are only moderately more severe than fully coupled implementation.

Accomplishment Description

Experimentally consistent oscillations are computed for an

axisymmetric base flow at Mach 14, and they are the first timedependent simulations of hypervelocity base flows that demonstrate unsteady behavior. The laminar mechanism of unsteadiness

in the computed solutions agrees well with qualitative and quantitative experimental observations. Gridding and resolution

requirements for capturing the unsteady features of the flow field

are determined and their impact on the character of unsteadiness

is assessed. The line relaxation technique results in simulation

times on 128 nodes of the iPSC/860 that are two times faster than

a traditional Gauss—Seidel solver running on a single pipe of the

Cray C-90.

Significance

Implicit algorithms can be mapped to parallel supercomputers

with significant performance gains without significantly increasKeywords

Hypersonics, Wakes, Parallel algorithms, Time-dependent

simulation

Computed Mach contours for a blunt body base flow at treestream Mach number 14.

10

—15

1.0 1.2 1.4 1.6 1.8

Time (milliseconds)

Computed pressure fluctuations at the wake neck.

Aeronautics/Hypersonics 23
n Experimental, Hanff

Experimental, curve fit

q CFD, NSS

----- (Fn r.urve fit

Effects of Roll Angle on Vortex Breakdown

Neal M. Chaderjian, Principal Investigator

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

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

Research Objective

To determine if the Reynolds-averaged Navier–Stokes equations

(RANS) can simulate an experimentally observed nonlinear variation of rolling moment with static roll angle.

Approach

The time-dependent, three-dimensional RANS equations are

used to compute the flow about a 65 degree swept delta wing at

high incidence. The Navier–Stokes simulation (NSS) code integrates the RANS equations using the implicit Beam–Warming

algorithm. The diagonal algorithm may result in spurious vortex

asymmetries at high angles of attack. Turbulent flow conditions

are treated with the Baldwin–Lomax turbulence model with the

Degani–Schiff modification, which properly accounts for crossflow separation at high incidence.

Accomplishment Description

The NSS code is used to compute vortical flow about a delta wing

with a free-stream Mach number of 0.27 at 30 degrees angle of

attack and a Reynolds number of 3.67 million based on the wing

root chord. Static roll angles up to 42 degrees are investigated.

The first figure shows an instant streamline visualization of the

leeward-side vortices that form when the delta wing is held at

–5 degrees of roll. There is a sudden growth of the vortex-core

diameters, which indicates that vortex breakdown occurs on both

sides of the wing. The helical path of the streamlines indicates a

spiral-type breakdown. The flow downwind of the initial breakdown positions is highly nonsteady. There was good agreement

between the computed and experimental breakdown positions.

The second figure shows that the computed and experimental

time-averaged rolling moment coefficients compare well. Note

the nonlinear variation of the rolling moment coefficient with roll

angle. There are five trim points (zero moment)—three that are

statically stable (negative slope) and two that are statically unstable (positive slope). Some differences between the computation

and experiment are attributed to the need for finer grids.

Significance

This study extends the validity of the present approach to higher

angles of attack and advances the goal of using numerical simulation to study aircraft maneuver aerodynamics at high angles

of attack.

Future Plans

Forced and tree-to-roll motions of a delta wing at high incidence

will be simulated and the effects of grid refinement will be

determined.

Keywords

High angle of attack, Unsteady flow, Vortex breakdown

Planform view of leeward-side vortices visualized with instant

streamlines. Flow conditions: Mach number = 0.27, angle of

attack = 30 degrees, roll angle = –5 degrees, Reynolds number =

3.67 million.

06

a

C:i 04

U

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m 0

E

E -.02

m

-.04

6

It 

–.06 i 1 1 i i 1 i 1 i i i

–50 –40 –30 –20 –10 0 10 20 30 40 50

Roll angle (deg)

Computed and experimental time-averaged rolling moment

coefficients. Flow conditions: Mach number = 0.27, angle of

attack = 30 degrees, Reynolds number = 3.67 million.

24 Aeronautics/Hypersonics
jet Interaction Effects on Aerodynamic Control of Missiles

Robert R. Chamberlain, Principal Investigator

Co-investigators: Anthony Dang and Ken Xiques

Adaptive Research Corporation

lc

Research Objective Keywords

To study the effects of a single supersonic pitch control jet and the Aerodynamic control, Missile aerodynamics

associated jet interaction (JI) phenomena on the aerodynamic

forces and moments experienced by an ogive/cylinder missile

configuration. Investigation of the scaling of cold jet wind tunnel

data to the hot jet flight environment is of particular interest.

Approach

To study the scaling from tunnel to flight, a parametric study

involving the variation of the control jet thrust coefficient was

carried out. Computational fluid dynamics (CFD) predictions for

the forces and moments both with and without a cold jet were

compared to tunnel data to verify the grid (2.7 million nodes) and

the computational procedure. Flight cases at the same free-stream

Mach number (2.2) and angle of attack (-6 degrees), including a

separate species for the hot jet exhaust, were also computed over

a wide altitude range to determine the effects on stability and

control of the full-scale prototype. Individual surface pressure

contributions to the integrated aerodynamic coefficients were isolated for the missile body and the fins to identify the JI effects.

Accomplishment Description

Favorable agreement with experiment was obtained for all tunnel

cases computed. The flight cases indicated that, at the higher altitudes, control of the missile is dominated by the thrust of the jet

alone QI effects are negligible). At low altitudes, however, )I

effects become significant because of the confinement of the jet

shock within the body bow shock and the resulting low-pressure

region immediately behind the jet. The pitching moment is up to

two times over that because of the jet thrust alone. The jet plume/

fin interaction accounts for a significant portion of the moment

amplification and also depends substantially on whether the jet

exhaust is hot (flight) or cold (tunnel). The size, shape, and position of the plume as it passes over the fins dictates how the cross

flow wraps around the body and determines the resulting surface

pressure distribution on the fins. A typical run for this problem

requires 64 megawords of memory and 80 Cray C-90 hours.

Significance

Battlefield engagement simulations, autopilot design, and

guidance and control algorithms all depend upon reliable fullscale flight information about JI control effectiveness. CFD predictions for the increment to the aerodynamic coefficients in

flight caused by JI effects significantly enhance confidence in current scaling procedures for obtaining the required data from cold

gas wind tunnel tests. In some cases, new scaling procedures

were suggested based on the CFD results.

Future Plans

Additional work is planned to examine the influence of Mach

number, angle of attack, and altitude on )I control effectiveness,

particularly when wind tunnel data indicate an instability or some

other unfavorable interaction. The effect of exhaust gas temperature and composition will also be investigated.

Pressure coefficients on the missile with the jet turned on.

Recirculation in thell plume and pressure contours in a cross-flow

plane through the first set of fins. Streamwise velocity goes into the

page.

Aeronautics/Hypersonics 25
.5 1.0 1.5 2.0 2.5 3.0 3.5

X (ft)

2.5

U

N 2.0

N

m 1.5

a^

1.0

0)

c

is .5

m

Transition Model for High-Speed Flow

Hassan A. Hassan, Principal Investigator

Co-investigator: Eric S. Warren

North Carolina State University

Research Objective

To develop a transition model that incorporates information from

linear stability theory and accurately reproduces experimental

data for transitional flows over a wide range of Mach numbers.

Previous work has established a model for including the effects of

first-mode disturbances into a transition model. An objective of

the present work is the extension of this model to account for

Mach number effects. Additionally, the present work includes the

effects of second-mode disturbances that dominate the transition

process at hypersonic speeds.

Approach

The model is based on the fact that the stress in the transitional

region can be expressed as a function of the turbulent and

nonturbulent stresses. The nonturbulent stress is a result of firstand second-mode oscillations. The transition model was incorporated into a Reynolds-averaged Navier—Stokes solver with a

one-equation turbulence model. A variable turbulent Prandtl

number was developed and included in the approach. A finitevolume method was used to evaluate the spatial terms in the

Navier—Stokes equations. An upwind approach based on Roe's

flux-difference splitting was used along with MUSCL differencing.

The solution was stepped in time using a modified Runge—Kutta

method until a steady state was obtained.

Accomplishment Description

Available experimental results for transitional flows are well

predicted by the transition model. Moreover, the results of the

model illustrate the inadequacies of using a constant turbulent

Prandtl number for high-speed flow. Typical runs took an average

of 1 Cray C-90 hour with central memory requirements of about

3-5 megawords.

Significance

The study of high-speed transition is important for the efficient

design of hypersonic vehicles, which encounter flows that may

be transitional over a significant portion of the vehicle. Within

this transitional region, design parameters such as skin friction

and heat transfer are rapidly increasing or are maximum.

Future Plans

Efforts are focused on predicting the transition onset location. A

method based on the reduced Navier—Stokes equations is being

developed with the goal of simulating the transition onset location and transition region. Additional work is beginning on an

improved turbulence model that more adequately accounts for

the presence of pressure gradients.

Keywords

Transitional flow, High-speed flow, Hypersonic

.900

.875 L ^i

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U

w

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.825 — — — — ONERA/CERT

Experiment

.800

O.00E0 5.00E6 1.00E7 1.50E7 2.00E7 2.50E7

Reynolds number

Comparison of model with experiment and linear combination

transition models for Mach 3.5 flow over a 5-degree half-angle

cone.

Comparison of model with experiment for Mach 8 flow over a

7-degree half-angle flared cone.

26 Aeronautics/Hypersonics
Titan Entry Flow Field of the Cassini—Huygens Probe

John Rakiewicz, Principal Investigator

Co-investigators: Scott Ward, Darren Fricker, and David Wyn-Roberts

Jet Propulsion Laboratory/McDonnell Douglas Aerospace/European Space Agency

Research Objective

To use state-of-the-art computational techniques to characterize

the forebody and wake flow-field environment about the

Cassini—Huygens probe as it enters the Titan atmosphere.

Approach

The McDonnell Douglas Navier—Stokes code designated

MDNS3D/ULTRA was used for this study. The code uses a structured zonal grid approach to discretize the full set of Reynoldsaveraged Navier—Stokes equations and includes a generalized

fully coupled finite-rate chemical kinetics model and a fully coupled two-equation turbulence model. The equations are solved in

a time-dependent manner using second-order-accurate central

differencing in space and an explicit fourth-order-accurate time

advancement algorithm. When used with an appropriate totalvariation-diminishing scheme, the applicability of the code

extends into the Mach 30+ range. To model the chemical kinetic

effects of the Titan atmosphere (composed of approximately

86.5 percent nitrogen, 10 percent argon, and 3.5 percent methane), a nine-species, seven-reaction finite-rate chemistry model

was implemented.

Accomplishment Description

Simulations were concentrated on the maximum total heating

trajectory point. Analyses completed at zero angle of attack provide detailed aerothermodynamic and chemical composition

data that address the effects of small perturbations in the freestream chemical composition (because of uncertainties in the

concentration of argon in the Titan atmosphere). A 5 degree

angle-of-attack solution was also completed to assess the effects

of the nonaxisymmetric forebody and wake flow-field structure

on vehicle heating. All the solutions provide connective heating

results and the flow-field temperature and chemical composition

distributions necessary to calculate the radiative component

of the total heat flux to the probe surface. A typical four-zone,

angle-of-attack solution of the forebody and wake flow fields

required approximately 30 megawords of memory and

100 Cray C-90 hours.

Significance

The computational fluid dynamics (CFD) predictions provide

direct assessment of the severity of the environments that will be

encountered during the Cassini—Huygens probe entry into the

Titan atmosphere. The data provided by these simulations cannot

be reproduced using any means other than CFD or measurement

during the actual mission.

Future Plans

This work is complete.

Keywords

Hypersonic, Planetary probe, Aerobrake, Nonequilibrium

chemistry

Pressure distribution around Cassini—Huygens probe enterin„ the

Titan atmosphere.

0.02

0.0

' 0.032

0.0

CN (top) and C (bottom) chemical composition distributions (mole

fractions) around probe.

Aerona uti cs/Hyperson ics 27
Mich number field from interaction of two rarefied free jets

separated by a distance of three exit diameters.

Simulation of Rarefied Jet Interaction

S. H. Konrad Zhu, Principal Investigator

Co-investigator: Leonardo Dagum

Rockwell International, Rocketdyne Division/NASA Ames Research Center

I.V'-1

Research Objective

To characterize the flow field produced by two interacting free

jets expanding into a vacuum and to establish an analytic model

to predict the flow field for the Space Station design.

Approach

The direct simulation Monte Carlo method was employed and

PSiCM, a three-dimensional (3-D) particle simulation code for

rarefied flows, was used to perform a parametric study with a

wide range of the penetration Knudsen number.

Accomplishment Description

PSiCM was validated in a previous work by comparison to

experimental results for the interaction between two rarefied free

jets, which shows excellent agreement. The present parametric

study on the jet plume interactions in the transitional regime has

revealed that a similarity does exist for the secondary jet resulting

from the interaction of two primary jets. The separation distance

between the primary jets is a proper length scale in studying the

plume interaction. The study also showed that predicting multiple

plume flow field as a linear superposition of each single plume is

not accurate, but the inaccuracy decreases as the interaction

approaches the free molecular regime. The study required

500 CM-5 hours and 2 gigabytes of memory.

Significance

This work is a continuation of a task prompted by the Space

Station design and the concern of possible excessive loads on the

solar array from the Orbiter reaction control system (RCS) plume

impingement. An analytical model is needed to determine the

flow field resulting from multiple RCS firings for the design of

the station and the proximity maneuver of the Orbiter. This study

provides a theoretical basis for establishing the model and a

method for quantifying the errors using the linear superposition

prediction.

Future Plans

The simulation will be extended further downstream and the

range of the penetration Knudsen number will be expanded.

Eventually, the analytic model will be established, and the

impingement from the multiple plume interactions will

be investigated.

Keywords

Flow field, Plume interaction

Publications

1. Dagum, L.; and Zhu, S.: Three Dimensional Particle Simulation of High Altitude Rocket Plumes. AIAA Paper 92-2913,

July 1992.

2. Dagum, L.; and Zhu, S.: DSMC Simulation of The Interaction

Between Rarefied Free Jets. AIAA Paper 93-2872, July 1993.

3. Zhu, S.; and Dagum, L.: A Parametric Study of Rarefied Jet

Interaction Using DSMC. AIAA Paper 94-2046, June 1994.

28 Aeronautics/Hypersonics
Subsonic High-Lift Analysis

Kenneth M. Jones, Principal Investigator

Co-investigators: Victor Lessard, Simha Dodbele, Kevin Kjerstad, Steve Yaros, and Mamad Takallu

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

r,r

Research Objective

To provide accurate analyses of complex high-lift systems for

advanced subsonic transport configurations and high-speed

research (HSR) configurations. The long-term goal of the project

is to assess a variety and hierarchy of methods for their usefulness

as analysis and design tools for three-dimensional (3-D) highlift systems.

Approach

Two approaches for analyzing high-lift systems were pursued.

One approach was to use a multiblock Chimera grid scheme with

a 3-D finite-volume Navier–Stokes code (CFL3D). The twodimensional (2-D) option of the code was used with Chimera

grids to analyze a multi-element airfoil. The second approach

was to use a 3-D unstructured-grid Euler solver (USM3D) to analyze a vortex-flow leading-edge-flap configuration that included a

partial span trailing-edge-flap deflection.

Accomplishment Description

The CFL3D code was used in the 2-D mode to analyze a

three-element high-lift airfoil using a Chimera gridding technique.

Five grids were used to model the configuration (first figure).

The analyses were performed over an angle-of-attack range of

4-21 degrees and at Reynolds numbers of 5 and 9 million.

The analyses were performed with the Baldwin–Lomax and

Spalart–Allmaras turbulence models. The surface pressure distributions for both models compare well with experimental data.

The Spalart–Allmaras model produced more accurate results, par

titularly in the leading-edge region of each element. The analyses

required up to 3 Cray-2 hours and 8 megawords of memory. The

USM3D code was used to model leading-edge vortex flaps coupled with deflected trailing-edge flaps on an HSR configuration.

Experimentally obtained force, moment, and flow visualization

results were compared with the computed solution. The analysis

performed with the original unstructured grid generated for this

configuration produced reasonable force and moment comparisons, but did not match the experimentally observed flow-field

pattern—there was no clearly defined reattachment line for the

primary leading-edge vortex. The grid was reclustered in the

vicinity of the core of the vortex, which produced an improved

solution (second figure). The analyses required 122 megawords of

memory and 8 Cray-2 hours.

Significance

Structured Navier–Stokes and unstructured Euler methods are

viable candidates for analyzing complex high-lift systems for

subsonic transports and HSR configurations.

Future Plans

The Chimera approach will be extended to a 3-D high-lift system

using the CFL3D. The unstructured Euler analysis will be

extended to include viscous effects by tying in the Euler solution

with a boundary layer method or by employing an unstructuredgrid Navier–Stokes solver.

Keywords

Navier–Stokes, Chimera, Unstructured grids, Euler

Five-block Chimera grids used to analyze the three-element

airfoil.

Unstructured flow-field grids (original and modified) and surface

pressure distributions with surface streamlines for an HSR configuration. Mach number = 0.22, angle of attack = 16.0.

Aeronautics/Subsonics 29
MFR (10-3)

000

m .025

;_0

F_ .050

m .075

100

- 0.243

= 0.400

= 0.700

Investigation of Forebody Flow Control

Lewis B. Schiff, Principal Investigator

Co-investigators: Ken Gee and Roxana Agosta

NASA Ames Research Center/MCAT Institute

Research Objective

To determine the effectiveness of forebody tangential slot blowing

for flow control on rounded forebodies.

Approach

Time-accurate solutions and a range of free-stream Mach number

and jet exit condition solutions are obtained for an isolated F-1 R

forebody. Solutions are obtained using the generic chined forebody at different angles of attack, slot locations, and jet exit conditions. A multizone, thin-layer Navier-Stokes code is used to

compute the flow fields and the computational results are compared with corresponding wind tunnel experimental data.

Accomplishment Description

At very low values of mass flow ratio (MFR), little or no yawing

moment occurs. At moderate values of MFR, the jet remains

attached to the forebody surface and generates a low-pressure

region. This effect, coupled with the movement of the nose vortices, results in a significant yawing moment. Jet underexpansion

occurs at high MFR resulting in early separation of the jet and

leveling off of the yawing moment with increased MFR. Similar

results are obtained from the generic chined forebody solutions.

At low blowing rates on the chined forebody, moment reversal is

observed. At very high blowing rates, there is a sharp decrease in

the yawing moment. The differences in the effect of blowing are

partly due to the different body geometries. Each solution

required an average of 20 hours Cray C-90 time and

16 megawords of central memory.

Significance

Tangential slot blowing remains an effective method of generating

yawing moment at transonic Mach numbers and over a range of

mass flow ratios.

Future Plans

Analysis is under way to fully understand the flow physics of

tangential slot blowing and force and moment generation.

ti

'1.

(a) (b) (0)

The effect of jet blowing rates on the flow field: (a) low, (b) moderate, and (c) high.

t Slot 1, comp, M_ - 0.2

--A- Slot 2, comp, M_ = 0.2

• 8

-j:1- Slot 1, Experiment, M_ = 0.06

-n- Slot 2, Experiment, M_ = 0.06

C .6

E .4

E 2 a,

.3

> n

-.2 1 1 0 1

MFR (10-3)

Comparison with experimental data shows that the effect of

tangential slot blowing on a generic chined forebody can be accurately predicted using computational methods.

Keywords

High angle of attack, Tangential slot blowing

Tangential slot blowing is effective over a range of free-stream

Mach numbers and blowing rates.

The surface flow pattern shows how blowing changes the flow

field near the surface of the generic chined forebody,

30 Aeronautics/Subsonics
Evaluation of Navier—Stokes Codes for Transport Configuration Analysis

N. J. Yu, Principal Investigator

Co-investigators: H. V. Cao and S. R. Allmaras

Boeing Commercial Airplane Group

Research Objective

To evaluate numerical and turbulence modeling capabilities of

Navier-Stokes codes for transport configuration analysis, and to

validate the direct iterative surface curvature (DISC) inverse

design method for wing and isolated nacelle design.

Approach

The multiblock Navier-Stokes codesTNS3DMB and CFL3D were

used to analyze the 747-200 configuration and other advanced

technology wing/body configurations. Existing wind tunnel data

were used to assess the accuracy of the codes and the applicability of various turbulence models for cruise and off-design analysis. DISC was incorporated in the Navier-Stokes system for wing

and isolated nacelle design. The incompressible Navier-Stokes

code INS2D was evaluated and enhanced for two-dimensional

(2-D) multi-element airfoil analysis.

Accomplishment Description

TheTNS3DMB and CFL3D codes were evaluated for transport

wing/body configurations. Using the same turbulence model, the

codes provide comparable wing pressure results over a wide

range of flow conditions. The two codes also give similar velocity

profiles within turbulent boundary layers. At cruise design

conditions, computed results using the Johnson-King,

Spalart-Allmaras, and k-w models agreed well with test data. For

off-design analysis, both the Spalart-Allmaras and k-w models

provide improved results over the Johnson-King model. The DISC

inverse design method was tested for cruise wing and isolated

nacelle designs. Wings with improved cruise-condition pressure

distributions, and nacelles with nearly shock-free pressure distributions were designed based on this method. An example of a

wing design to improve outboard loading is shown in the figure.

A typical design run required 3 Cray C-90 hours and 60 megawords of memory. The INS2D code was tested on various highlift, multi-element airfoil configurations over a wide range of flow

conditions. It was incorporated in a 2-D Navier-Stokes analysis

system that consisted of an automatic multiblock grid generator

for multi-element airfoils and a method for predicting boundary

layer transition. Numerical studies were conducted to determine

the effects of grid distribution, boundary layer transition,

Reynolds number, turbulence models, and wind tunnel walls.

The results of the numerical studies led to a better interpretation

of the 2-D wind tunnel data and, in many instances, a better

understanding of the physics of high-lift systems.

Significance

Evaluation and improvement of 2-D and 3-D Navier-Stokes

codes are important steps toward transferring the technology from

CFD developers to airplane designers. Designers can use validated codes in the airplane design process. This project works

well for cruise-wing and high-lift wing design.

Future Plans

The feasibility of using Navier-Stokes codes for stability and control problems, such as airplane pitch up, and spoiler aerodynamic

reversals, will be explored. Evaluation of the OVERFLOW and

INS2D codes for complex configuration analysis, such as high-lift

systems, is planned.

Keywords

Transonic, High lift

Wing design using DISC method for input (top), target (middle), and design

(bottom); light yellow = low pressure and green = high pressure.

Aeronautics/Subsonics 31
Supersonic Laminar Flow Control Experiment

Michael C. Fischer, Principal Investigator

Co-investigator: Chandra S. Vemuru

NASA Lewis Research Center/AS&M, Inc.

i_

Research Objective

To develop detailed grids and to analyze F-16XL flight configurations to reduce risk for the F-16XL supersonic laminar flow

control experiment.

Approach

Three-dimensional Euler and Navier—Stokes codes were used

to analyze the F-16XL configurations. The GRIDGEN program

was used to generate the grids for the F16-XL shock-fence

configurations.

Accomplishment Description

Using GRIDGEN, a 27-block grid was generated containing

approximately 2 million grid points. A family of shock fences of

varying height and length were modeled to determine the minimum height required to effectively block the shock originating

from the inlet diverter. Without the fence, the computational fluid

dynamics (CFD) solution revealed a strong shock that traversed

the lower surface and caused undesirable, peaky upper-surface

leading-edge pressures downstream of the intersection with the

leading edge. These peaky pressures prevent the achievement of

laminar flow on the upper surface. With a fence of adequate

height and length at the 65-in. wingspan location, the shock was

blocked and was reflected toward the fuselage. Before, at the end

of the fence, the shock disturbance wrapped around the fence

and traversed outboard, but with the fence it crossed the leading

edge outboard of the laminar flow glove region. The CFD predictions with and without the fence agree well with precursor

F-16XL flight data. Each analysis used about 70-80 megawords of

memory and approximately 4 Cray C-90 hours.

Significance

A baseline fence configuration was selected and is being

evaluated on a 1:15 scale F-16XL model in the Langley Research

Center Supersonic Unitary Plan Wind Tunnel. This fence design

will greatly reduce the F-16XL flight experiment risk.

Future Plans

Refinements to the fence design to improve its effectiveness may

be explored. Other unique F-16XL experiment-related flow problems will be modeled to develop workable solutions that reduce

risk or provide design information.

Keywords

High-Speed Research Program, High-speed civil transport,

Laminar flow

F-16XL lower-surface isobar contours.

F-16XL lower-surface isobar contours with 20-in. shock fence at

65-in. wingspan location.

32 Aeronautics/Supersonics
Unstructured Grid/Flow Solver Calibration

Steve L. Karman, Jr., Principal Investigator

Co-investigators: Tracy J. Welterlen, Elizabeth A. Fuchs, and Doug Howlett

Lockheed Engineering and Sciences Company

Research Objective

To validate a Cartesian unstructured grid Euler/Navier—Stokes

solver for complex three-dimensional (3-D) geometries with

special emphasis on weapons separation.

Approach

An unstructured Cartesian grid computational fluid dynamics

(CFD) code is used to analyze complex 3-D multibody flow

fields. The grids are generated by recursively refining a root

Cartesian cell that encompasses the entire domain. The grid generation is initially based on user-defined boundary spacing and is

later refined based on flow-field gradients. The flow solver uses

upwind differencing for inviscid fluxes and central differencing

for viscous fluxes. A point implicit scheme, which makes use of

Jacobi subiterations for improved stability, is used to advance the

solution to steady state. Prismatic grids can be added for resolving

the boundary layer regions near the bodies in viscous flow analyses. The outer layer of the prismatic grids interfaces with the

Cartesian grid in a conservative manner.

Accomplishment Description

The 3-D version of the code was tested on several simple singleand multiple-body configurations. Calibration of the results for

the ONERA M6 wing showed that the code was in agreement

with the experimental data. The inviscid analysis of an MK-84

bomb separating from a fully loaded F-16 began in late 1993.

Computation of the captive case is nearly completed. The

analysis compares the forces and moments on the MK -84 to

wind tunnel balance results. The current solution is limited to

850,000 cells and is shown in the figure. A shedding vortex

inboard of the MK-84 pylon makes determining forces difficult.

The amount of memory required for this case is 220 megawords

and a run of 100 iterations requires 2.9 Cray C-90 hours.

Significance

The development of this unstructured CFD capability reduces the

analysis time for complex 3-D geometries and flow fields. The

automatic grid generation capability permits rapid problem setup

and configuration changes, enabling the timely evaluation of

parametric studies of complex geometries and analysis of phenomena such as store separation.

Future Plans

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

will continue. The analysis of the MK -84 bomb at several

instances in time will be compared with the wind tunnel captive

trajectory simulation results. A quasi-steady trajectory simulation

will be performed using the integrated CFD forces and moments

and a six-degree-of-freedom package. The computed trajectory

will be compared with the experimental result.

Keywords

Store separation, Adaptive

Mach number contours displayed on solid surfaces and the symmetry plane. A cutting

plane through the MK-84 centerline (inset) shows the grid adaptation results and Mach

number contours.

Aeronautics/Supersonics 33
Laminar Flow Supersonic Wind Tunnel

E. Tu, Principal Investigator

Co-investigator: G. H. Klopfer

NASA Ames Research Center/MCAT Institute

r"3

17

Research Objective

To develop computational tools so that the design of a test model

and its placement within the test section of the Laminar Flow

Supersonic Wind Tunnel (LFSWT) can be verified by numerical

simulation of the Navier—Stokes equations before the model is

constructed. For transition studies in the supersonic Mach regime

it is important to know the extent of clean and undisturbed flow

over the test model.

Approach

Modified versions of the upwind parabolized Navier—Stokes

(UPS) and the compressible Navier—Stokes finite volume

(CNSFV) codes were used to solve the thin-layer Navier—Stokes

equations for laminar flow about the test model inside the LFSWT

The faster UPS code was used for the higher Mach numbers

investigated and the multizonal CNSFV for the lower supersonic

Mach regime.

Accomplishment Description

Computations were performed for a NACA 64A010 wing with a

70-degree leading-edge sweep mounted on the top wall of the

LFSWT at Mach 1.6 for inviscid and viscous flows with the UPS

and CNSFV codes. Various model locations were studied to verify

that the top wall mounted position provided the largest extent of

undisturbed flow on the model. The figure shows the inviscid

shock pattern obtained with the UPS code. Shown are the

impinging shocks on the tunnel walls and the model. The undisturbed region on the model is the triangular region in front of the

reflected shock wave impinging on the model. The flow field

behind the impinging shock wave is no longer undisturbed and

the part of the model behind the shock is useless for any natural

transition study.

Significance

An accurate prediction of the flow field inside the LFSWT with

various test models is important for maximizing the usefulness of

the tunnel, especially when relatively small test sections are considered. Numerical simulation can be used for designing tunnel

modifications and innovative passive and active tunnel devices to

minimize the impact of reflected shock waves on the test model.

This is a cost-effective means of increasing the usable size of

the LFSWT.

Future Plans

The CNSFV code and technique will be extended to unsteady

flows to study test model blockage and tunnel unstart problems.

Keywords

Laminar flow control, Supersonic

Numerical verification of the LFSWT shock pattern of a 70-degree sweep NACA 64A010 wing at Mach 1.6.

34 Aeronautics/Supersonics
Pressure contours on a high-wing subsonic MinspotI at cruise Mach number.

Transonic Flows Around Complex Geometries

Jassim A. AI-Saadi, Principal Investigator

Co-investigators: Peter A. Cavallo and William D. Smith

NASA Langley Research Center/George Washington University

r^s

Research Objective

To accurately and efficiently compute transonic flows around

complex subsonic transport aircraft geometries.

Approach

An unstructured Euler code, USM3D, was used to predict

pressure distributions and flow characteristics over advanced subsonic transport configurations. An interacting boundary layer

(IBL) method was added to USM31D to improve pressure predictions in the transonic range. A high-wing transport with engine

nacelles, pylons, and winglets was analyzed with the Euler code

to assess engine interference effects against previous clean-wing

solutions. The interacting boundary layer code was applied to this

complex configuration and to a low-wing transport to compare its

ability to predict pressure distributions against the Euler

code alone.

Accomplishment Description

The Euler code successfully captured shocks at the wing/pylon

intersections and other flow characteristics observed in wind

tunnel tests of the high-wing transport. The grid for the full highwing configuration contained 1.4 million cells. Both the Euler

and Euler/IBL solver solutions required 4 Cray C-90 hours and

250 megawords of memory. Addition of the IBL method to the

Euler code results in a marked improvement in the high-wing

transport pressure distribution predictions when compared with

experimental data. Accurate prediction of wingledwing interaction required the IBL method. To verify the utility of the Euler/IBL

solver, a low-wing transport solution was obtained with a grid of

286,000 cells. This solution took 1 Cray C-90 hour and required

50 megawords of memory. The IBL method also showed significant improvements in pressure predictions for the low-wing transport configuration.

Significance

Unstructured-grid Euler methods can effectively compute

transonic flows over complex transport aircraft configurations.

The addition of a boundary layer calculation for viscous flow simulation further improves pressure predictions for complex geometries without requiring the grid generation time associated with

structured-grid Navier—Stokes solvers.

Future Plans

Reynolds number effects with either the unstructured Euler/IBL

solver or a structured Navier—Stokes solver will be investigated.

Work is under way to predict propulsion installation effects using

a structured-grid Navier—Stokes solver on a low-wing transport

configuration. The full viscous solutions will be compared against

the unstructured-grid Euler and Euler/IBL solutions and experimental data.

Keywords

Unstructured grids, Propulsion integration, Interacting boundary

layer

Euler (Inviscid)

— — — — Euler and boundary layer

O Experimental data, wing upper surface

q Experimental data, wing lower surface

0 .2 .4 .6 .8 1.0

Effect of interacting boundary layer on pressure predictions for a low-wing transport.

—1.0

.5

C 

0

.5

1.0

X/C

Aeronautics/Transonics 35
Simulation of Coupled Fluid

Christopher A. Atwood, Principal Investigator

Overset Methods, Inc.

and Flightdynamic Systems

s^

Research Objective compared to canard-fixed simulations. However, because of the

To develop and demonstrate a multidiscipline simulation capabil- high store ejection rate, the improvement was modest. The

ity involving the nonlinear coupling of aerodynamics, flight accompanying figure shows the 3-D controlled store during the

dynamics, and controls. separation process. A typical 3-D simulation required 16 megawords of main memory and approximately 70 Cray C-90 hours

Approach on a single processor.

This effort addresses the nonlinear interaction of multidiscipline

problems through numerical simulation. The disciplines of aero- Significance

dynamics (Reynolds-averaged Navier–Stokes equations), body This computation represents one of the first 3-D simulations to

dynamics (Euler rigid body dynamics equations), and controls couple complex viscous aerodynamics and control models

(state-feedback controls model) are included. The simplifications together in the same computation. This capability, when fully

typically used in computing aerodynamic interactions and loads developed, offers a means of reducing aircraft design cycle time

for control law design are not included in this study. In order to and cost while enhancing safety and performance.

assess the accuracy of the initial computations, a problem that

could be compared against analytic and experimental results was Future Plans

chosen—a cavity store separation problem with and without More detailed simulations involving reasonably complete aircraft

active control of a canard control surface was used. and the inclusion of a structural dynamics model to allow for aircraft flexibility will be completed.

Accomplishment Description

Application of the control law to both two- and three- Keywords

dimensional (3-D), three-degree-of-freedom cavity store separa- Unsteady, Controls, Navier–Stokes

tion problems revealed improved trajector y characteristics as

x/L

(b) z/L

Uncontrolled _-_-----

------ Controlled

o n Experiment oQpO°°Ohco

CP

e dA^^

oCra

s__

(c)

a'

(d)

`I ` M ` , ^ '`I'

n

A

_8o

-.1 0 .1 .2

Time (sec)

Three-dimensional controlled store: (a) Mach contours on symmetry plane, (b) center of gravity position, (c)

body attitude (0) and canard deflection (S), (d) force coefficient, and (e) moment coefficient.

of .05

v

v

m

ro

CL

E 0

d

10

a

a^

m 5

v

r

ID 0 H

20

a

d

m 0

E

Ql

H

-20

60

(a)

36 Aeronautics/Transonics
^^ r

r _

1^ -

E'

F/A-18 Aerodynamic Assessment

Raymond R. Cosner, Principal Investigator

Co-investigators: Robert H. Bush, William W. Romer, and Paul G. Willhite

McDonnell Douglas Corporation

s

Research Objective

To predict the model support hardware impact on the measured

performance of the F/A-18.

Approach

Parallel computational fluid dynamic (CFD) and experimental

predictions of the sting and distortion performance increment will

be performed to determine if CFD can adequately predict this

increment. The CFD study precisely paralleled the experimental

study, including the exact wind tunnel model geometry, the wind

tunnel walls, and wing tip mounting system in the calculations.

Non-overlapping grids were used to model the geometry. Zone

breaks in the aft end region isolated the grid changes to areas

where the geometry changed from the real to the distorted aft

end. The modified differential approximation proprietary code

NASTD, using the Baldwin—Barth turbulence model option, was

used to compute the flow field. The Arnold Engineering Development Center wind tunnel wall porosity model was implemented

for this study to ensure consistency between the wind tunnel and

CFD wall effects.

Accomplishment Description

This study demonstrates the ability of CFD to accurately predict

sting and distortion performance increments. Accurate increment

predictions require tight controls on the geometry, grid topology

and resolution, and consistent application of the flow solver to

the incremented cases. The flow solver took approximately

200 Cray C-90 hours for each run and used 50 megawords of

memory. The figure shows predicted surface pressure contours on

a boom-supported, distorted, F/A-18A model at Mach 0.85.

Significance

The ability to accurately predict sting and distortion increments

with CFD provides a cost effective means of correcting wind

tunnel performance data for sting and distortion effects.

Future Plans

The ability of CFD to quantitatively predict the performance of

full configuration aircraft configurations opens an array of potential applications. Future work will focus on further validation of

the accuracy of CFD for complex steady and unsteady flows of

engineering interest.

Keywords

Wind tunnel interference, Sting and distortion increments,

Afterbody drag

A full configuration Navier—Stokes calculation of a distorted aft end flow field provides qua ntita tive performance predictions.

Aeronautics/Transonics 37
Launch Vehicle Flow-Field Simulation

Jerry E. Deese, Principal Investigator

Co-investigator: Ramesh K. Agarwal

McDonnell Douglas Corporation

Research Objective

To develop accurate and efficient methods for the prediction of

flow fields over complete launch vehicle configurations including

the effects of plumes.

Approach

The Reynolds-averaged Navier—Stokes equations are solved

on body-conforming curvilinear grids using a finite-volume

Runge—Kutta time-stepping algorithm. Euler, thin-layer, slenderlayer, and full Navier—Stokes options are available. Complex configurations are modeled with a multiblock zonal implementation.

The code includes algebraic, Johnson—King, and two-equation

turbulence models. Chemistry models for nonequilibrium air,

hydrogen/air, and hydrocarbon/air are available.

Accomplishment Description

This effort was focused on the prediction of multiple body launch

vehicle flow fields. A number of calculations were run for

multiple body configurations that consist of a core vehicle with

solid rocket boosters attached. The surface pressure distribution

on a 120-degree segment of a three-booster configuration is

shown in the figures. Initial calculations modeled the solid rocket

plumes with hot air; a 15-species, 13-reaction solid rocket motor

chemistry model is being added to the code for improved plume

modeling. Typical runs required 100-200 megawords of memory

and 50 Cray C-90 hours.

Significance

The ability to predict complete configuration flow fields allows

designers to take component interference into account and optimize the aerodynamic performance of the complete vehicle. An

accurate prediction method allows designers to build an aerodynamic performance database for new configurations that have not

been tested experimentally.

Future Plans

Flow solver and chemistry model validation with experimental

data will continue. The code will be applied to the analysis of

new launch vehicle configurations. Once the methodology is

established for steady flows, attention will turn to unsteady flows

for application to booster-stage separation.

Keywords

Transonics, Navier—Stokes, Base flows

Aorpr'

Front view (top) and aft view (bottom) of the surface pressure

distribution on a 120-degree segment of a three-booster

configuration.

38 Aeronautics/Tran son ics
PM

Space Shuttle Flow Fields with Plume

Daniel F. Dominik, Principal Investigator

Co-investigators: K. Rajagopal, C. 011ing, K. Mani, S. Vuong, J.

Rockwell International, Space Systems Division

r^

Simulation

Wisneski, G. Hock and J. Sikora

Research Objective

To predict transonic aerodynamic loads for the Space

Shuttle Launch Vehicle (SSLV) during ascent by performing

Navier—Stokes computations with plume simulation at full-scale

flight Reynolds number.

Approach

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

ICEM, a computational fluid dynamics (CFD) computer aided

design grid software. Most of the interzonal boundaries were

aligned. The numerical solutions were obtained using the

Reynolds-averaged Navier—Stokes formulation of the USA-RG3D

code with the algebraic Baldwin—Lomax turbulence model.

Calculations include simulation of the three Space Shuttle main

engine (SSME) plumes and the two solid rocket booster (SRB)

plumes. Results were correlated with available flight-test and

wind tunnel data.

Accomplishment Description

The SSLV model simulated the forward and rear attach hardware

connecting the Orbiter, the external tank, and the two SRBs. The

primary protuberances on the body surfaces, including the liquid

oxygen feed line on the right side of the external tank, were modeled to enhance the accuracy of the multibody flow-field simulation. The complete, asymmetric grid model contained 269 zones

and approximately 6 million points. The USA-RG3D flow solver

was used to simulate the turbulent, flight Reynolds number transonic flow over the geometry. The model was broken into

6 sections so that no section required more than 64 megawords of

main computer memory. Each section had overlapping zones and

the solutions were updated by cycling through the sections to

ensure proper communication. Numerical solutions were computed for the redesigned solid rocket booster at Mach = 1.25 with

—5.1 and —3.3 degrees of angle of attack, and the advanced solid

rocket booster at Mach = 1.25 with —5.1 degrees of angle of

attack. Variable specific heat gases were simulated for the SRB

and SSME plumes; the plumes affect the flow field in the aft half

of the vehicle and increase the wing loads.

Significance

In transonic flow regimes, the influence of plumes can significantly alter the forebody load distributions. These computations

demonstrate the CFD simulation of engine plumes and their effect

on the flow field. However, the pressures were over predicted on

the Orbiter base and on the wing lower surface. Better comparison between CFD results and flight data will require a better turbulence model, plume chemistry model, and grid resolution.

Future Plans

SSLV flow fields will be simulated with a turbulence model better

suited for multiple shear flows.

Keywords

Plume effects, Multibody flow fields

Pressure coefficient at Mach = 1.25, angle of attack =

—5.1 degrees; blue = —0.4 and magenta = 0.55.

Pressure coefficient on Orbiter wing lower surface at Mach —

1.25, angle of attack = —5.1 degrees with plumes inactive (left)

and active (right). CFD in contours and flight data in circles;

blue = —0.4 and magenta = 0.55.

Aeronautics/fransonics 39
Reynolds Number Effects on a Delta Wing

W. Kelly Londenberg, Principal Investigator

ViGYAN, Inc.

Research Objective

To continue the calculations for a grid about a delta wing with

a medium leading-edge radius including the sting at varying

Reynolds numbers and angles of attack. The calculations will be

assessed on how well Reynolds number, angle of attack, and geometrical effects on the flow over a highly swept wing are computationally predicted.

Approach

Calculations were performed using the CFL3D code. This multiblock code solves the thin-layer approximations to the threedimensional, time-dependent, compressible Navier—Stokes equations. Coupled with this code are the Baldwin—Lomax algebraic

and the Spalart—Allmaras one-equation turbulence models.

Accomplishment Description

Numerous solutions were obtained for a 65-degree delta

wing configuration using the CFL3D code. Initial pressure coefficients obtained using the Baldwin—Lomax algebraic turbulence

model compared well with experiment for attached flow conditions at 6 x 106 Reynolds number, but the solution became

unsteady at separated conditions. Solutions obtained using the

Spalart—Allmaras one-equation turbulence model predicted pressure distributions that compared equally well for attached flow

cases, and steady solutions were obtained for separated flow conditions. Using the one-equation turbulence model, solutions were

obtained for Reynolds numbers 6 x 10 6 < 120 x 106 for several

angles of attack. The predicted pressure distributions compared

very well with experiment at all Reynolds numbers for attached

flow conditions. The code, however, predicted separation aft of

the experimental location. This resulted in the predicted vortex

forming later than in the experiment, which resulted in a weaker

vortex being computed. Also, the cross-flow shock apparent in

the experiment was not predicted. On average, the solutions

required 6.33 Cray C-90 hours for convergence and used approximately 60 megawords of memory.

Significance

Even though the vortex location is predicted well, the

Spalart—Allmaras one-equation turbulence model does not

predict the separation location accurately for this geometry.

Future Plans

Methods to improve the comparisons will be investigated. The

use of advanced one- and two-equation turbulence models for

comparison improvement will be studied. The effect of transition

location will also be investigated.

Keywords

Computational fluid dynamics, Turbulence models, Transonic

flow

—2.0

—1.6

—1.2

C p —.8

—.4

4

—2.0

x/cr = 0.40

—1.6

—1.2

x/cr = 0.60

%-

I

I

• I

• I

I P

---4

---p

I I ^

x/cr = 0.80

CFL3D upper surface

--- CFL3D lower surface

• Experiment upper surface

o Experiment lower surface

Cp —' 8 i I ;_• •

• I I le

_. 4 •• • I

I I

I

^j(I

d

- ---^-- n-rob 1

4 0 — .2 .4 .6 .8 1.0 0 .2 .4 .6 .8 1.0 0 .2 .4 .6 .8 1.0

y/(b/2) y/(b/2) y/(b/2)

Predicted pressure distributions show poor comparison with experiment for separated flow case. Large-radius leading

edge, Spalart-Allmaras turbulence model, Mach = 0.85, a = 16.59 degrees, Reynolds number = 120 million.

40 Aeronautics/Tran son ics
Pitching Moment Coefficient

Cn —4

—8

—12

High-Angle-of-Attack Aerodynamics

Christopher L. Reed, Principal Investigator

Lockheed Engineering and Sciences Company

r^3

Research Objective

To develop and validate three-dimensional computational fluid

dynamics (CFD) techniques for the aerodynamic analysis of

bodies at arbitrary attitudes.

Approach

CFD analysis techniques were developed starting with a simple

configuration at relatively benign attitudes and then progressing

to more complex configurations at extreme attitudes. The analyses results were compared with experimentally obtained force

and moment data. These analyses are completed using a central

difference finite-volume Navier—Stokes analysis code. The code

uses a diagonalized lower—upper implicit technique to solve the

governing equations. A two-equation k-ka turbulence model was

employed with a wall layer model along no-slip walls.

Accomplishment Description

An extensive wind tunnel test was completed for a 920,000 point

symmetric grid and a 1.8 million grid point full grid for a

600 gallon F-22 tank plus pylon configuration. This allowed a

significant number of analysis-to-experiment comparisons. The

analyses were run at Mach 0.9 and 1.2, from 0 to —90 degrees

angle of attack, and with sideslip angles from 0 to 40 degrees. A

typical analysis required approximately 5,000 iterations to reduce

the mean flow residual 3 to 4 orders of magnitude. The normal

force, axial force, and pitching moment were also monitored to

determine solution convergence. Each case required approximately 50 megawords of memory and 34 Cray C-90 hours. Comparisons with the test data were very favorable.

Significance

The capability to analyze bodies at arbitrary attitudes is important

(particularly to store separation engineers). Bodies ejected during

flight achieve extreme attitudes compared to flight vehicles.

Methods that can accurately predict store trajectories are desirable and CFD analysis promises to fill that requirement. Methods

exist that use a number of steady-state solutions to aid in the

prediction of store trajectories. Lessons learned from steady state

predictions will aid in the development of fully time-dependent

methodologies. It is believed that significant gaps will exist in the

F-22 store force and moment database because of test limitations.

The ability to accurately provide these data using CFD may significantly impact the F-22 program.

Future Plans

The methodology will be further developed and validated. A

low-speed test will be completed using the F-22 600 gallon tank

plus pylon configuration. The configuration will have a number

of pressure taps and force and moment balances. This will allow

more extensive comparisons. The test program will include runs

at extreme attitudes.

Keywords

Store separation, Computational fluid dynamics

r

a.o ;,o

t

,.o

o.o

_,4o

An F-22 tank with pylon is colored by pressure and shown at

various angles of attack. Mach = 1.2 and Reynolds number =

1.2 million.

Normal Force Coefficient

UIAN

0000

00

Cm —4 00

O

—8 000 000

CbOO 000

—12

0 —30 —60 —90 —120 —150 —180

a

Comparisons with available test data indicate excellent

agreement for normal force coefficient (top) and pitching moment

coefficient (bottom). The triangles represent the CFD analysis and

the circles represent the Arnold Engineering Development Center

test data.

Aeronautics/Transonics 41
C 

0

C 

C 

0

Computational Studies of Advanced Wings at High Incidence

Arvin Shmilovich, Principal Investigator

Co-investigators: K. C. Chang and R. A. Pelkman

McDonnell Douglas Aerospace

r_'%

Research Objective

To calibrate advanced computational fluid dynamics (CFD)

methods with various turbulence models for simulating flow

fields about wing/fuselage configurations at buffet conditions.

Approach

The one-half-equation turbulence model of Johnson—King and the

one-equation models of Baldwin—Barth and Spalart—Ailmaras

were used to evaluate numerical methods for predicting wing

separated flows at transonic buffet. The models were implemented in TLNS3D and CFL3D.

Accomplishment Description

Several wing/body configurations representing contemporary

wing designs and advanced wing concepts were used to evaluate

the turbulence models in TLNS3D and CFL3D. Comparisons of

the numerical predictions with wind tunnel and flight-test data

were made for conditions representing cruise-to-buffet and for a

range of Reynolds numbers. Results indicate that better predictions can be obtained for separation-dominated flows using oneequation turbulence models, and thus the capability of the

numerical schemes to simulate flows beyond buffet onset is

improved. For a grid of about 1 million points, a typical run

required 2 Cray C-90 hours and approximately 60 megawords

of memory.

Significance

In contrast to the success accomplished for attached flow conditions and for flows with mild separation, in general, computational tools are of limited use for predicting flow fields at

conditions that result in sizable separated flow regions because

the predictions gradually depart from the experimental data. The

flow conditions examined in this study at buffet onset and post

buffet indicate improved correlations with the aid of new turbulence models.

Future Plans

The validation study will be extended to include other promising

turbulence models. Possible extensions of the turbulence models

to include more adequate wake modeling will be investigated.

Proper wake modeling is a crucial element for the treatment of

buffet flows because the attainment of acceptable accuracy is

contingent on the ability to simulate the interaction between the

shear layers and the wake during flow reversal conditions.

Keywords

Separation, Turbulence, Buffet

Experiment

TLNS3D, Johnson-King

ULM Johnson-King

CF3D. Spalart-Allmaras

CF31D. Baldwin-Barth

J .2 .4 .6 .8 1.0

X/C

Percent semispan

X/C

CFL3D and TLNS3D advanced wing predictions with various turbulence models at transonic buffet (same incidence

and free-stream velocity).

42 Aeronautics/1'ransonics
=1 I I
Page intentionally left blank 
Density contours of an axial cut through the jet; yellow = low

density and red = high density.

Simulation of Supersonic jet Screech

Alan B. Cain, Principal Investigator

Co-investigators: William W. Bower and Yutaka Ikeda

McDonnell Douglas Corporation

Research Objective

To simulate supersonic jet screech to provide modeling

information on the receptivity and acoustic radiation from the

shock—vortex interaction.

Approach

A conservation, finite-volume formulation of the Navier—Stokes

equations are solved using an upwind biased spatial scheme

together with a Runge—Kutta time integration. A hole cut grid

permits the needed grid smoothness to be achieved with the com

promise of multizone interfaces. Damping is added near the com

putational boundaries to minimize boundary reflections because

of imperfect boundary conditions.

Accomplishment Description

The basic three-dimensional simulation of a self-sustained oscilla

tion in a round jet flow was demonstrated. This is the first of the

necessary steps to establish an accurate simulation of screech.

The geometry of this demonstration case matches that of a labora

tory study at NASA Langley. The early stages of this work were

based on two-dimensional and axisymmetric cases. A single

case requires about 200 Cray Y-MP hours and 64 megawords

of memory.

Significance

Screech acoustic amplitudes can reach 160 dB and are known

to produce damage to the aft end of certain military aircraft. An

accurate and rapid analytical design tool is needed to ensure that

future designs will not produce screech at an amplitude that

results in structural damage. Numerical simulation is the only

method of obtaining detailed information to validate or guide

analytical models of the acoustic radiation produced by the

shock—vortex interaction.

Future Plans

Numerical simulations will be used to guide or verify analytical

modeling of nozzle—lip receptivity and shock—vortex interactions

Keywords

Acoustic resonance, Jet plumes

Publication

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

of the 14th IMAC World Congress, Atlanta, Ga., July 1994.

Pressure contours of a streamwise cut through the jet centerline;

dark blue = low pressure and white-red = high pressure.

Aeronautics/Propulsion 45
NASA rotor 37 near stall showing clearance vortex passage—shock interaction.

Multiblock Solver for Turbomachinery Flows

Rodrick V. Chima, Principal Investigator

NASA Lewis Research Center

Research Objective

To develop and validate a multiblock Navier—Stokes code for

analysis of three-dimensional viscous flows in turbomachinery

Approach

The SWIFT multiblock code solved the thin-layer Navier—Stokes

equations with the Baldwin—Lomax turbulence model using a

finite-difference scheme. An explicit, multistage Runge—Kutta

scheme with implicit residual smoothing was used to solve the

flow equations. The code uses H-type grids for inlet duct regions,

periodic C-type grids around the blade and in the wake, and

0-type grids in the tip clearance region above the blades. Asynchronous input/output is used on the Cray to move the grid and

solution blocks between the solid state storage device and the

core memory used by the solver.

Accomplishment Description

A transonic compressor rotor (NASA rotor 37) was analyzed as

part of a blind test case for turbomachinery codes developed by

the American Society of Mechanical Engineers. Detailed experimental measurements of the rotor flow field were withheld from

the test case participants until the computational results were submitted. The computational grid had just over 1 million grid points

in 3 blocks, with 13 points across the tip clearance gap. Nine

operating points along the 100 percent speed line were computed. Required processing time was about 4 Cray C-90 hours for

each operating point.

Significance

High-resolution solutions of the flow in the tip clearance region

allow detailed study of the losses because of tip leakage and also

allow evaluation of simpler tip models used in other codes.

Future Plans

Detailed comparisons will be made with the experimental data

to verify the code. The SWIFT code will be extended to allow

steady calculations of multiple stages using an averaging

plane approach.

Keywords

Turbomachinery, Compressors, Tip clearance

46 Aeronautics/Propulsion
Airframe/Inlet Aerodynamics

Wei J. Chyu, Principal Investigator

Co-investigator: Tom I-P. Shih

NASA Ames Research Center/Carnegie Mellon University

Research Objective

To develop an analytical capability to predict integrated

performance of forebody/inlet systems for highly maneuverable

aircraft.

Approach

The OVERFLOW Navier—Stokes code is used to study inlet bleed

systems. The PEGSUS code Chimera grid-embedding techniques

are used to integrate grids encompassing bleed holes and freestream and plenum flow fields. A variety of two-equation turbulence models are evaluated to identify the most robust model.

Accomplishment Description

The inlet bleed process under experimental investigation is

numerically studied in a case where the fluid in the boundary

layer was bled through multiple staggered rows of circular holes

into a plenum to control shock-induced flow separation. The

modeling techniques use symmetry boundary conditions to simu

late an infinite row of holes. Turbulence models based on the

standard k—s and k--(o formulations were developed and incorporated into the Navier—Stokes code. The capability of the models

to predict separated flows is being studied.

Significance

This study permits analysis of complex inlets, making possible the

investigation of a broader range of design variables associated

with the integrated forebody and inlet systems and the evaluation

of inlet performance as influenced by the boundary-layer control.

incident shock

- inflow

Turbulence models validated for accurate flow prediction are

essential in the computation of viscous-dominated inlet

flow fields.

Future Plans

The study will be extended to other experimentally investigated

bleed configurations. The effects of critical design parameters on

bleed-system efficiency will be studied. Turbulence models will

be investigated for their ability to predict inlet performance and

associated flow physics.

Keywords

Inlet, Bleed, Turbulence

Publications

1 . Mysko, J. M.; Chyu, W. J.; and Chow, C. Y.: Navier—Stokes

Simulation of External/Internal Transonic Flow on the

Forebody/Inlet of the AV-813 Harrier II. AIAA Paper 93-3057,

July 1993.

2. Chyu, W. J.; Rimlinger, M. J.; and Shih, T. I-P.: Effects of BleedHole Geometry and Plenum Pressure on Three-Dimensional

Shock-Wave/Boundary-Layer/Bleed Interactions. AIAA

Paper 93-3259, July 1993.

3. Rimlinger, M. J.; Shih, T. I-P.; and Chyu, W. J.: ThreeDimensional Shock-Wave/Boundary-Layer Interactions with

Bleed Through Multiple Holes. AIAA Paper 94-0313,

Jan. 1994.

outflow -

do

Boundary layer bleed through multiple staggered rows of circular holes.

Aeronautics/Propulsion 47
Turbine Blade Tip Clearance Flows

Frederik J. de Jong, Principal Investigator

Co-investigator: Tony Chan

Scientific Research Associates, Inc.

Research Objective

To analyze the tip clearance region flow for the gas generator

oxidizer turbine (GGOT) blade, and to study advanced concepts

proposed in an effort to reduce tip clearance losses.

Approach

Turbine blade tip flow fields were computed by solving a

finite difference form of the compressible, three-dimensional

Navier—Stokes equations for a single blade row in a rotating

frame of reference using a linearized block-implicit alternatingdirection implicit procedure. The grids were generated by stacking a series of two-dimensional grids from hub to casing; grid

points were clustered near solid boundaries and in the tip clearance region. At the inflow boundary flow profiles were specified,

while at the outflow boundary the static pressure was prescribed,

Accomplishment Description

Previously, results were obtained for the baseline GGOT blade at

the full-scale Reynolds number on a grid containing approximately 216,000 points. Particle traces in the near-tip region show

vortical flow behavior of the fluid that passes through the clearance region and exits at the downstream edge of the gap. In an

effort to reduce clearance flow losses, the mini-shroud concept

was proposed, which adds a fence to the pressure side of the

blade at the blade tip. However, calculations performed on the

GGOT geometry with the mini-shroud indicate that the minishroud has only a marginal effect on the tip clearance losses at

the design tip clearance. Additional calculations were performed

at two and four times the design tip clearance, both with and

without the mini-shroud, to assess the effect of tip clearance. The

analysis of these calculations is in progress. Each calculation

required approximately 10 Cray Y-MP hours on a single processor

and 11 megawords of memory.

Significance

The ability to accurately predict the tip clearance flow and to

perform parametric studies on changes in the geometry (both

blade shape and tip clearance) is essential to the design of highperformance turbines.

Future Plans

The present code will be applied to the study of aircraft engine

compressor tip treatments.

Keywords

Rocket propulsion, Navier—Stokes analysis, Computational fluid

dynamics

Publication

Chan, Y-T.; and de Jong, F. J.: Navier—Stokes Analysis of an

Oxidizer Turbine Blade with Tip Clearance with and without a

Mini-Shroud. 11 th Workshop for Computational Fluid Dynamics

Applications in Rocket Propulsion, NASA CP-3221, Part 2,

Apr. 1993, pp. 1,397-1,422.

Pressure contours and Particle traces on the GGOT blade with the mini-shroud and at twice the design clearance.

48 Aeronaut ics/PropuIsion
Vortex Induced Mixing Behind Ramp Injectors

W. W. Follett, Principal Investigator

Co-investigators: J. V. Madison and S. Palaniswamy

Rockwell International, Rocketdyne Division/Rockwell International Science Center

s

Research Objective

To investigate the effects of geometry and aspect ratio of

ramp injectors for high-speed vortex mixing through numerical

modeling.

Approach

The unified solution algorithm (USA) was used to compute the

flow fields of several axial ramp injector elements. The USA code

uses the third-order accurate total-variation-diminishing (TVD)

scheme in space, Roe's approximate Reimann solver for convective fluxes, and the scalar approximate factorization scheme for

time advancement. Once the computational fluid dynamics

(CFD) code is validated with the test data, it will be used to design

an optimal ramp injector geometry. The velocities of the injection

jet were matched to the free stream to minimize the effects of

shear mixing and to isolate the vortex structure and its interaction

with the fuel jet.

Accomplishment Description

Four sets of three-dimensional tip-to-tail solutions were obtained

for simulating ramp injectors at an inlet Mach number of 2.07

with an axial injection jet of iodine seeded air at Mach 1.7. The

geometries included a 10-degree swept ramp with 9.5-degree

swept sidewalIs and a base aspect ratio of 1.4, and a 10-degree

straight sidewall ramp with base aspect ratios of 2.0, 1.0,

and 0.71. All ramps had the same length and base height and

were run with the same free-stream and injection conditions.

Each calculation used a 3-zone grid with 820,000 cells

and required approximately 32 megawords of memory and

12 Cray Y-MPhours. Solutions were run two ways—either

laminar or as a Baldwin–Lomax turbulence model to determine

the impact on the vortex development and persistence. All solutions showed reasonable agreement with the experimental data.

The first figure shows the wall static pressure contours and the

vortex/jet particle traces for the swept ramp configuration. The

second figure shows the same quantities for the high-aspect-ratio

straight ramp.

Significance

The solutions provide insight into the structure of complex

vortical flow fields generated by ramp injectors, which will aid in

developing an optimal configuration for high-speed fuel injectors.

Validation of the methodology with experimental data was

accomplished, and it justifies the extension to higher enthalpy

flight conditions. Applying CFD to this class of problems is necessary because of the expense and difficulties associated with fullscale testing for high-Mach-number flows.

Future Plans

The effects of higher Mach number and higher enthalpy flows

will be investigated. A more refined one-equation turbulence

model will be implemented to determine the impact of physical

modeling on vortex strength and persistence. More complex

geometries will be simulated to study interference effects.

Keywords

Mixing, Numerical modeling

Particle traces show the vortex structure and jet interaction for the

10-degree swept ramp case. Wall static pressures vary from 22 psf

(dark blue) to 1,150 psf (white).

Particle traces and wall static pressure contours for the highaspect-ratio straight ramp injector. Wall static pressures vary from

22 psf (dark blue) to 1,150 psf (white).

Aeronautics/Propulsion 49
Unsteady Turbomachinery Computations

Karen L. Gundy-Burlet, Principal Investigator

NASA Ames Research Center

r^

Research Objective

To investigate unsteady rotor/stator interaction in multistage

turbomachines.

Approach

The three-dimensional (3-D) Navier—Stokes codes STAGE-3 and

Rotor-4 incorporate the most modern, high-order upwind-biased

schemes for solving thin-layer Navier—Stokes equations. The

schemes are set in an iterative implicit framework. The codes use

a multizone grid with some zones moving relative to others.

Information is transferred between the various zones using temporally and spatially accurate zonal boundary conditions.

Accomplishment Description

A fine-grid 3-D calculation of a 1.5 stage turbine was completed.

Results from the code compare well with experimental data

for time-averaged surface and cross-flow pressure distributions

and time-averaged surface streamlines. The code requires

93 megawords of main memory for the 2.7 million point grid.

This 1.5 stage turbine calculation required approximately

500 Cray C-90 hours for the time-averaged solution to converge.

The figure shows time-averaged surface pressures overlaid with

surface streamlines within the turbine.

Significance

Unsteady turbomachinery flow fields are extremely complex,

especially in the latter stages of multistage turbomachines. Wake/

wake and wake/airfoil interactions cause complex time-varying

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

A fine-grid, 3-D computation of a 2.5 stage compressor is in

progress. The results of the 2.5 stage compressor computation

will be compared with experimental data from a 2.5 stage compressor. Future computations will include a 4.5 stage compressor

geometry, a multistage/multi-airfoil turbine computation, and an

investigation of the effects of 3-D hot streaks on the flow within

the 1.5 stage turbine.

Keywords

Turbine, Multistage

Publication

Gundy-Burlet, K. L.; Rai, M. M.; and Madavan, N. K.: Unsteady

Three-Dimensional Navier—Stokes Simulations of Multistage

Turbomachinery Flows. AIAA Paper 93-1979, June 1993.

Time-averaged pressures and surface streamlines within a 1.5 stage turbine.

50 Aeronautics/Propulsion
Contours of the time-averaged static pressure inside the tip gap.

Unsteady Three-Dirnensional Flow in a Transonic Compressor

Chunill Hah, Principal Investigator

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

NASA Lewis Research Center/ICOMP/Wright Patterson Air Force Base

Research Objective

To numerically simulate the unsteady flow field inside a

transonic, high-through-flow axial compressor stage, and to

compare the computed results with experimental data.

Approach

The three-dimensional, unsteady Navier—Stokes equations are

solved on structured grids using second-order accurate implicit

time integration, third-order accurate upwind differencing for the

inviscid fluxes, and second-order accurate centered differencing

for the viscous fluxes. A two-equation turbulence model with

a low-Reynolds-number modification is used for turbulence

closure.

Accomplishment Description

The flow field inside a transonic, low-aspect-ratio compressor

rotor was numerically analyzed by solving the unsteady,

Reynolds-averaged Navier—Stokes equations. Previous experimental studies have indicated that some inherent flow unsteadiness exists near the shock and shock-boundary layer interaction

areas even though the compressor rotor was tested as an isolated

blade row. In order to study the origins of the unsteadiness, especially of the shock oscillations, an unsteady calculation of the

flow field as an isolated rotor was performed. Computed distributions of the time-averaged static pressure and the nondimensionalized pressure fluctuation parameter are shown near the shroud

in the figures. The tip-clearance vortex with particle traces is

shown in the first figure. These numerical results agree well with

the unsteady pressure measurements. The unsteady results from

the isolated rotor calculation will be compared with results from

an unsteady rotor/stator stage calculation.

Significance

This work is a direct extension to unsteady flow analysis of an

existing steady Navier—Stokes solver that has been tested and

applied to a wide range of turbomachinery flows. Unsteady analysis is necessary for accurate flow-field prediction for closely coupled blade rows within a stage or between successive stages.

Future Plans

Efforts are under way to analyze the unsteady effects of inlet

distortions and to perform parametric studies for design

optimization.

Keywords

Navier—Stokes, Tip clearance

Contours of the unsteady pressure fluctuation parameter inside

the tip gap.

Aeronautics/Propulsion 51
Transonic Core Compressor Rotor/Stator Interaction

Edward J. Hall, Principal Investigator

Co-investigators: Kurt F. Weber, K. V. Rao, and R. A. Delaney

Allison Engine Company

Research Objective

To assess the effect of rotor/stator interaction on the aerodynamic

performance of aircraft engine axial compressors. Existing compressor data show a significant effect on performance of the

unsteady interaction between rotors and stators in axial compressors. Substantial performance gains can be achieved by optimizing the spacing between adjacent blade rows.

Approach

The analytical tool utilized in this study was the advanced ducted

propfan analysis code (ADPAC). ADPAC was developed for

steady state and time-dependent flow-field simulation of

advanced turbofan engines. The code employs a flexible, multiblocked mesh discretization scheme. The block gridding technique permits coupling of complex multiple-region domains with

common grid interfaces, and permits simultaneous coupling of

solutions for adjacent, relatively rotating blade rows. The solution

procedure employs a finite-volume numerical formulation and a

Runge—Kutta time-marching procedure.

Accomplishment Description

Numerical solutions were obtained for the very high pressure

ratio, two-stage axial compressor design advanced small

turboshaft compressor (ASTC). Numerical solutions were

obtained for individual blade rows operating independently, as

well as time-averaged solutions employing the circumferential

averaging concept. A time-dependent solution was generated for

a realistic representation of the two-stage compressor. The original compressor blade row airfoil counts were slightly modified to

permit a reduced geometric representation of the whole machine

(the unsteady problem can be reduced to one quarter of the entire

wheel with circumferential periodicity). The unsteady solution

requires modeling 4 blade passages of the first rotor, 9 blade passages of the first stator, 8 blade passages of the modified second

rotor, and 15 blade passages of the second stator. The final mesh

contained over 2,000,000 points and required approximately

400 Cray C-90 hours and 125 megawords of central memory to

reach a quasi-time periodic state. An instantaneous snapshot of

the computed velocity field from the time-dependent solution for

the advanced two-stage compressor is shown in the figure. The

interaction between the bow-shock system in the second rotor

and the trailing edge of the first stator is evident, and comparisons

of the time-dependent prediction and the time-averaged predictions for compressor performance show substantial differences.

Significance

Optimization of the blade row spacing in advanced highly loaded

axial compressors is essential for minimizing performance degradation associated with rotor/stator aerodynamic interaction. This

research is relevant to aerodynamic aspects of the Integrated High

Performance Turbine Engine Technology goal of doubling the

thrust-to-weight ratio of fighter engines by the year 2000.

Future Plans

Detailed comparisons of the time-dependent predictions with

measured test data for the ASTC compressor will be performed.

Keywords

Compressor, Interaction

Instantaneous predicted velocity contours at midspan for the ASTC compressor.

52 Aeronautics/Propulsion
(b) {d) (C)

Advanced Ducted Propfan Analysis Code Certification

Edward J. Hall, Principal Investigator

Co-investigator: G. Scott McNulty

Allison Engine Company

Research Objective

To certify a three-dimensional (3-D) Navier—Stokes code for

turbomachinery flow computations intended to form the basis for

an aerothermodynamic/structural optimization program for turbomachinery airfoils planned by the Multidisciplinary Analysis and

Design Industrial Consortium. Certification is required to establish an accurate, robust, user-friendly solver that can be used in

an optimization environment. Particular emphasis was placed on

accurate prediction of airfoil surface aerothermodynamics and

overall aerodynamic performance.

Approach

The 3-D Navier—Stokes advanced ducted propfan analysis code

(ADPAC) was selected for the investigation because it is applicable to a wide variety of turbomachinery flow/heat transfer problems. Certification of the code was performed by obtaining

detailed comparisons of numerical predictions with experimental

data for two test cases. A transonic fan rotor, NASA Rotor 67, was

tested to provide data for code validation. The rotor has 22 lowaspect-ratio blades, a design relative tip Mach number of 1.38,

and a design mass flow rate of 73.3 pounds per second. Laser

anemometer measurements taken within the rotor provide excellent definition of the passage shock structure at the peak efficiency and near-stall operating points. To assess the ADPAC's

ability to predict turbine airfoil heat transfer, the C3X turbine cascade was used. The C3X airfoil, representative of a vane section

from a transonic turbine vane row, was tested in a rectilinear cascade. Both aerodynamic and heat transfer data were acquired for

a range of expansion ratios, Reynolds numbers, and wall-to-gas

temperature ratios. The computations were performed on grids

that were fine enough to yield grid independent heat transfer

solutions.

Accomplishment Description

An array of numerical predictions was obtained for the NASA

Rotor 67 and C3X turbine vane cascade. The rotor calculations

were performed on H-type and body-centered 0-type mesh systems. Details of the tip clearance flow were computed using a

degenerate O-type mesh that discretized the clearance flow

region. Mesh-independent heat transfer predictions were defined

for the C3X turbine cascade by comparing solutions on successively finer meshes. Essentially mesh independent heat transfer

predictions could be achieved when the near-surface mesh spacing resulted in a turbulent velocity profile y+ value of 3 or less.

Typical calculations required 1 Cray C-90 hour and 16 megawords of central memory.

Significance

The use of modern aerothermodynamic analytical techniques

requires extensive validation to achieve the levels of confidence

necessary to base decisions related to design trade-off studies on

numerical simulations. This program is the first step to incorporate an advanced aerothermodynamic analytical tool into a multidisciplinary optimization design program for turbomachinery

airfoils. The program is relevant to the Integrated High Performance Turbine Engine Technology goal of doubling the thrust-toweight ratio of fighter engines by the year 2000, and is applicable

to the High Performance Computing and Communication

Program initiative.

Future Plans

Additional compressor and turbine validation tests will be

performed.

Keywords

Advanced ducted propfan analysis code, Optimization,

Fan, Turbine

Predicted results for NASA Rotor 67; (a) 3-D surface pressure contours at near stall, (b) blade-to-blade Mach number contours for choke,

(c) peak efficiency, and (d) near stall at 90 percent span.

Aeronautics/Propulsion 53
Fan Rotor/Endwall Treatment Interaction

Edward J. Hall, Principal Investigator

Co-investigators: Andrew). Crook and Kathy P. Nardini

Allison Engine Company

Research Objective

To determine the effects of endwall treatments (slots, grooves,

embedded vanes) on the aerodynamic performance of aircraft

turbofan engine fan rotors. Existing fan rotor performance data

indicate a significant improvement in fan stall margin with endwall treatment compared to a similar untreated geometry.

Advanced Navier—Stokes analysis techniques can aid in the

understanding of the complex time-dependent rotor/endwall

treatment aerodynamic interactions that determine the overall

performance gains associated with endwall treatments.

Approach

The analytical tool utilized during this study was the advanced

ducted propfan analysis code (ADPAC). ADPAC was developed

for steady state and time-dependent flow-field simulation of

advanced high-bypass turbofan engines. The code employs a

flexible, multiblocked mesh discretization scheme. The block

gridding technique permits coupling of complex, multiple-region,

relatively rotating domains with common grid interfaces. The

solution procedure employs a finite-volume numerical formulation and a Runge—Kutta time-marching procedure.

Accomplishment Description

Numerical solutions were obtained for a modern turbofan engine

fan rotor design with and without endwall treatments. Fan

pressure ratio and efficiency versus mass flow operating characteristics were predicted at the design speed to determine the baseline rotor stall characteristics. Similar results were generated for

various endwall treatments to determine the effectiveness of the

treatments on enhancing stall margin. Several solutions were

obtained using a time-averaged fan rotor/endwall treatment coupling technique as well as a complete time-dependent solution of

the rotor/treatment aerodynamic interaction problem. Timedependent solutions for the rotor/treatment interaction analysis

typically required 100 Cray C-90 hours and 64 megawords of

central memory.

Significance

Effective application of endwall treatments permits high levels of

blade aerodynamic loading while maintaining adequate stall

margin safety. This research is relevant to aerodynamic aspects of

the Integrated High Performance Turbine Engine Technology goal

of doubling the thrust-to-weight ratio of fighter engines by the

year 2000, and has significant applications for reducing acoustic

signatures of advanced subsonic transport engines.

Future Plans

Additional calculations are under way to examine the effectiveness of endwall treatments for fan rotors operating in the presence

of inlet aerodynamic distortion. Inlet distortion can drastically

reduce engine stall margin. Endwall treatments may ultimately

be designed to accommodate off-design rather than designflow conditions.

Keywords

Fan, Casing treatment, Stall

Publication

Hall, E.; Crook, A.; and Delaney, R.: Aerodynamic Analysis of

Compressor Casing Treatment with a 3-D Navier—Stokes Solver.

AIAA Paper 94-2796, June 1994.

Instantaneous surface static pressure contours for rotodendwall treatment aerodynamic interaction analysis.

54 Aeronautics/Propulsion
Three-Dimensional Steady and Unsteady Turbulent Flows in Turbomachines

Budugur Lakshminarayana, Principal Investigator

Co-investigators: S. Fan, J-F. Gallardo, Y-H. Ho, J. Luo, and W. S. Yu

Pennsylvania State University

r^

Research Objective

To develop advanced turbulence models and efficient numerical

techniques for the prediction of complex steady and unsteady turbulent and thermal fields in turbomachinery applications.

Approach

Three-dimensional (3-D) time-marching and pressure-based

Navier—Stokes codes are used to predict steady and unsteady flow

and heat transfer fields in turbomachines. In the pressure-based

technique, a predictor-corrector type algorithm is applied to the

governing equations to ensure the time accuracy. The timemarching technique incorporates 3-D steady and unsteady nonreflecting boundary conditions based on characteristics theory. A

variety of turbulence models have been incorporated in these

codes to capture complex turbulent flow features such as the

effect of rotation, curvature and anisotropy, and unsteady effects.

Accomplishment Description

The codes were used to predict a variety of turbomachinery flow

fields, providing validation of the turbulence models and numerical techniques. A subsonic cascade flow subject to vortical

incoming waves was computed to demonstrate the effectiveness

of the unsteady nonreflective boundary conditions at the inlet and

the outlet of the flow field. A typical unsteady Euler solution

requires about 1 Cray Y-MP hour and nearly 3 megawords of

memory. The pressure-based method was used to investigate the

structure of secondary flows in a turbine rotor cascade. The

development of secondary flow is captured accurately, including

blade-to-blade variation of all three components, pitch and yaw

angles, and losses and blade static pressures. Each flow computation required approximately 8 Cray Y-MP hours and 28 megawords of memory. The turbulent flow in a strongly curved turnaround duct has been computed using various turbulence models

and the results were compared with experimental data. The

Reynolds stress model (RSM) and algebraic Reynolds stress model

(ARSM) successfully captured the convex curvature effects and

the large anisotropy of the turbulence through the bend. Significant improvement was obtained with the k--F-, ARSM, and RSM

models with a modified dissipation equation. The computation

required about 2 Cray Y-MP hours and 4 megawords of memory.

Significance

The ability to simulate the unsteady viscous flow inside the blade

passages is important in understanding the rotor/stator interaction

effects in turbomachinery. Improved understanding of the complex flow field inside the turbomachinery leads to more efficient

turbomachinery design.

Future Plans

Three-dimensional time-accurate predictions of unsteady flow

caused by rotor/stator interaction are being investigated. The

emphasis is to understand the effects of upstream rotor wakes,

secondary flow, and tip leakage flow on the blade row.

Keywords

Turbomachines, Turbulent flow, Thermal fields

Unsteady pressure and velocity fields caused by wake/blade

interaction.

Turbulent intensity distribution in a 180-degree duct.

Aeronautics/Propulsion 55
Computational Propulsion Applicatio

C. L. Merkle, Principal Investigator

Co-investigators: S. Venkateswaran, J-Z. Feng, H-H. Tsuei, and J

Pennsylvania State University

'I

ns

M. Grenda

Research Objective

To enhance physical understanding in three areas of propulsion

application—film-cooled chemical thrusters for auxiliary propulsion, super-critical coolant channel flow fields, and combustion

instability in liquid rocket engines.

Approach

All three problems involve solutions of the Navier–Stokes equations coupled to additional physics in two and three dimensions.

The chemical thruster analysis involves detailed hydrogen/

oxygen chemistry along with three-dimensional ( 3 -D) and

unsteady effects in the reacting shear layer adjacent to the

thruster walls. The combustion instability problem is 3-D,

time-dependent, and involves two-phase effects. A Lagrangian

droplet-tracking model is coupled to the Eulerian gas phase. The

coolant channel analysis is also 3-D and involves high-speed

flow of super-critical hydrogen.

Accomplishment Description

Several representative calculations for the three problems were

performed. For the chemical thruster problem, a series of

unsteady reacting computations were completed for the wallbounded shear layers and full-engine geometry. Parametric trends

of the effects of molecular weight and wall proximity have been

established (see figures). Typical memory requirements for the

calculations (for 15 periods of oscillation) are 12 megawords and

about 40 Cray-2 hours. For the combustion instability problem,

3 -D response to incipient instability and bombs were simulated

including the effects of droplet atomization and vaporization. For

the coolant channel analysis, heat transfer and pressure drop

have been characterized at several representative Reynolds

numbers. Calculations typically required about 50 Cray-2 hours

and 64 megawords of memory.

Significance

In the auxiliary chemical thruster engines, more than 50 percent

of the fuel is used for wall-film cooling. Computations provide

insight into fuel-film integrity and combustion performance of the

engines. The combustion instability results demonstrated that

droplet vaporization is a likely mechanism for triggering unstable

pressure response in liquid rocket engines.

Future Plans

Computations of 3-D chemical-thruster geometry including

turbulence and finite rate chemistry are being performed.

Unsteady reacting computations with large-eddy simulation

models for subgrid turbulence will also be performed. For the

coolant channel analysis, bifurcated channels and coupled heat

transfer in solid walls and fluid dynamics are being studied.

Keywords

Combustion, Instability

Publications

1. Grenda, J. M.; Venkateswaran, S.; and Merkle, C. L.: Analysis

of Liquid Rocket Instabilities Using a Computational Testbed.

Presented at the International Symposium on Combustion,

Irvine, Calif., July 25-31, 1994.

2. Tsuei, H-H.; and Merkle, C. L.: CFD Analyses of Combustor,

Nozzle and Plume Flowfields. AIAA Paper 94-0553,

Jan. 1994.

3. Yagley, J. A.; Feng, J.; and Merkle, C. L.: CFD Analyses of

Coolant Channel Flowfields. AIAA Paper 93-1830, June 1993.

(a)

(b)

(c)

Instantaneous radical contours in unsteady reacting (hydrogen/

oxygen) shear layer with effects of wall proximity. Distance of wall

from shear layer: (a) 2 cm, (b) 1 cm, and (c) 1 mm.

56 Aeronautics/Propulsion
Complex Three-Dimensional Flows in the Advanced Solid Rocket Motor

Edward J. Reske, Principal Investigator

Co-investigators: Lee A. Kania and James W. Zuercher

NASA Marshall Space Flight Center

r

Research Objective

To characterize the internal flow environment and internal nozzle

aerodynamics for the advanced solid rocket motor (ASRM).

Three-dimensional (3-D) computational fluid dynamic (CFD)

models are tools to address key design issues.

Approach

A 3-D Navier—Stokes flow solver was used to analyze complex

flows in the aft segment and nozzle of the ASRM. Input on physical reference quantities and boundary conditions was obtained

from the Joint Army-Navy-NASA-Air Force Solid Performance

Prediction Ballistics program.

Accomplishment Description

Gimballing a submerged nozzle in a solid rocket motor produces

a pressure differential and subsequent circumferential flow in the

aft-dome region. The hot gases entering, swirling in, and then

exiting the aft-dome region can produce severe environments for

motor components. A fine-grid CFD analysis has been conducted

for the ASRM aft segment with an 8-degree gimballed nozzle to

assess the thermal environment and heat transfer to components

in the ASRM, such as the flex seal, case-to-nozzle joint, and casing insulation. The CELMINT Navier—Stokes flow solving code

has been used to model the flow and thermal environment on a

numerical grid of 311 axial points x 127 radial points x 37 circumferential planes (1.46 meter grid points). Such a large grid

was needed because the computational domain has a length

scale of 600 inches and diameter scales of 54.5-130.0 inches,

whereas the smallest grid spacing was required to be a thousandth of an inch to enable sufficient boundary layer resolution

to facilitate the calculation of heat transfer (despite the fact that

wall functions were being used). This calculation required

125 Cray C-90 hours and 64 megawords of memory. A code validation effort has been performed to benchmark the CELMINT

code for use in modeling the internal aerodynamics of submerged

nozzles in solid rocket motors by comparing it with the 8 percent

ASRM cold flow model. Flow in the cold flow model was simulated by running CELMINT on a numerical grid consisting of

160 axial points x 85 radial points x 37 circumferential planes.

Good agreement was obtained for the static wall pressures predicted by CELMINT and the experimental values. The model

required 25 Cray C-90 hours and 24 megawords of memory.

Significance

The results from these models have provided useful tools and

results needed for designing the ASRM. In particular, these models allow designers to address key issues for the ASRM, such as

the internal environment and nozzle aerodynamics.

Future Plans

The results of the CFD analyses performed over the past three

years will be documented and submitted for publication in professional journals. The results from this work will provide a basis

for the solid rocket motor community to address key design issues

for future solid rocket motor projects.

Keywords

Computational fluid dynamics, Nozzle aerodynamics

Flow visualization for the ASRM fine-grid model showing circumferential flow in the aft-dome region when the nozzle is gimballed

8 degrees. Mach number = 0.224, a = 0.0, Reynolds number = 4.42 x 106.

Aeronautics/Propulsion 57
High-Speed Inlet Design and Analysis

William C. Rose, Principal Investigator

Rose Engineering and Research, Inc.

Research Objective

To examine the flow in a subsonic diffuser located downstream of

a high-speed inlet.

Approach

The three-dimensional (3-D) flow code OVERFLOW is used to

obtain the solution through the supersonic inlet and then determine the flow characteristics (with the attendant thick boundary

layers) within an S-duct diffuser as the flow transitions from a

rectangular to a circular shape.

Accomplishment Description

The OVERFLOW code was used to solve the flow within several

prospective diffusers. The general shape of a diffuser is shown in

the first figure. The diffuser area ratio and the amount of S-duct

offset were varied to achieve a reasonable diffusion rate and flow

quality measurement at the outlet of the subsonic diffuser. The

second figure shows the Mach number contours from a 3-D solution for half of the symmetrical inlet.

Significance

This study shows the technical feasibility of using the

OVERFLOW code to calculate subsonic diffuser flows very

rapidly. Each solution takes approximately 15 minutes. The rapid

solution turnaround time allows the code to be used in an iterative, parametric manner to arrive at a manually optimized design

in a very short period of time.

Future Plans

The study is temporarily suspended pending funding for fiscal

year 1995.

Keyword

Subsonic diffuser flow

View of the surface geometry for the transitioning, S-duct subsonic diffuser.

Mach number contours obtained with the OVERFLOW code in a

representative diffuser. Mach number = 5.020, a = 0.0, Reynolds

number= 1.85 x 106.

58 Aeronautics/Propulsion
Computed and measured pressure contours in the

symmetry plane.

Computed and measured injectant mass fraction in the

symmetry plane.

0 10 20 30 40 50 60 70

Dimensions (mm)

1.0

.8

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.2

Three-Dimensional Mixing Flows in High-Speed Combustors

Balu Sekar, Principal Investigator

Wright Patterson Air Force Base

®^

Research Objective

To apply the state-of-the-art three-dimensional (3-D) computer

codes to predict the flow development and performance of the

high-speed propulsion components, such as supersonic combustors and nozzles.

Approach

The 3-D, time-dependent, compressible Navier—Stokes equations

were solved with algebraic turbulence models and with fully

coupled chemistry for various reaction mechanisms to compute

the mixing and diffusion phenomena that occur in a typical

scramjet combustor.

Accomplishment Description

The GASP code was used to simulate the parallel injection of air

from a swept-ramp base into a Mach 2 free stream of air. GASP

has the capability of computing complicated flows by a general

multiblock (zonal) feature that allows piecing the individual

blocks of grids so that they align. The ease of applying various

boundary conditions to complex 3-D arbitrary internal/external

configurations and the implementation of efficient numerical

algorithms make this code particularly attractive and suitable for

high-speed combustor modeling. Numerical studies carried out

for the swept-ramp combustor configuration were used to determine the key performance parameters, such as penetration and

mixing. The figures show the computed pressure, injected air

mass fraction distribution inside the combustor, and the axial

decay distribution of the injectant at the symmetric plane. Flow

convergence for this configuration of 6 zones with 656,850 grid

points was obtained with 16 Cray C-90 hours on a single processor and 6 megawords of memory. The computed flow compares

moderately well with the experiments.

Significance

Numerical modeling of the complex 3-D flows is necessary to

understand the critical issues associated with fuel mixing. The

validated numerical model aids in pre- and posttest analysis and

gives an in-depth analysis of the complexities of these flows.

These studies are essential tools for the improved design and

development of supersonic combustors with increased mixing.

Future Plans

Research is under way to determine the relative performance,

improvements of different combustors, and fuel injector conceptual designs applicable to scramjet propulsion. This will enhance

the operating characteristics of the propulsion systems because of

increased mixing and combustion and the reduction in the net

flow loss.

Keywords

High-speed propulsion, Supersonic combustion, Mixing,

Diffusion

Computed injectant decay in the symmetry plane.

Aeronautics/Propulsion 59
One view of the 3-D Chimera grid.

Heat transfer coefficient on a leading surface. Rate of

rotation = 22,000.

Flow in Turbine-Blade Cooling Passages

Tom I-P. Shih, Principal Investigator

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

Carnegie Mellon University/NASA Lewis Research Center

r_

Research Objective

To develop, evaluate, and apply advanced computational tools to

study three-dimensional (3-D) flow and heat transfer inside cooling passages of turbine blades.

Approach

When 3-D flow fields inside geometrically complex turbine-blade

cooling passages are computed, the hours involved in the grid

generation process account for most of the time required to

obtain a solution. To reduce grid generation time, overlapping

structured (Chimera) grids were employed. Solutions to the compressible Navier—Stokes equations on the Chimera grids were

obtained using a modified version of the OVERFLOW code.

Accomplishment Description

The OVERFLOW code was adapted to compute flow fields in

cooling passages of turbine blades that have steady state solutions

with respect to a rotating frame of reference. Solutions for flow

and heat transfer were obtained for several cooling passages

including a branched duct where experimental data are available

for code validation and a cooling passage inside a rotating radial

turbine blade. These computations elucidate physics as a function of design and operating parameters. The 3-D solution for the

branched duct with 1.6 million grid points required 28 megawords of memory and 10 Cray C-90 hours. The 3-D solution for

the cooling passage inside the rotating radial turbine blade with

0.7 million grid points required 16 megawords of memory and

5 Cray C-90 hours.

Significance

Currently, cooling passages are designed by using either quasione-dimensional analyses or multidimensional analyses with sim

plified geometries. This study produced a code that allows the

correct geometry to be analyzed in a timely and cost-effective

manner. Also, application of the code produced better understanding of flow physics.

Future Plans

The accuracy of predictions will be improved by using turbulence

models that account for rotation and streamline curvature. Physics of flow and heat transfer inside coolant passages as a function

of design and operating parameters will be studied.

Keywords

Internal blade cooling, Heat transfer

Publication

Stephens, M. A.; Rimlinger, M. J.; Shih, T. I-P.; and Civinskas,

K. C.: Chimera Grids in the Simulation of Three-Dimensional

Flowfields in Turbine-Blade-Coolant Passages. AIAA

Paper 93-2559, June 1993.

Heat transfer coefficient on a trailing surface. Rate of

rotation = 22,000.

60 Aeronautics/Propulsion
Analysis of Nacelle/Pylon/Wing Installations

Arvin Shmilovich, Principal Investigator

Co-investigator: Lie-Mine Gea

McDonnell Douglas Aerospace

wrC4

Research Objective

To validate the OVERFLOW/Chimera code for the transonic

analysis of multicomponent wing/fuselage/nacelle/pylon/winglet

configurations.

Approach

A major difficulty in using computational methods for the

integration of new engine technology on advanced wings is solving the flow problem in regions of complex geometrical shapes.

Combined with the requirement to achieve proper scales of discretization, it is desirable to generate a set of structured grids with

each grid adequate for the flow discretization about an individual

component. In the Chimera approach the analysis is performed

independently on each component grid and it allows for proper

interaction within the regions of overlap. OVERFLOW solves the

time-averaged Navier—Stokes equations using an implicit approximate factorization scheme.

Accomplishment Description

Several contemporary and advanced wing/fuselage/nacelle/pylon

configurations were used to calibrate OVERFLOW and the

Chimera approach. Strategies for generating the individual component grids and prescribing the overset regions were established

to ensure proper communication among the individual subgrids.

For all configurations analyzed, the OVERFLOW/Chimera results

were in excellent agreement with test data.

Significance

Considerations of both noise and propulsive efficiency have

enhanced the attractiveness of high-bypass-ratio engines. This

trend, coupled with the accompanying requirements toward

higher cruise speeds, is making the task of efficiently integrating

the advanced engines with modern wings very complex.

OVERFLOW can analyze these flow situations and its Chimera

implementation offers great geometric flexibility. This study demonstrates the ability of the code to predict the impact of nacelle/

pylon installation on the wing flow development. OVERFLOW

predictive capabilities for wing/winglet installation analyses were

also established.

Future Plans

Further studies of the advanced wing installation effects for

engines of very-high-bypass ratio (up to 12) will be completed

using OVERFLOW to establish the range of applicability.

Keywords

Overset grids, Engine integration

Velocity contours for a complex advanced wing fuselage/nacelle/pylon configuration at transonic conditions;

blue = low velocity and white = high velocity.

Aeronautics/Propulsion 61
ets-in-Cross-Flow Mixing;

Clifford E. Smith, Principal Investigator

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

CFD Research Corporation/NASA Lewis Research Center

Research Objective

To identify improved mixing schemes for rich burn/quick mix/

lean burn (RQL) combustors applicable to advanced supersonic

aircraft engines, and to assess the effects of design parameters on

mixing effectiveness and emission reduction.

Approach

Numerical parametric studies were performed to asses the effect

of varying key design parameters. The computational tool used

was a three-dimensional Navier—Stokes flow solver (CFD-ACE)

designed to analyze turbulent reacting flows in complex geometries. The computational results were examined using the interactive software visualization package, CFD-VIEW.

Accomplishment Description

The CFD-ACE code was used in analyzing over 30 parametric

mixing simulations. The analysis was performed to identify the

effect of orifice aspect ratio and jet mass flow ratio on jet mixing

in rectangular geometries. The geometry modeled consisted of a

rectangular duct with inline rows of jets on the top and bottom

walls. The computations were performed assuming isothermal

and reacting flow conditions. The numerical calculations in this

study used grids varying in size from 50,000 to 100,000 nodes.

An average parametric case took 2-4 Cray C-90 hours and

8-12 megawords of memory.

Significance

The computations show that the slot aspect ratio had little effect

on the jet penetration and overall mixing. Circles and equivalent

area slots (e.g., squares and 4:1 slots) were shown to have similar

mixing characteristics. The jet wake recirculation region

increased in size as the slot aspect ratio was reduced. Jet mass

flow ratio was shown to significantly affect jet penetration; the

design correlation parameter for optimum mixing (orifice

spacing-to-duct height ratio x the square root of the jet-to-crossflow momentum flux ratio) increased by a factor of two as the jetto-cross-flow mass flow ratio increased from 0.5 to 2.0. These

results were used to design advanced RQL combustors.

Future Plans

Continuing analysis will focus on turbulent reacting flow in jets in

cross flow. The calculations will assess what effect heat release

has on jet mixing. Jet mixing configurations that reduce gaseous

emissions will also be studied.

Keywords

Combustors, Cross flow

Publication

Bain, D. B.; Smith, C. E.; and Holdeman, J. D.: CFD Assessment

of Orifice Aspect Ratio and Mass Flow Ratio on Jet Mixing in

Rectangular Ducts. AIAA Paper 94-0218, Jan. 1994.

Numerical simulation ofjet mixing in a rectangular geometry. Inline square orifices: S/H = 0.425, J = 36, mass flow

ratio = 2.0. Temperature: red = 2,550 K and blue = 950 K.

62 Aeronautics/Propulsion
10 20 30 40

25

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m

Parallel Multidisciplinary Stochastic Optimization

Robert H. Sues, Principal Investigator

Co-investigator: Graham S. Rhodes

Applied Research Associates, Inc.

Research Objective

To develop a general purpose parallel computing environment for

large-scale multidisciplinary stochastic optimization (MSO) problems encountered in the design of advanced, high-performance

aerospace systems.

Approach

A problem was formulated on the aerodynamic design of an

advanced propfan blade. A single-processor code representing

key components of the MSO methodology was developed. Probabilistic analysis was performed using Monte Carlo simulation.

The finite-element code, NIKE3D, was used for structural analysis, and a linear-potential panel-method code was used for aerodynamic analysis. The Automated Design Synthesis library was

used for optimization. Parallel implementations of the singleprocessor code were developed using the parallel virtual

machine programming language. These implementations

exploited the multiple levels of parallelism inherent in MSO

problems. Performance was evaluated on the 128-processor

Intel iPSC/860 hypercube and on a network of 32 IBM RS/6000

workstations.

Accomplishment Description

The feasibility of effectively parallelizing some of the key

computational elements of MSO problems was demonstrated.

Influence coefficients were computed in parallel to determine

blade surface pressure coefficients. An average analysis required

5 minutes to process (using 32 hypercube processors) and

2 megabytes of memory per processor. High speedups achieved

on the hypercube and workstation networks demonstrate the

portability and parallel processing potential of the methodology.

High speedups were also achieved for the sensitivity coefficient

computations. The effect of multilevel parallel processing was

investigated. Optimization was performed in clusters of processors to compute each sensitivity coefficient, and each processor

in a cluster computed a group of influence coefficients. Speedup

results for single- and multilevel parallel processing of the sensitivity coefficients are shown in the figure. Single-level speedup

results reach a plateau at 20 processors because there are only

20 sensitivity coefficients to compute.

Significance

There are several levels of parallelism in MSO problems that must

be fully exploited to maximize parallel processing performance.

MSO can be applied to high-speed civil transport and advanced

propulsion systems research to reduce design and development

time and test costs.

Future Plans

A fully integrated design system for solving MSO problems on

parallel computers will be developed. The system will integrate

several principal components of parallel processing technology.

Efficient multilevel parallel algorithms will be developed to effectively distribute aerodynamic and structural analysis computations across multiple processors.

Keywords

Turbomachinery, Turbojet, Propfan, Turboprop

Number of processors

Relative speedup results for multilevel parallelism.

Aeronautics/Propulsion 63
Simulation of Scramjet Flow Fields

R. C. Swanson, Principal Investigator

Co-investigators: J. Korte and E. Turkel

NASA Langley Research Center/ICASE

r].

Research Objective

To develop and apply the Langley Research Code for Chemical

Kinetics (LARCK) for the solution of the three-dimensional (3-D)

compressible Navier—Stokes equations for turbulent high-speed

flows with and without chemistry.

Approach

A cell-centered, finite-volume spatial discretization with

upwind-biased differencing was used. A multistage time-stepping

scheme with implicit support was used as a smoother in a multigrid process to obtain stationary solutions and a multiblock capability was included to provide flexibility for treating complex

geometries.

Accomplishment Description

Before considering high-speed flows with complex flows, LARCK

was evaluated for a Mach 14.1 flow over a 3-D compression

ramp at an angle of 24 degrees. The Reynolds number was

72,000, the free-stream temperature was 160 °R, and the wall

temperature was 535 °R. This problem has viscous/inviscid interactions typical of those found in the flow field within the propulsion system of a hypersonic vehicle. In addition, it has separated

flow, which makes convergence to a stationary solution more

difficult. The difficulty is a consequence of significant changes in

the downstream solution with the evolution (in pseudo-time) of

the separation bubble. The figure shows the computed pressure

contours on a mesh of 53 points in the streamwise direction,

41 points in the normal direction, and 33 points in the cross-flow

direction. The leading-edge shock, separation shock, and shock/

shock interaction region are shown and the separation extent is

delineated. A comparison of the predicted wall pressures in the

symmetry plane with experimental data exhibits fairly good

agreement.

Significance

This project provided a capability to effectively compute

high-speed flows with a multigrid/multiblock method and

upwind-biased spatial differencing. The LARCK code was used to

obtain good predictions for flows similar to those occurring in

hypersonic propulsion systems.

Future Plans

The ability of the present 3-D flow solver to efficiently compute

high-speed flows with chemistry will be demonstrated.

Keywords

Hypersonic, Propulsion

Mach 141 1 laminar flow over a compression ramp.

64 Aeronautics/Propulsion
Hydrocarbon Scramjet Combustor

Jong H. Wang, Principal Investigator

Rockwell International, North American Aircraft Division

Research Objective

To validate Navier—Stokes methodology for predicting hydrocarbon scramjet combustor flows.

Approach

The unified solution algorithms code is applied to existing

combustor-flow test cases to validate the code and to provide

direction for future numerical and chemical model development.

Accomplishment Description

A dual-mode direct-connect hydrocarbon scramjet combustor

was numerically investigated. The combustor was first modeled

as a two-dimensional (2-D) flow with the fuel injector opening

adjusted for a correct fuel-air ratio. At a fuel-air ratio of 1 in the

injection zone, the 2-D model underpredicted the measured pressure data by about 50 percent. This deviation indicated that the

mixing for the 2-D numerical model was much less than that for

the physical flow model. A study was also performed to investigate the effects of three-dimensional (3-D) flow. The detailed configuration of the combustor, including all the fuel injectors, was

considered in the numerical model. The grid system consisted of

347,030 points. The model required 19 megawords of memory,

the solutions took 76 Cray C-90 seconds for each iteration, and

the entire process took about 60 hours to converge. The predicted

temperature contours in the combustor are shown in the first figure. The predicted distributions of centerline wall static pressure

are compared with data in the second figure. The present 3-D

numerical model gives predictions that are in good accord with

experimental data.

cient mixing and combustion require an accurate and robust

computational tool.

Future Plans

Based on the 3-D numerical results, an injection model will be

devised for the 2-D computations. The injection model is necessary for promoting mixing in the 2-D numerical model.

Keywords

Propulsion, Scramjet combustor, Combustion

50

40

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° 30

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1v^

20 40 60 80

Distance (in.)

Significance Predicted temperature contours in the scramjet combustor.

Efficient hypersonic combustors are essential to the design of airbreathing hypersonic vehicles. Flow physics complexity and effiComparisons of predictions with data for pressure on port side (top) and starboard side (bottom) combustor walls.

Aeronautics/Propulsion 65
Reusable Launch Vehicle Flow Simulations

Scott Ward, Principal Investigator

Co-investigator: Darren Fricker

McDonnell Douglas Aerospace

r_-

Research Objective

To use state-of-the-art computational techniques to predict the

aerodynamic and aerothermal performance of advanced launch

vehicle configurations including power-on effects.

Approach

The Navier—Stokes code MDNS3D/ULTRA uses a structured

zonal grid approach to discretize the full set of Reynoldsaveraged Navier—Stokes equations and includes a generalized

fully coupled finite-rate chemical kinetics model and a fully coupled two-equation turbulence model. The equations are solved in

a time-dependent manner using second-order-accurate central

differencing in space and an explicit fourth-order-accurate time

advancement algorithm. Total-variation-diminishing dissipation

ensures stability of the central difference scheme.

Accomplishment Description

Analysis of a nonaxisymmetric low-fineness-ratio multi-engine

reusable launch vehicle configuration with and without power

effects was conducted. The complex structure of the fundamentally unsteady base flow field, characterized by large regions of

separation and entrainment into the engine exhaust flows, was

studied. Base pressure and total vehicle force predictions from

perfect gas simulations were compared to the pre-existing ground

test database for a range of different ascent power settings.

Although the large, essentially unsteady regions of separated flow

reduced code convergence rates for most power-on cases, agreement between predictions and test data was good. Typical perfect

gas solutions required approximately 8 Cray C-90 hours and

15 megawords of memory.

Significance

A computational fluid dynamics capability to model the complex

base flow field for a multi-engine reusable launch vehicle was

validated. This capability augments the existing ground test database by allowing a detailed study of underlying flow features not

recorded by test instrumentation. The necessity for further ground

testing may also be diminished.

Future Plans

A hydrogen/oxygen plus air chemical kinetics model will be

implemented and several ascent cases calculated. Results will be

compared to the perfect gas simulations to obtain increments that

may be used to scale ground test data to flight conditions. Aerothermal studies of other mission critical trajectory points will

be undertaken.

Keywords

Computational fluid dynamics, Aerodynamics, Aerothermodynamics, Propulsion, Base flow

Surface pressure contours (red = high pressure and blue = low

pressure) and flow-field streamlines for the Delta Clipper experimental single-stage-to-orbit vehicle during power-on ascent. Freestream Mach number = 0.3, altitude = sea level.

66 Aeronautics/Propulsion
Chimera Domain Decomposition Applied to Turbomachinery Flow

Kurt F. Weber, Principal Investigator

Co-investigator: Dale W. Thoe

General Motors Corporation, Allison Gas Turbine Division

Research Objective

To upgrade the POVERFLOW code for turbomachinery applications, apply it to conventional and unconventional turbomachinery applications, and evaluate its performance and accuracy.

Approach

Domain decomposition was applied using Chimera grid embedding to couple the calculated solutions of local flow fields in gas

turbine fans and compressors. The equations were solved using

Pulliam's diagonal version of the Beam—Warming implicit

approximate factorization algorithm. A composite mesh for complex flow fields was generated using the Chimera grid embedding

technique. The code worked with the output from the multiplegrid data management code PEGSUS, which established the

interpolation between individual grids. The flow field was geometrically decomposed and grids were generated for the individual elements of a turbomachinery flow field that cannot be

handled with a single mesh (blade rows for a compression system

or air feed pipes for an engine test cell). PEGSUS was run to set

up the interpolations for the mesh boundary communication. For

parallel processing, data decomposition was applied and the parallel processors distributed so that the solution on each mesh was

calculated concurrently.

Accomplishment Description

Upgrades to account for source terms due to a constant rotation

rate were completed. Coding to calculate time varying metrics for

a rotating mesh system, and an inlet boundary procedure specifying radial profiles of total pressure, total temperature, and velocity

components were added. Exit boundary procedures and periodic

boundary procedures for 0-grids, C-grids, and H-grids were completed. The most computationally intensive calculation was for

the flow through the six feed pipes and plenum and into the bell

mouth inlet to the test section for an Allison Engine Company Test

Cell 886. The calculation involved 8 grids, 904,000 points, and it

required 116 processors with 8 megabytes of storage each. The

calculation was run for 6,000 iterations and it required the equivalent of fifty 32-node iPSC/860 hours. The accompanying figure

shows particle traces released at the inlets of the six pipes that

feed the plenum supplying the air to the bell mouth inlet. The calculation confirmed the degree of total pressure distortion at the

inlet of the engine test section.

Significance

The use of three-dimensional flow codes for aerodynamic design

improvements in aircraft gas turbine engine compressors and fans

is widespread. There is increasing emphasis on extending the

capability of analyses by coupling previously isolated flow-field

calculations to determine the effect of coupling on component

performance. With cost-effective parallel computers, more complex turbomachinery flow calculations can be used in the routine

design process.

Future Plans

The routine written specifically for turbomachinery and the

accuracy of the code for calculating the flow through transonic

fans and compressors will be verified. Calculated values of overall performance characteristics and blade-to-blade mass averaged

quantities will be compared with component rig test data for

NASA rotor 37 and NASA stage 37.

Keywords

Turbomachinery, Domain decomposition

Particle traces released at the inlets of the six feed pipes supplying

air to an Allison Engine Test Cell 886. The traces show the extent

of mixing and swirl in the test cell plenum.

Aeronautics/Propulsion 67
National Aero-Space Plane Nozzles with External Burning

ShayeYungster, Principal Investigator

Co-investigator: Charles J. Trefny

ICOMP/NASA Lewis Research Center

r— ^

Research Objective

To develop computational fluid dynamics (CFD) prediction

techniques to study the subscale and full-scale performance

of external burning nozzles at transonic flight conditions.

Approach

The Reynolds-averaged Navier—Stokes equations with finite-rate

chemistry were solved using two- and three-dimensional multiblock, fully implicit codes. The spatial discretization was based

on a second-order total variation diminishing scheme and the

LU-SSOR implicit factorization scheme. The study focused on the

effects of external heat addition and introduced a simplified injection and mixing model based on a control volume analysis. Using

this simplified approach, different configurations were analyzed

quickly and inexpensively.

Accomplishment Description

A CFD tool for analyzing single-expansion ramp nozzles with or

without external burning was developed. Extensive code validation studies were conducted for several nozzle configurations.

These configurations included a baseline and an extended cowl

with and without the use of a flame holder. The computed pressure distributions, axial and normal forces, and general flow

features were in very good agreement with experimental data.

The first figure shows a typical pressure coefficient distribution

along the expansion ramp surface. The results were compared

with experimental data given at the centerline and off-centerline

(61 percent semi-width) locations. A parametric study of nozzle

performance as a function of nozzle pressure ratio at different

Mach numbers was completed. A parametric study of nozzle performance as a function of external burning fuel injection pressure

was also completed. The second figure shows Mach number contour plots for two different external burning fuel pressures. One of

the main effects of increasing the fuel pressure is a reduction in

the amount of overexpansion on the initial part of the ramp. An

analysis of the external burning flow field at full scale was completed, and axial and normal forces and Isp estimates were provided for all the cases studied. A typical calculation required

about 2 Cray Y-MP hours and 8 megawords of memory.

Significance

National aero-space plane nozzles are highly overexpanded at

transonic speeds, which results in significant drag increases. One

way to eliminate this drag is to increase the pressure along the

afterbody by using external combustion. This investigation provides a method for determining the subscale and full-scale performance of external burning nozzles at transonic flight conditions.

Future Plans

This study is complete.

Keywords

External burning, Transonic drag reduction

Publication

Yungster, S.; and Trefny, C. J.: Computational Study of SingleExpansion-Ramp Nozzles with External Burning. NASA

TM-106550, April 1994.

C 

_1

—.4 —.2 0 .2 .4 .6 .8 1.0

( x—xr)/L

Pressure coefficient along the expansion ramp: Mach

number = 1.8, nozzle pressure ratio = 8.58, external burning

fuel pressure = 113 psi.

Mach number contours for two external burning fuel pressures

(top = 50 psi and bottom = 250 psi). Mach number ranges from

blue (0.0) to magenta (2.4).

68 Aeronautics/Propulsion
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Calculations of High Performance Rotorcraft

W. J. McCroskey, Principal Investigator

Co-investigators: E. P. N. Duque, S. Ko, J. Ahmad, and A. C. B. Dimanlig

NASA Ames Research Center/U.S. Army Aeroflightdynamics Directorate/Sterling Software Systems/University of California, Davis

r

Research Objective

To compute the viscous, three-dimensional (3-D), unsteady flow

field around advanced helicopters with particular emphasis

placed on the aerodynamics of high-performance rotor blade

tips, vortical wake structure, and rotor—body aerodynamic

interference.

Approach

The unsteady 3-D Euler/Reynolds-averaged Navier—Stokes

equations are solved by embedded structured-grid, unstructuredgrid, and hybrid embedded structured/unstructured methods.

The codes are validated through detailed comparisons with

experimental data.

Accomplishment Description

The Reynolds-averaged thin-layer Navier—Stokes code

OVERFLOW was used to solve the flow field of the Comanche

helicopter in forward flight. The first figure illustrates the surface

pressure distribution over the fuselage and at the main rotor disk.

The main and tail rotors are approximated by an actuator disk

with a constant pressure jump based on a given rotor thrust. Significant progress was made in coupling an unstructured-grid rotor

code to a structured-grid rotor code. In this case, a structured grid

optimized for viscous dominated flows was used for the nearbody flows, whereas the unstructured grid was used for the inviscid flow fields away from the body. The two grid systems were

embedded into each other with the flow quantities passed

between the grid systems by interpolation. The second figure

shows the structured/unstructured grid system for a rotor in hover

with surface pressure contours and pressure along a cross section

of the blade surface. In addition, an iso-surface of vorticity illustrates the structure of the computed wake.

Significance

The Comanche helicopter is the newest Army aircraft. Viscous

fuselage calculations provide detailed flow-field information

unattainable by other methods. These calculations can be used to

help evaluate and improve aircraft performance. The hybrid

structured/unstructured method incorporates the best features of

each technology. Generally, structured grids perform better for

near-body flows because the grid can cluster near the walls without severe time-step limits. The unstructured grids can refine flow

features more readily than the structured grids. This feature potentially allows more accurate wake calculations.

Future Plans

The unsteady flow field of the Comanche, including the rotor hub

and blades in forward flight, will be computed and compared

with future experiments. The structured/unstructured methodology will be validated for hover and forward flight including realistic motion of the blades. Dynamic grid adaptation of the

unstructured mesh will be implemented and demonstrated.

Keywords

Unstructured embedded grids, Tip vortex, Wakes, Solution

adaptation, Unsteady rotor—body interaction

Publications

1. Berry, J. D.; Chaffin, M. S.; and Duque, E. P. N.: Helicopter

Fuselage Aerodynamics Predictions: Navier—Stokes and Panel

Method Solutions and Comparison with Experiment. 1994

AHS Aeromechanics Specialists Conference, San Francisco,

Calif., Jan. 19-21, 1994.

2. Duque, E. P. N.; and Dimanlig, A. C. B.: Navier—Stokes

Simulation of the RAH-66 (Comanche) Helicopter. 1994 AHS

Aeromechanics Specialists Conference, San Francisco, Calif.,

Jan. 19-21, 1994.

3. Duque, E. P. N.: A Structured/Unstructured Embedded Grid

Solver for Helicopter Rotor Flows. AHS 50th Annual Forum,

Washington, D.C., May 11-13, 1994.

Comanche helicopter in forward flight, free-stream Mach

number = 0.26, free-stream Reynolds number = 14.0 million, and

rotor thrust = 0.009.

Rotor in hover using the structured and unstructured grid pressure

contours at the blade and an iso-surface of constant vorticity.

Aeronautics/Rotorcraft 71
Solution-Adaptive Computations of Helicopter Acoustics

Roger C. Strawn, Principal Investigator

Co-investigators: Rupak Biswas and Michael Garceau

U.S. Army Aeroflightdynamics Directorate/NASA Ames Research Center/RIACS/Stanford University

Research Objective

To develop innovative solution-adaptive methods to accurately

convect vorticity and acoustic waves in helicopter aerodynamics

computations. Successful calculations require that numerical dissipation be kept to a minimum.

Approach

An unstructured-grid flow solver for the Fuler equations is used to

model helicopter aerodynamics and acoustics. An initial solution

is obtained on a coarse tetrahedral mesh and local error estimates

determine where additional mesh resolution is required. Mesh

points are locally added or deleted to capture vortices and acoustic waves with an efficient distribution of grid points.

Accomplishment Description

A new three-dimensional mesh coarsening and refinement procedure was developed for tetrahedral meshes. Several demonstration cases show this procedure to be a highly efficient and

effective way to adaptively resolve helicopter flow-field features

in three dimensions. One of these cases is shown in the figure,

where a solution-adaptive surface mesh and pressure contours

are shown for a high-speed hovering rotor. In addition to the high

resolution of the shock wave on the rotor surface, this solution

shows excellent resolution of the acoustic wave that extends radially from the blade tip. This solution required approximately

5 Cray C-90 hours and 50 megawords of memory.

Significance

Noise reduction is one of the major problems facing the designers

of future civilian and military rotorcraft. Accurate prediction of

helicopter noise is a key component in its reduction and control.

Helicopter acoustic signals require extremely fine-grid resolution

away from the rotor blades. Solution-adaptive grids provide a

computationally efficient method for obtaining this grid resolution in the far field.

Future Plans

Future computations will use the newly developed dynamic mesh

adaption scheme to compute rotor tip vortices and acoustic

waves. Development of local error indicators for vortexdominated flow fields is also a goal.

Keywords

Unstructured, Acoustics, Tip vortex, Wakes, Solution adaption,

Euler solver, Tetrahedra

Publications

1 . Biswas, R.; and Strawn, R. C.: A New Procedure for Dynamic

Adaption of Three-Dimensional Unstructured Grids. Appl.

Num. M., vol. 13, 1994, pp. 437-452.

2. Strawn, R. C.; Garceau, M.; and Biswas, R.: Unstructured

Adaptive Mesh Computations of Rotorcraft High-Speed Impulsive Noise. AIAA Paper 93-4359, Oct. 1993.

Solution-adaptive mesh (top) and pressure contours (bottom) near

the tip of a hovering rotor blade. Results are shown in the plane of

the rotor. The hover-tip Mach number = 0.95 and the blade aspect

ratio = 13.71.

72 Aeronautics/Rotorcraft
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Transition in a Highly Disturbed Environment

David E. Ashpis, Principal Investigator

Co-investigator: Philippe R. Spalart

NASA Lewis Research Center/Boeing Commercial Airplane Group

r

eC4

Research Objective

To understand the physical mechanisms of nonlinear stability

and bypass transitions to provide guidance for engineering

model developers.

Approach

Direct numerical simulations of controlled disturbances in the

boundary layer were performed. Spatial simulations were performed using a Spalart's "fringe code." Numerical experiments

simulating single-frequency excitation (two-dimensional (2-D)

vibrating ribbon, harmonic point source) and broadband excitation (blowing pulse) were performed. The disturbances are introduced on the wall by specifying localized suction and blowing.

The flow field, frequency, and wave-number spectra were computed for incompressible Blasius boundary layer flow.

Accomplishment Description

The fringe code was modified to allow specification of

time-dependent wall boundary conditions. The prior series of 2-D

and three-dimensional (3-D) runs were continued for a range of

disturbance frequencies, amplitudes, and Reynolds numbers.

New simulations of harmonic-point-source excitations were performed. Parameters of these simulations matched prior calculations and experiments. The simulations were performed for a

range of excitation amplitudes, from linear to nonlinear. Observa

tions demonstrated the differences between linear and nonlinear

disturbances. Computing resources varied according to the

domain size, grid resolutions, and the length of simulation.

Typical runs required 4-32 megawords of memory and

4-20 Cray Y-MP hours.

Significance

A highly disturbed environment prevails in the gas turbine flow.

The simulations of the associated nonlinear and bypass stability

mechanisms are valuable for developers of models for applied

engineering calculations. The harmonic-point-source simulations

reveal the effects of a single frequency in 3-D flow without the

complexity of a broadband excitation spectrum.

Future Plans

Postprocessing will continue with the data acquired. The effects

of pressure gradients and curvature will be examined. The studies

will be extended to flows with heat transfer.

Keywords

Disturbances, Receptivity, Bypass transition, Stability, Turbulence,

Point source, Excitation

Disturbance energy for harmonic-point-source excitation in a

Blasius boundary layer. Frequency parameter is 60. Excitation

amplitude is 0.0001 of free-stream velocity. The top shows the

X-Z plane and the bottom shows the X-Y plane at the centerline.

Linear response is observed.

Disturbance energy for harmonic-point-source excitation in

Blasius boundary layer. Frequency parameter is 60. Excitation

amplitude is 0.01 of free-stream velocity. The top shows the

X-Z plane and the bottom shows the X-Y plane at the centerline.

Nonlinearities and breakdown to turbulence are evident.

Aeronautics/Turbulence 75
Incompressible Plane Wake Computation Comparisons

Brian J. Cantwell, Principal Investigator

Co-investigators: Nagi N. Mansour, Rolf Sondergaard, and Eric Monsen

Stanford University/NASA Ames Research Center

i^

Research Objective

To gain an improved understanding of the linkage between the

late stages of wake development and properties of the initial flow

field. Topological methods will be developed for revealing the

significant flow-field features contained in large data sets and

universal features will be sought.

Approach

Direct numerical simulations of the three-dimensional (3-D)

incompressible wake using a Fourier spectral algorithm for flows

with one infinite and two periodic directions have been run.

Parameters such as Reynolds number and initial disturbance

fields were varied to study the effects of changes in the early flow

field on the development and structure of later turbulence.

Matching simulations have been run on the Cray C-90 and the

Intel iPSC/860 to compare performance.

Accomplishment Description

A code for direct numerical simulation of planar free shear flows

was run on the Cray C-90 and Intel iPSC/860. Simulation comparison showed excellent agreement. A typical moderate

Reynolds number simulation up to a convective time of 150

required approximately 20 Cray C-90 hours and 12 megabytes of

memory. A matching simulation on the iPSC/860 using the parallel code required approximately ten 32-node hours. Wakes

started from a variety of initial conditions were run on both

machines. The computations showed that (1) the presence of an

oblique disturbance had a significant effect on initial 3-D development in the wake; (2) flow Reynolds number had a strong effect

on the late time growth of the wake; (3) the presence of longer

wavelength disturbances allowed scale changes in the flow; and

(4) the relative phase of initial disturbances had an effect on the

detailed structure of the wake, but only a minor effect on the

growth rate. The simulations studied included incompressible

and compressible isotropic turbulence, sheared turbulence, and

mixing layer flows.

Significance

Improved understanding of turbulence development and the fine

scale motions in inhomogeneous turbulent flows is necessary

before improved subgrid-scale turbulence models can be developed. This project significantly improved the understanding of

the role of subharmonic disturbances on wake growth.

Future Plans

Direct numerical simulation of several high Reynolds number

3-D incompressible planar shear flows will be performed to further study the effect of initial conditions on the late stages of flow

development. Of particular interest is the effect of initial conditions on the fine scale topology of the turbulence.

Keywords

Wake, Turbulence, Parallel

Entropy contours for a vertical cut through a Reynolds number

2,768 wake at a time of 101.0. Contour levels are 0.1 (blue) to 5.0

(red). Free-stream flow is from left to right.

76 Aeronautics/Turbulence
Numerical Simulation of Riblets

Haecheon Choi, Principal Investigator

Co-investigators: Parviz Moin and John Kim

Stanford University/NASA Ames Research Center

7C4

Research Objective

To provide an accurate estimate of drag reduction performance

by riblet configurations supplied by the Boeing Company.

Approach

Direct numerical simulation is used to simulate turbulent flows

over longitudinal riblets. A second-order-accurate finite difference method is used to solve unsteady incompressible

Navier—Stokes equations in generalized coordinate systems.

Accomplishment Description

Direct numerical simulations of turbulent flows over riblets were

performed at Reynolds number 180 based on the flat plate shear

velocity and channel half-width. Two sawtooth riblet configurations (one with and one without a flat valley) in turbulent channel

flow were considered. Both had riblet spacing of s+ = 15, which

is an optimum size for drag reduction. Twelve riblets were

included in the computational domain, and 32 grid points for

each riblet were used in the spanwise direction to accurately predict skin friction on the surface of riblets. The computations

revealed about 9 percent drag reduction by the riblet with the flat

valley and about 3 percent drag reduction by the riblet without

the flat valley. The grid contained 16 x 129 x 384 grid points in

the sreamwise, wall-normal, and spanwise directions. A typical

run required 300 Cray C-90 hours and 10 megawords of memory.

Significance

Direct numerical simulation is an important tool in predicting

complex turbulent flow phenomena occurring in a configuration

of practical interest.

Future Plans

The direct numerical simulation range of applicability is limited

to low Reynolds numbers. Large-eddy simulation and the

dynamic subgrid-scale model will be used to access higher

Reynolds number flows.

Keywords

Drag reduction, Turbulent boundary layer

Publications

1 . Choi, H.; Moin, P.; and Kim, J.: On the Effect of Riblets in Fully

Developed Laminar Channel Flows. Phys. Fluids A., vol. 3,

1991, p. 1,892.

2. Choi, H.; Moin, P.; and Kim, J.: Direct Numerical Simulation

of Turbulent Flow Over Riblets. J. Fluid Mech., vol. 255,

1993, p. 503.

Flow field above a riblet-covered wall. Streamwise vorticity with cross velocity (top) and streamwise velocity (bottom).

Aeronautics/Turbulence 77
Direct and Large-Eddy Simulation of Wake Flows

George E. Karniadakis, Principal Investigator

Co-investigators: D. Newman and R. D. Henderson

Brown University

r- —

Research Objective

To understand and model turbulent wakes formed behind bluff

bodies of arbitrary shape using high-order direct and large-eddy

simulations (LES).

Approach

The numerical simulations were performed using the NEKTAR

code, which is based on a new generation of spectral element

algorithms. NEKTAR uses triangular instead of quadrilateral spectral elements and a hierarchical expansion basis, which produces

well-conditioned matrix systems for high-order polynomial

expansions. In the NEKTAR code, the entire three-dimensional

(3-D) field is decomposed into Galerkin modes in all directions

that can be used to identify which scales are resolved, at what

accuracy, and how the quality of the solution can be improved by

adaptively varying the hierarchical basis. In such a framework it is

much more natural to implement subgrid models for LES. This

adaptive strategy also provides a more efficient approach in direct

numerical simulations (DNS) because its resolution requirements

are optimal. In turn, this implies that a higher Reynolds number

regime can be accurately investigated.

Accomplishment Description

Direct and large-eddy simulations of turbulent wakes were completed for a family of nominally two-dimensional planar and axisymmetric bluff bodies including circular and half-cylinders,

ellipses of different aspect ratios, flat plates, and spheres. Comparisons were performed with results obtained in companion

experiments for vorticity- and velocity-averaged and instantaneous fields. Three topics were addressed in this study: (1) validation of the hybrid DNS-LES approach where the laminar

boundary layer on the body is entirely resolved and the turbulent

wake is modeled at the subgrid level; (2) investigation of spanwise variations of the time-averaged flow in wakes associated

with instabilities of the large-scale turbulence structure; and

(3) simulation of vortex-induced vibrations in 3-D flow past

cylinders and spheres.

Significance

Bluff-body wake flows are perhaps the most common flows in

aeronautical engineering; however, there are no numerical

models for predicting the dynamics of such flows nor even the

average body forces. The current work is an attempt to systematically simulate bluff-body flows and provide low-dimensional predictive models for turbulent wakes.

Future Plans

High-Reynolds number turbulent wake and flow-structure

interactions will be studied.

Keywords

Turbulent wakes, Spectral elements, Simulations

Iso-surfaces of streamwise vorticity for flow over an axially spinning sphere at Reynolds number 300.

78 Aeronautics/Turbulence
Simulation of Complex-Geometry Wall-Bounded Flows

George E. Karniadakis, Principal Investigator

Co-investigators: Catherine H. Crawford and Ronald D. Henderson

Brown University

Research Objective Keywords

To correlate turbulent structures in the near-wall region of Near-wall turbulence, Drag reduction, Spectral elements,

complex-geometry flows with drag-reducing and drag-inducing Simulations

flow events using high-order direct and large-eddy simulations.

Approach

Direct numerical simulations of turbulent flow over ribletmounted surfaces were performed using an unstructured spectral

element flow solver. This high-order flow solver used nonconforming spectral elements so that a large degree of mesh adaptivity and mesh refinement resolved the small-scale structures of the

turbulence, especially the flow features near the riblet tips. The

first figure shows a channel domain with its lower wall mounted

with V-shaped grooves or riblets. The riblet geometry was chosen

to correspond with available experimental data. The channel

dimensions and mesh structure were chosen after a study of the

resolution requirements for a spectral element simulation for a

channel with two smooth walls.

Accomplishment Description

Direct numerical simulations of turbulent flow over riblets at a

moderate Reynolds number were completed for both conforming

and nonconforming riblet meshes. Turbulence statistics for both

the lower riblet wall and the upper smooth wall were collected

and averaged over 250 convective units. The riblets for this case

did not reduce or induce drag at the wall. Because the riblet wall

had over twice the wetted area as the smooth wall, there was a

reduction in the skin friction. The turbulence intensity profiles

and Reynolds stress profiles showed a slight reduction of the

streamwise intensity and Reynolds stress at the riblet wall. The

statistics from the smooth wall were in excellent agreement with

previous numerical and experimental results. A quadrant analysis

study was done to understand how small-scale turbulent structures were altered by the presence of a riblet wall. The riblets

seemed to redistribute the instantaneous velocity and vorticity

at about y+ = 10 at the riblet wall compared to y+ = 10 at the

smooth wall. Also, the drag-enhancing events for both walls were

correlated using instantaneous wall-shear distributions and veloc

ity vector plots. The near-wall sweep events caused an instantaneous increase in the wall shear, while the near-wall ejection

events caused an instantaneous decrease in the wall shear.

Significance

Although most researchers now agree that riblets of the correct

height and span geometries do cause drag reduction, there exists

no consistent explanation of why this happens. Through studies

of the velocity and vorticity quadrant analysis and the identification of flow structures associated with increased or decreased

drag, an explanation for riblet drag reduction and the drag reduction found in other near-wall flows will be found.

Future Plans

Higher Reynolds number turbulent flows and different riblet

geometries will be studied. The hypothesis that the drag reduction

could be caused by flow randomization will be validated.

Streamwise velocity contours in the riblet domain. Note the

unstructured mesh at the riblet wall.

Close-up of streamwise velocity contours at the riblet wall.

Aeronautics/Turbulence 79
Turbulence Modeling for Three-Dimensional Flow Fields

Linda D. Kral, Principal Investigator

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

McDonnell Douglas Corporation

Research Objective

To investigate turbulence models used in numerical simulations

of complex viscous flows for critical aircraft components such as

the forebody, wing, nozzle, inlet, diffuser, and afterbody. The

focus is on computations of turbulent, three-dimensional (3-D)

flows using one-equation and two-equation turbulence models.

Approach

The 3-D zonal Navier-Stokes code, NASTD, was used to predict

the turbulent flow fields around flight vehicles. An implicit,

approximately factored, upwind scheme is employed to solve the

3-D, compressible, Favre-averaged Navier-Stokes and energy

equations together with the k,c two-equation turbulence model.

Three zero-equation algebraic models and two one-equation

models are available in the code. Six low-Reynolds number forms

of the k-c turbulence model are available. The Wilcox k-w model

and the Menter baseline and shear-stress transport-blended

k-(o/k-e models were included for evaluation. The Sarkar compressibility correction for free-shear layers was also employed.

Accomplishment Description

A hot-gas-exhaust mixing system was successfully analyzed using

the Baldwin-Barth one-equation turbulence model. The analysis

started from the hot engine exhaust, included ejector air at the

turning duct inflow, continued through the mixing flanges,

and out to the end of the tail boom where a back pressure was

specified. Surface static temperatures of the mixer assembly and

downstream duct are shown in the first figure. Prediction of

downstream core temperature agreed with empirical predictions,

and calculated surface temperatures indicate adequate mixing of

the exhaust and bypass gases (second figure). The solution indicates there is significant cross flow near the approach to the

mixing flanges, which could be minimized by proper vane placement. The exhaust mixer system calculation required 2.5 million

points in 16 zones and nearly 20 Cray C-90 hours.

Significance

The data used to develop turbulence models were obtained from

canonical flows and it is not clear how these turbulence models

will perform for more realistic geometries. Successful simulation

of turbulent flows of the complexity found on aircraft will be a

valuable design tool.

Future Plans

The Spalart-Allmaras one-equation model, the k-£ low-Reynolds

number models of So and Huang-Coakley, and the Menter

blended two-equation model will be evaluated. Corrections for

curvature and compressibility will also be evaluated.

Keywords

k-c, One-equation, Two-equation

Surface static temperature of the mixer ,issembly and downstream duct region; red = high temperatures and blue =

low temperatures.

Temperature contours of the mixer showing the quick mixing

present as the exhaust and fan air come together aft of the mixing

flanges; red = high temperatures and blue = low temperatures.

80 Aeronautics/TurbuIence
Large-Eddy Simulation of jet Combustors

Thomas S. Lund, Principal Investigator

Co-investigators: Parviz Moin and Knut Akselvoll

Stanford University/NASA Ames Research Center

Research Objective

To study mixing of fuel and air in a coaxial jet combustor to better

understand the mechanisms responsible for lean blowout.

Approach

A passive scalar was used to trace the mixing of fuel and air in

nonreacting jet combustors. Reynolds numbers on the order

of 104 were considered and corresponded to those numbers

achieved in laboratory experiments. Because of the relatively

high Reynolds number, the large-eddy simulation (LES) approach

was adopted. The dynamic subgrid scale model accounts for the

effect of the unresolved turbulent motions. The code is based on

finite differences and is second-order accurate in space and time.

Accomplishment Description

First, LES was used to calculate turbulent flow over a backward

facing step to assess the accuracy and suitability of the dynamic

subgrid scale model in a complex flow geometrically similar to

the jet combustor. Simulations of the backward facing step were

performed at Reynolds numbers of 5,100 and 26,000, based on

step height and inlet free-stream velocity. For the low Reynolds

number case, the mesh contained 192 x 48 x 32 points in the

streamwise, wall-normal, and spanwise direction, respectively. A

typical run required about 20 Cray C-90 hours and used about

2.5 megawords of memory. For the high Reynolds number case,

244 x 96 x 96 grid points were used, and about 70 Cray C-90

hours and 9 megawords of core memory were required. Both

simulations are in excellent agreement with experimental data.

The low Reynolds number simulation also agrees well with direct

numerical simulation data. The second part of the study involved

the simulation of mixing in a planar jet combustor. The geometry

differs from the coaxial jet combustor only in that it is planar

rather than axisymmetric. This part of the study was undertaken

to evaluate the dynamic subgrid-scale model for the residual scalar flux and to gain some insight into the mixing characteristics

of jet combustors. The grid contained 192 x 76 x 16 points.

A typical run required about 20 Cray C-90 hours and about

2 megawords of memory. To prepare for the coaxial combustor

simulation, an axisymmetric code was written and validated for

turbulent pipe flow.

Significance

The LES technique, used in conjunction with the dynamic subgrid

scale model, can accurately predict a complex turbulent flow at

low and high Reynolds numbers. This technique also successfully

describes the convection and diffusion of a passive scalar.

Future Plans

The coaxial jet combustor at Reynolds number 35,000 will be

simulated and the mixing of fuel and air will be studied. Mechanisms responsible for lean blowout will be identified and control

strategies aimed at preventing this condition will be investigated.

Keywords

Backward facing step, Scalar transport

Publication

Akselvoll, K.; and Moin, P.: Application of the Dynamic

Localization Model to Large Eddy Simulation of Turbulent

Flow Over a Backward Facing Step. Engineering Applications

of Large Eddy Simulations, 1993, S. A. Ragab and U. Piomelli,

eds., ASME Fluids Engineering Conference, Washington, D.C.,

June 20-24, 1993.

Passive scalar contours show mixing of fuel and air in the planar combustor; red = fuel,

blue = air. Three times are shown from earliest (top) and latest (bottom).

Aeronautics/Turbulence 81
Boundary-Layer Transition on a Heated Flat Plate

Nateri K. Madavan, Principal Investigator

MCAT Institute

Research Objective

To conduct direct numerical simulations of boundary layer transition on a heated flat plate with elevated free-stream turbulence.

Approach

The present approach uses a high-order-accurate, upwind-biased,

finite-difference technique in conjunction with an iterativeimplicit time advancement scheme to solve the unsteady

Navier—Stokes equations in nonconservative form. The iterative

time advancement ensures that the fully implicit form of the

equations are solved at each time step. Factorization and linearization errors are thus driven to zero at each time step. Zonal

methodologies are used to allow selected regions of the flow field

to be resolved more accurately. The methods being used can be

extended to general geometries.

Accomplishment Description

An existing computer code was modified for this effort. A coarsegrid computation (16 million grid points) was completed and the

results were encouraging. The coarse-grid computation was

obtained in stages, starting from an initial solution obtained on an

extremely coarse grid and then interpolating to successively finer

grids. This interpolation helped reduce the computation time. For

each grid point, the code requires 6 microseconds of singleprocessor Cray C-90 time for each iteration.

Significance

When completed, this work will provide valuable insight into the

physics of transition and turbulence in high free-stream disturbance environments. This work has a wide range of applications,

including gas turbine cooling system design, and it lays the foundation for future efforts toward direct or large-eddy simulations in

complex configurations at more realistic Reynolds numbers.

Future Plans

After a larger statistical sample is acquired on the present

grid, computations will be initiated on a finer grid (24 million

grid points).

Keywords

Direct simulation, Turbulence

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350,000 and 550,000.

82 Aeronautics/Turbulence
Direct Numerical Simulation of Shock/Turbulence Interaction

Parviz Moin, Principal Investigator

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

Stanford University

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Research Objective

To guide turbulence model development for the interaction of

shock waves with turbulent shear flows.

Approach

The three-dimensional Navier—Stokes equations were numerically solved. The sixth-order Pade scheme was used to compute

spatial derivatives and the third-order Runge—Kutta scheme was

used for time advancement. The shock wave could be treated

either through resolution of its thickness using a nonuniform

mesh or by using a high-order shock-capturing scheme. Turbulent

fluctuations were specified at the inflow boundary, nonreflecting

boundary conditions were specified at the outflow and normal

boundaries, and the spanwise direction was homogeneous.

Accomplishment Description

The interaction of a shock wave of mean Mach number 1.2

with a developed shear flow was computed. A mesh of

191 x 65 x 65 was used to resolve both the turbulence and the

thickness of the shock wave. The code ran at 300 MFLOPS with

10 Cray C-90 seconds for each time step. The computation

required about 325 Cray C-90 hours, 8 megawords of core memory, and 40 megawords of scratch disk space. The results were

consistent with a linear analysis that predicted large amplification

of turbulent kinetic energy and a drop in the Reynolds shear stress

across the shock wave (a behavior not predicted by current models). The computation also clearly illustrated the importance of

entropy fluctuations in the interaction of shear flows with a shock

wave. Currently, shock capturing is used to compute the interaction of a Mach 2 shock with a shear flow.

Significance

The simulations provide a valuable data base to validate turbulence models as well as to examine the physics involved.

Future plans

Shock capturing will be used to complete computation of

shock wave/shear flow interaction.

Keywords

Turbulence, Shock wave, Shear

Instantaneous Reynolds shear stresses in an X-Z plane. The mean flow is from left to right and the shock wave is at the center of the

domain. Note the presence of more regions of positive shear stress immediately behind the shock wave.

Aeronautics/Turbulence 83
Large-Eddy Simulation of Separated External Flows

Parviz Moin, Principal Investigator

Co-investigators: Thomas S. Lund, Patrick Beaudan, and Haecheon Choi

Stanford University/NASA Ames Research Center

Research Objective

To assess the accuracy of the dynamic subgrid-scale model for the

large-eddy simulation of separated external flows.

Approach

Large-eddy simulations were performed for two external separated flows: a circular cylinder and an airfoil at high angle of

attack. A fifth-order-accurate upwind-biased finite-difference

scheme was used for the circular cylinder, and a second-order

finite-difference scheme was used to simulate the airfoil. For the

cylinder wake, simulation statistics were compared with experimental data to assess the effectiveness of the subgrid-scale turbulence model.

Accomplishment Description

Three large-eddy simulations of the circular cylinder flow were

performed with a dynamic model, a conventional Smagorinsky

model, and a no-subgrid-scale model. The mesh contained

144 x 136 x 48 points in the radial, azimuthal, and spanwise

directions, respectively. The wake was resolved accurately for a

distance of 10 diameters downstream from the cylinder. Each

simulation required approximately 300 CPU hours and about

10 megawords of core memory on the Cray C-90. The simulation

with the dynamic model predicts the mean velocity and the random and periodic components of the Reynolds stresses more

accurately than the other two models. The figure shows the timeand spanwise-averaged subgrid-scale eddy viscosity obtained

with the dynamic and conventional Smagorinsky model. The

dynamic model predicts a maximum level of eddy viscosity near

the bubble closure, where mean turbulent production is maximum. Correct prediction of the eddy viscosity in the shear layers

is necessary to simulate the transition process accurately, and the

dynamic model simulation is superior in this regard. Simulation

of flow over a NACA 4414 airfoil at Reynolds number 1.6 million

and 12 degrees angle of attack was completed. A closed separation bubble appears near the trailing edge for these conditions.

The grid contains 257 x 70 x 65 points in the streamwise, normal,

and spanwise directions, respectively. A typical simulation takes

50 Cray C-90 hours and requires 26 megawords of core memory.

Significance

Large-eddy simulation in conjunction with the dynamic subgridscale model accurately predicts external flows involving separation. The dynamic model shows a clear advantage over a conventional Smagorinsky model for regions of laminar-turbulent

transition.

Future Plans

Simulations of the airfoil at high angles of attack will continue

and results will be compared with experimental data.

Keywords

Subgrid-scale modeling, Wake flow, Airfoil

Contours of time- and spanwise-averaged subgrid-scale eddy viscosity normalized

by the molecular value, dynamic model (top) and conventional Smagorinsky model

(bottom).

84 Aeronautics/Turbulence
Large-Eddy Simulation of Complex Flows

Parviz Moin, Principal Investigator

Co-investigators: Thomas S. Lund and Hans J. Kaltenbach

Stanford University/NASA Ames Research Center

Research Objective

To evaluate the technique of large-eddy simulation in conjunction

with the dynamic subgrid-scale model for complex turbulent

flows where Reynolds-averaged techniques are subject to

considerable error.

Approach

Simulations were performed for two complex turbulent flows: the

boundary layer flow over a concave wall and the separated flow

in a diffuser. The incompressible Navier—Stokes equations are

solved in a curvilinear coordinate system using second-order

finite differences. The time advancement is fully implicit and is

also second order. The dynamic subgrid-scale model is used to

account for the effect of the unresolved turbulent motions.

Accomplishment Description

A large-eddy simulation was conducted where a flat-plate boundary layer encountered a 90-degree constant radius of curvature

bend. The Reynolds number based on the momentum thickness

of the boundary layer at the start of the curve was 1,200 and the

radius of curvature was 18 boundary layer thicknesses. Streamwise Taylor—Gortler vortices induced by the convex curvature

were found to significantly alter the mean velocity and turbulent

stress profiles and resulted in a 35 percent increase in skin friction. Comparisons with experimental data indicate that the simulation provides an accurate prediction of this flow and shows an

improvement over the results of a full Reynolds stress transport

model. The mesh contained 178 x 40 x 64 points in the streamwise, wall-normal, and spanwise directions, respectively. A typical run required about 20 Cray C-90 hours and 16 megawords of

memory. A second simulation of turbulent flow through a onesided, 10-degree plane diffuser was performed. Fully developed

channel flow enters the diffuser at Reynolds number 18,000

based on bulk velocity and inlet channel height. Results show a

significant separation in the rear part of the diffuser and attendant

losses in pressure recovery. The mesh contained 160 x 64 x 64

grid points. A typical run required roughly 40 Cray C-90 hours

and 12 megawords of memory.

Significance

The results of the two simulations indicate that the technique

of large-eddy simulation, used in conjunction with the dynamic

subgrid-scale model, can accurately predict complex turbulent

flows for which Reynolds-averaged techniques are inaccurate.

Future Plans

The relaxation of a boundary layer undergoing transition from

convex curvature to a flat section will be simulated. Reynoldsaveraged models are inaccurate for this flow. Disturbances will

be added to the diffuser flow at the corner in an attempt to control

the separation in the aft portion of the device. The simulations

will be used to determine optimal disturbance waveforms and to

test feedback control strategies.

Keywords

Subgrid-scale modeling, Boundary layer, Diffuser

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Reynolds shear stress profiles for the curved-wall boundary layer.

From left to right: a flat section that is 8 boundary layer thicknesses ahead of the curve, and 15-, 30-, 45-, and 60-degree

stations along the curve. Solid line = large-eddy simulation,

dashed line = full Reynolds stress transport model, dots = experimental data.

Instant spanwise velocity for the diffuser flow.

Aeronautics/Turbulence 85
Unsteady Turbulence Model Simulations

Steven A. Orszag, Principal Investigator

Co-investigator: William S. Flannery

Cambridge Hydrodynamics, Inc.

Research Objective

To assess the importance of developing unsteady turbulence

transport models for application to real-world engineering

problems.

Approach

A very-large-eddy simulation (VLES) model, using turbulence

transport equations based on the renormalized group (RNG)

theory of turbulence and implemented in the commercial code

FLUENT, was used to perform the computations.

Accomplishment Description

The incompressible flow past a simulated compressor trailing

edge was considered as a prototype unsteady turbomachinery

application. The Reynolds number for the flow was 56,400 based

on the diameter of the half-cylinder trailing edge. The inlet was

taken at 10.6 diameters upstream where a symmetrical zero pressure gradient turbulent boundary layer was set to match the

experimental conditions. A grid of 30,000 cells was used. The

time-dependent computations were run until a quasi-steady state

was achieved and typically required 25-30 periods of the final

shedding cycle. Approximately 8 megawords of memory and

about 2 Cray-2 hours were required for each period. The first

figure shows the vortex shedding behavior in the wake. Steady

state computations of the trailing edge failed to predict the

strength of the vortex in the wake and yielded significantly overpredicted wake base pressures. The results were identical to those

obtained with a splitter plate in the wake. In the second figure,

the time-averaged VLES computation gives rise to a strong pressure minimum in the wake, as does the experiment. The computed Strouhal frequency is 0.2 and is in reasonable agreement

with experiment. Also shown in the second figure is a timeaveraged computation using the standard k-e model. Apparently,

excessive turbulent production leads to large-eddy viscosities and

mean wake vortices that are too weak. The base pressure in the

wake is significantly overpredicted by the standard model.

Significance

Unsteady simulations are required to accurately predict the

effects of large-scale time-dependent structures. The RNG VLES

models compute these flows and provide an accurate prediction

of unsteady effects.

Future Plans

Further studies of other unsteady flows are planned to understand

what time-dependent behavior the VLES models capture. In addition to two-dimensional flows, three-dimensional effects on nominally two-dimensional problems will be explored. Finally, some

fully three-dimensional studies will be performed.

Streamlines in the wake of a compressor trailing edge over a shedding cycle (Reynolds number = 56,400).

Time = averaged RNG VLES

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Very-large-eddy simulation, Turbulence transport equations s/t — i - x/t

Comparison of time-averaged distributions of pressure with experiment for the flow past a compressor trailing edge (Reynolds

number = 56,400).

86 Aeronautics/Turbulence
Three-Dimensional Liquid Sloshing Flow Simulation

Richard H. Pletcher, Principal Investigator

Co-investigators: S. Babu, F. J. Kelecy, and W-P. Wang

Iowa State University

r^

Research Objective

To develop accurate and efficient simulation methods for flows

having a free surface, particularly liquid sloshing flows in moving

containers, and to evaluate the predictive capability of large-eddy

simulations (LES) to accurately resolve flows with complexities

such as heat transfer and recirculation.

Approach

Current free-surface research focused on using surface fitting and

surface capturing to compute three-dimensional (3-D) liquid

sloshing flows in moving containers. The incompressible

Navier—Stokes equations were solved using the artificial compressibility method and a fully coupled, implicit discretization in

both schemes. The LES study employed a preconditioned (for

low-Mach number efficiency), fully coupled, upwind finitevolume discretization of the filtered compressible Navier—Stokes

equations in primitive variables. The convective terms were

discretized with third-order quick-type upwinding while fourthorder central differencing was used for the viscous terms.

Accomplishment Description

A new calculation procedure that iteratively approaches secondorder temporal resolution of the free-surface motion was devised.

Results for the 3-D broken dam problem were in good agreement

with experimental data (first figure). Several free-surface flows

were computed using the surface capturing approach, including

the broken dam problem, the Rayleigh—Taylor instability, and

an axisymmetric spin-up in a spherical tank. The current LES

approach was tested on a planar channel flow and was in

reasonable agreement with published results. LES of channel

flow with a low level of heat transfer (with both the uncoupled

passive scalar equation and the coupled fully compressible

energy equation) was accomplished. A typical computation used

250,000 grid points with 50-90 megawords of memory and

25-50 Cray C-90 hours. The second figure shows the instantaneous temperature contours at different spanwise locations for

the case of compressible LES channel flow with low heat transfer

rate and a Mach number of 0.01.

Significance

The ability to numerically simulate the motion of liquids in

partially filled containers and the resulting interactions with structures is important for the design of liquid management systems for

many space applications and more traditional applications. Turbulent flow represents one of the major pacing items restricting

the usefulness of computational fluid dynamics for realistic

design applications. A goal of the LES research is to take advantage of rapidly developing computer technology to advance the

predictive capabilities of the approach toward larger and more

relevant problems.

Future Plans

The surface fitting code will be used to compute flows under

microgravity environments. Work is under way to perform the LES

of variable-property channel flow with massive heat transfer to

study the changes in turbulent statistics.

Keywords

Large-eddy simulation, Free surface, Turbulence

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Comparison of predicted and measured surge front location for

the 3-D broken dam problem.

Instantaneous nondimensionalized temperature contours at three

spanwise locations blue = low temperature and red = high temperature. Flow is from left to right.

Aeronautics/Turbulence 87
Direct Simulation of Turbulent/Transitional Airfoil Flow

Man Mohan Rai, Principal Investigator

NASA Langley Research Center

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Research Objective

To perform direct simulations of turbulent/transitional flow over a

turbine airfoil.

Approach

High-order-accurate, upwind-biased, finite-difference methods

were used to perform direct simulations of compressible

turbulent/transitional flows over complex geometries.

Accomplishment Description

A program to compute turbulent/transitional flows over an airfoil

cross section was developed. This code was used to compute the

flow over a turbine airfoil cross section using a grid with about

20 million points (a grid with 40 million grid points probably will

be required for the final computation). The figure shows instantaneous velocity contours very near the airfoil surface. The figure

shows the isolated patches of vorticity immersed in relatively quiescent regions (near the leading edge). Further downstream there

is an abrupt appearance of vortical structures indicating that the

transition process is essentially complete.

Significance

When complete, this work will be the first direct simulation of

this kind for turbulent/transitional flow over an airfoil cross section. The computed data will be used to understand the transition

process on a turbine airfoil and also to develop improved turbulence and transition models for large-eddy simulations and

Reynolds-averaged Navier—Stokes simulations. The methodology

and computer program developed during this effort will serve

as a basis for future direct simulations of flow over complex

geometries.

Future Plans

This computation will be continued on a grid with approximately

40 million grid points and a statistical sample will be obtained.

Keywords

Direct simulation, Turbulence

Instantaneous velocity contours just above the suction surface (y+ = 1.0) of a turbine airfoil with flow direction from left to

right. The leading edge is at the upper left section of the figure and the trailing edge is at the lower right section of the figure.

88 Aeronautics/Turbulence
Self-Similar Turbulent Plane Wakes

Michael M. Rogers, Principal Investigator

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

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

Research Objective

To study the effects of two-dimensional (2-D) forcing on

structures, statistics, and chemical reactions in incompressible

fully turbulent self-similar plane wakes.

Approach

Direct numerical simulations of temporally evolving incompressible turbulent plane wakes were generated with various levels of

2-D forcing, using previously simulated turbulent boundary layers

to provide a realistic turbulent initial condition. Pseudo-spectral

numerical methods were used to solve the three-dimensional

Navier—Stokes equations.

Accomplishment Description

Two fully turbulent wakes with half-width/velocity-deficit

Reynolds numbers of about 2,000 were simulated. A passive

scalar quantity was also computed to permit the study of "fast"

chemical reactions. In one case, no 2-D forcing (beyond what is

present in the initial boundary layers) was used, whereas the

other case was strongly forced. The unforced case evolved selfsimilarly, with mean velocity and Reynolds stress profiles that collapsed when scaled by the layer thickness (which increases like

the square root of time) and the velocity deficit (which decreases

like the square root of time). The forced case underwent a shorter

period of more approximate self-similarity. The unforced wake

spreads more slowly than experimental wakes, but the forced

case spreads much more rapidly. This increased spreading rate is

associated with a corresponding difference in flow structure.

Whereas the unforced wake appears as a somewhat homogeneous slab of vortical fluid, the forced wake exhibits a more organized large-scale structure that is suggestive of a Karman vortex

street (see figure). The Reynolds stress levels also increase when

forcing is used, but normalizing them by the wake growth rate

improves the agreement between the two cases and experiments.

The simulations require up to 600 x 260 x 160 modes. The two

simulations required a total of about 750 Cray C-90 hours,

10 megawords of core memory, and up to 170 megawords of

solid state device space.

Significance

These simulations are the first turbulent wake direct numerical

simulations to achieve self-similarity. The strong effect of forcing

on the wake growth rate and structure helps explain discrepancies in experimental observations by different investigators.

Future Plans

To numerically simulate turbulent plane wakes subjected to mean

strain in order to provide data that will assist in the development

of models for high-lift airfoils.

Keywords

Direct numerical simulation, Free-shear flows

Spanwise vorticity contours (red = positive, blue = negative) in the self-similar period for

the unforced wake (top) and strongly forced wake (bottom). Views are at typical spanwise

locations, with a horizontal streamwise direction.

Aeronautics/Turbulence 89
Origins of Near-Wall Turbulence

Bart A. Singer, Principal Investigator

Co-investigator: Ronald D. Joslin

High Technology Corporation/NASA Langley Research Center

Research Objective

To identify and understand the mechanisms involved in the

formation and growth of boundary layer turbulence.

Approach

Spatial direct numerical simulations (DNS) of the incompressible

Navier—Stokes equations were used to generate the data for this

study. Data visualization was performed using the flow analysis

software tool.

Accomplishment Description

In the past, hairpin vortices were generated by a variety of

localized wall disturbances. In one case, the hairpin vortex grew,

elongated, and spawned a secondary vortex head upstream of the

primary, and a subsidiary vortex developed underneath the hairpin vortex legs. This study continued the DNS into the early

stages of turbulence. The resolution requirements for the calculation required over 20 million grid points and the memory requirement was approximately 440 megawords. Approximately

1,000 Cray C-90 hours were used in the computation. Although

the mean flow did not develop a log-layer profile, many features

characteristic of fully turbulent flow were observed. Multiple

additional vortices developed in the flow, which generated newer

vortices in the near-wall region. The regeneration of new vortical

structures is a poorly understood feature of turbulent flows. In the

calculations, one particularly common sequence for vortex

regeneration was illustrated. The figure shows activity on a crosssectional plane that cuts through the middle of a large quasistreamwise vortical structure. One long spiraling line on the right

indicates the circulating flow associated with the leg of the large

vortex. This vortex induces a flow of near-wall fluid to the left.

Two streamlines show flow from right to left near the wall. The

fluid cannot continue into the region of higher pressure near

the wall and must separate from the wall. A single streamline

restricted to the plane is shown going from left to right. Where

it meets the streamlines that go from right to left, all of them

lift away from the wall and concentrate in a relatively narrow

band. The rising fluid creates a new vortex to the left of the

ejection zone.

Significance

This work provides a better understanding of the way in which

coherent flow structures evolve from a single localized disturbance into a group of structures that form the early stages of turbulence. It provides clues as to what kinds of structures dominate

the flow and how these structures interact and regenerate, making

the flow more turbulent. These insights will lead to more realistic

models for predicting turbulent and transitional flows.

Future Plans

The calculation will be continued until a fully developed

turbulent boundary layer is obtained. Data will be analyzed to

understand how the various types of vortices form, interact, and

regenerate. The role that vortices play in engulfing new fluid into

the turbulent region will be studied.

Keywords

Turbulence, Transition, Vortices, Turbulent spot

Cross-sectional plane at x = 98.7 and t = 97.2. Insert at left indicates pressure levels,

high pressure is near top and low pressure is near bottom. The black surface on right

is a portion of large vortex that protrudes upstream of the plane. The black lines are

streamlines restricted to plane.

90 Aeronautics/Turbulence
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Page intentionally left blank 
Hybrid Massively Parallel Aeroacoustics Scheme

Lyle N. Long, Principal Investigator

Pennsylvania State University

Research Objective

To analyze sound propagation and radiation in high-bypass-ratio

ducted fans. This research is focused on sound propagation

in nonuniform engine inlets and radiation of that sound to

free space.

Approach

A hybrid scheme that solves the three-dimensional Navier—Stokes

equations in the near field and then uses a moving Kirchhoff surface to predict the far field is implemented. The equations are

solved in a time-accurate manner using a finite-difference

scheme and Runge—Kutta time marching. This code is fourthorder accurate in space and time. The solution from the code is

passed to the Kirchhoff surface routine and works well. Nonreflecting boundary conditions are important in computational

aeroacoustics and also work well. Preliminary results show that

the schemes are effective on the massively parallel computers.

Accomplishment Description

This scheme was validated by comparing steady-state and

unsteady predictions to analytical and experimental data. The

scheme predicted the transonic flow over a transonic airfoil, a

moving and pulsating sphere, and the flow field of a high-bypassratio engine inlet. The code requires more than 1 gigabyte of

memory for each simulation.

Significance

Future commercial aircraft engines will have very high bypass

ratios. The noise from these engines will be difficult to predict

because of the very short ducts. In addition, fan noise will

become a significant noise source. The hybrid scheme in this

research addresses these noise issues.

Future Plans

Turbulence and engine liner models will be incorporated in

the scheme.

Keywords

Computational fluid dynamics, Navier—Stokes, Noise

Publications

1 . Chyczewski, T. S.; and Long, L. N.: An Efficient Higher Order

Accurate Parallel Algorithm for Aeroacoustics Applications.

AIAA Paper 94-2265, June 1994.

2. Ozyoruk, Y.; and Long, L. N.: A Navier—Stokes/Kirchhoff

Method for Noise Radiation from Ducted Fans. AIAA

Paper 94-0462, Jan. 1994.

3. Weinberg, Z.; and Long, L. N.: A Massively Parallel Solution

of the Three-Dimensional Navier—Stokes Equations on

Unstructured, Adaptive Grids. AIAA Paper 94-0760,

Jan. 1994.

Euler solution of a pulsating sphere moving at Mach 0.3.

Navier—Stokes solution of flow field over a General Electric engine

at Mach 0.82.

General/Aeroacoustics 93
Evolution of Planetary Rings

Creon Levit, Principal Investigator

NASA Ames Research Center

Research Objective

To gain insight into the dynamics of the f-ring region of Saturn.

The techniques can be applied to the study of other planetary ring

systems, and asteroidal, cometary, and planetary systems stability.

Approach

A celestial mechanics algorithm was implemented that integrates

the trajectories of a small number of mutually interacting massive

perturbers with a large number of massless test particles in orbit

around a large central body. The algorithm is based on a technique that assumes near-Keplerian orbits for all of the bodies.

Collision detection, a nonspherical central force field, and

smooth initiation of tidal acceleration are interesting features

of the code.

Accomplishment Description

The code was implemented, calibrated, and put into production.

Several simulations of the evolution of Saturn's f-ring region

were run with different initial conditions for the test particles

beyond a million orbits. Complete data for all bodies were saved

every 500 orbits and analyzed graphically. Approximately

200 CM-5 hours were used for the simulations.

Significance

The detailed structure of Saturn's rings still contains many structures whose origins are a mystery. The pattern of ringlets in

Saturn's f-ring or the overall structure of the b-ring can be simulated and analyzed with the code, as can Neptune's arc-like rings.

The code is also being used to simulate the orbital stability of

planets around binary stars.

Future Plans

Dynamics of the f-ring region will be simulated in more detail by

including additional perturbers, a wider variety of initial conditions, and dissipative forces on the test particles. The dynamics of

the b-ring region will be simulated. Stability of planetary orbits in

hypothetical solar systems will be tested.

Keywords

Gravitation, Dynamics, Solar system

(a)

(b)

Distribution of bodies in Saturn's f-ring region (0.10755E +

07 orbits). Surviving test particles are cyan, and Saturnian moons

Prometheus and Pandora are red. (a) Eccentricity distribution

(note the resonance at the far right and far left). (b) Physical distribution (note the lanes cleared by collisions and the thin bands of

particles in horseshoe orbits).

94 General/Astronautics
Research Objective 10-4

To construct an incident versus detected energy response matrix

of the cosmic x-ray experiment (CXE) high-energy detector 1

(HED1), and to use this response matrix to examine the spectral 10-5 measurement of the cosmic x-ray background from the CXE.

Approach

A Monte Carlo simulation of the HED1 gas proportional counter

T

j

a) 10-6

was applied. Simulation programs were parallelized and used a

high-quality random number generator with very large periods. V)

The simulation code included the photon and electron interac- 0 10-7

tion physics, the atomic rearrangement physics following active m xenon gas ionizations, the mechanical copper collimator, and the cc

detector housing and shielding. The photons were isotropically 10-8 incident on the top of the collimator. The CXE detector consisted

of two layers of two alternating detection cell types that were

below two different mechanically collimated fields of view

(FOV). Taking the difference between signals removes many 10-9

inherent detector backgrounds. 0 10 20 30 40 50

Detected energy (keV)

10 20 30 40 50

Simulating the Cosmic X-Ray Experiment

James Caristi, Principal Investigator

Co-investigators: Brian Ramsey, Martin Weisskopf, and Jeff Youngen

Valparaiso University/NASA Marshall Space Flight Center

Accomplishment Description Detector response for 50 keV incident photons for the large

Over 40 billion photons have been simulated in runs of mono- (thin line) and small (thick line) FOV detection cells in the first

energetic incident photons from 5 to 85 keV (in 5 keV steps). The layer of HED1.

simulations indicate that effects typical at energies above 25 keV

may not have been fully accounted for in the original analysis.

Significance 10- 4

The CXE is the only experiment that has measured the spectrum

of the cosmic x-ray background. A response matrix different from

that used in the original analysis will change the measured spectrum, which will have a profound effect on the cosmological 10-5

models currently employed to produce the spectrum of the x-ray

background. Y 10-6

Future Plans

Higher fidelity simulations in excess of 100 billion photons will iiv

be run with a continuous incident spectrum from 3 to 100 keV.

a 10-7 The high-fidelity simulations will be used to construct a new (n

detailed response matrix for the CXE HED1 detector. a)

Er

Keywords 10-8

Proportional counters, Instrumental response function

10-9

Detected energy (keV)

Response difference between the large (thin line) and small (thick

line) FOV detection cells in the first layer of HED 1.

General/Astronomy 95
Studies of Galaxy Formation and Fvolution

Bruce F. Smith, Principal Investigator

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

NASA Ames Research Center/University of Chicago

r^

-7

Research Objective

To use numerical experiments to understand the important physical processes at work in the formation and evolution of galaxies.

Approach

The development of particle and gas dynamics codes was

exploited to investigate the dynamics of galaxy models and interactions of galaxies in clusters of galaxies. The codes allow

the investigator to follow the time development of fully selfconsistent, three-dimensional galaxy representations. The codes

follow up to 107 particles representing both stars and the gas in

the interstellar medium. Forces and potentials are computed on a

2563 grid.

Accomplishment Description

The particle codes were used to extend earlier experiments that

indicated that galaxies are not in a steady state. Global and local

modes of oscillations in galaxy centers were identified. These

results led to the consideration of the long-term behavior of galaxy oscillations. The long-term survival of multiple galactic

nuclei was investigated using the Cray C-90. Galaxy interactions

were studied to understand the stellar and gaseous responses in a

collision. These experiments used a disk galaxy model composed

of stars and gas embedded in a dark matter halo.

Significance

Observations of galaxies provide improved higher spatial and

spectral resolution. The experiments, significantly improved, are

needed to understand observations from the Hubble Space Telescope and other observational instruments. Recent observations

indicate complex galaxy nuclei that are also indicated in the

numerical experiments. The experiments on the collision of a disk

galaxy with another galaxy will help in understanding the star formation processes in starburst galaxies.

Future Plans

The star formation process in interacting galaxies and galaxies in

general will be investigated. This requires a more sophisticated

treatment of the interstellar medium in a galaxy. The dynamic

evolution of disk galaxies composed of both stars and galaxies

will also be investigated.

Keywords

Dynamics, Stellar systems

The development of a ring structure (bottom left) in a disk galaxy after a collision with an

elliptical galaxy (top right).

96 General/Astronomy
Intraseasonal Atmospheric Oscillations

Christian L. Keppenne, Principal Investigator

Jet Propulsion Laboratory

Research Objective

To use a global multilevel atmospheric model to investigate

intraseasonal atmospheric oscillations in the Northern

Hemisphere.

t

Approach

A global multilevel spectral-transform model of Earth's

atmosphere was run for 18 years at a resolution equivalent

to a grid point model on a 2° longitude x 2° latitude grid. The

data corresponding to the last 15 years of the run history were

analyzed using a variant of empirical orthogonal function analysis in time and space. The results were compared with results

from observational studies.

Accomplishment Description

The intraseasonal variability of the model is dominated by

3 oscillations with average periods near 70, 40, and 25 days.

These periods are also present in a data set of observed atmospheric angular momentum (AAM). The spatial variability associated with the three oscillations is examined by compositing the

stream function-anomaly fields of the model. As was found in

observational studies, the resulting composites show that the

40-day oscillation is dominated by a standing zonal wavenumber-two pattern over the Northern Hemisphere topography.

Substantial mass transfer occurs between this midlatitude, wavenumber-two component of the 40-day oscillation and a zonally

symmetric, tropical component. The 70-day oscillation is associated with mixed wave-number-one to -three patterns and—even

though a similar oscillation is present in a single-layer version of

the model—has a pronounced vertical shear at all latitudes. The

25-day oscillation is reminiscent (in some aspects) of the

westward-traveling wave studied by Branstator and Kushnir.

It also exhibits noticeable shear in the tropics and is absent from

the single-layer model. Phase relationships between adaptively

filtered time series of the AAM for both layers over the cycle of

each oscillation are examined as well.

Significance

Our findings suggest that the 70-, 40-, and 25-day intraseasonal

oscillations detected in observational studies are associated with

non-zonal instabilities of the Northern Hemisphere jet caused by

its interaction with the mountain field.

Future Plans

The bifurcation structure of the model will be analyzed in detail

to identify the mechanisms that lead to non-zonal instabilities.

Keywords

Climate, Subannual variability

Opp _

A `

Model composite maps of the stream function-anomaly field at

the 500 millibar level showing four stages of the dominant 40-day

Northern Hemisphere oscillation. From top to bottom: phases

associated with high, decreasing, low, and increasing levels of

AAM.

General/Atmospheric Science 97
Dynamics of the Martian Atmosphere

James B. Pollack, Principal Investigator

Co-investigator: Robert Haberle

NASA Ames Research Center

i10

Research Objective

To simulate the present and past climate regimes on Mars. The

primary tool used is a general circulation model (GCM) of the

Martian atmosphere that predicts the time-evolving threedimensional wind and temperature fields, the surface pressure,

temperature, and carbon dioxide (CO2 ) ice abundance fields, and

atmospheric water vapor and dust abundance fields.

Approach

The Mars GCM is based upon the primitive equations of meteorology. Physics unique to Mars, including condensation and sublimation of the COZ (the main atmospheric constituent) are

included. The Mars GCM is interfaced with an aerosol physics

model to study dust transport and the impact of the absorption of

sunlight by dust on temperatures and winds. Water vapor transport and its sublimation and condensation are also included in

the model.

Accomplishment Description

A complete year (669 Martian days) was simulated. This

simulation allowed investigation of the radiative properties

(albedo, emissivity) of the COZ ice in the polar regions and their

effects upon the annual atmospheric mass (CO2 ) cycle. Calculated surface pressures at model grid points closely corresponding

to the Viking lander locations favorably compared with the

observed pressures (see figure). A complete simulation required

12 Cray C-90 hours, 3 megawords of memory, and generated

12 gigabytes of output for analysis.

Significance

The Martian atmosphere is the best solar system analogue for the

terrestrial atmosphere. Application of similar models to both

atmospheres and analyses of their results illustrate similarities and

differences between the two environments. On Earth, condensation and sublimation of water are the important energy sources

and sinks of heat to the atmosphere. On Mars, which is much

more arid, the absorption and emission of radiant energy by dust

is the important source of atmospheric heating.

Future Plans

The ability to accurately simulate the Martian atmospheric circulation will be improved to come closer to the ultimate goal of

having a single model to study the interactive nature of the circulation and the annual and multi-annual cycles of Mars' geophysically important COZ , dust, and water cycles.

Keywords

General circulation model, Martian atmosphere

Le

90 120 150 180 210 240 270 300 330 360 30 60

10.0 6 ! l

i

9.5

9.0

cn 8.5

^a

8.0

7.5

7.0

6.5 i i 1 1 1 1 1 1 1 1 1 1 1 1

0 50 100 150 200 250 300 350 400 450 500 550 600 650

Time (day)

Daily averaged surface pressure (millibars) measured at the Viking lander 1 site (dashed line) and

calculated at the model equivalent location (solid line) from an annual simulation representative of

the first Viking year. Le is a seasonal indicator, with Le = 90 at northern summer solstice, Le = 180

at northern autumn equinox, Le = 270 at northern winter solstice, and Le = 360 at northern spring

equinox.

98 General/Atmospheric Science
99

Simulation of Volcanic Aerosol Clouds

Richard E. Young, Principal Investigator

Co-investigators: O. B. Toon and H. Houben

NASA Ames Research Center

rc^

Research Objective

To numerically simulate the behavior of the El Chichon and

Mt. Pinatubo volcanic aerosol clouds in the stratosphere to better

understand stratospheric transport and aerosol microphysical

processes. An associated goal is to assess the climatic impact

of such large volcanic eruptions on stratospheric wind and

temperature fields.

Approach

A three-dimensional (3-D), spectral, primitive-equation model is

used to compute wind and temperature fields in the stratosphere.

This model is coupled with a 3-D aerosol transport and microphysical model that computes the dispersion of the volcanic aerosol cloud using the computed winds from the circulation model.

The model accounts for sedimentation and coagulation of the

aerosol particles.

Accomplishment Description

Work concentrated on simulating the behavior of the El Chichon

volcanic cloud and comparing it to the behavior of the

Mt. Pinatubo cloud. Prior to the Mt. Pinatubo eruption in

June 1991, El Chichon represented the largest volcanic eruption

this century (April 1982). In terms of dispersion in the stratosphere, the two volcanic clouds behaved quite differently. A

series of 2-3 month simulations indicate that most of the difference in behavior between the volcanic clouds is due to interannual variability in stratospheric wind fields. The difference in

cloud absorption of radiative energy from the troposphere also

contributes to the difference. A 3-month simulation requires

about 4.5 Cray C-90 hours and 18 megawords of memory. Sinnilation results are generally in good agreement with the extensive

data sets that exist for both aerosol clouds. The figure compares

the dispersion of both clouds 21 days after each volcano erupted.

Significance

The principal goal of several major NASA programs is the study of

the stratospheric climatology. Further, the High-Speed Research

Program is sponsoring research to assess the climatic impact of

high-speed aircraft operating in the lower stratosphere. Unique

opportunities to better understand stratospheric transport and

aerosol microphysical processes occurred when the El Chichon

and Mt. Pinatubo volcanoes erupted. By simulating the behavior

of the volcanic clouds theoretical understanding of transport

and aerosol processes in the stratosphere can be assessed

and improved.

Future Plans

Multi-year simulations of the Mt. Pinatubo and El Chichon volcanic aerosol clouds will be conducted using passive tracer and

interactive modes. Climatic feedback about the stratosphere will

be assessed.

Keywords

Stratospheric transport, Climate

Publication

Young, R. E.; Houben, H.; and Toon, O. B.: Radiatively Forced

Dispersion of the Mt. Pinatubo Volcanic Cloud and Induced Temperature Perturbations in the Stratosphere During the First Few

Months Following the Eruption. Geophys. Res. Lett., vol. 21,

1994, p. 369.
Chemically Reacting Compressible Shear Layer

Sang-Wook Kim, Principal Investigator

University of Toledo

Research Objective

To develop a numerical analysis and simulation method that can

resolve advanced combustion physics such as the ignition delay,

flame extinction, and combustion instability occurring in

aeropropulsion systems.

Approach

The Favre-averaged compressible flow equations and the

convection-diffusion equations for the chemical species are

solved by a finite volume method that incorporates an incremental pressure equation for the conservation of mass. The chemical

reactions are described by 9 chemical species and 24 pairs

of reaction steps. The turbulence field is described by multipletime-scale turbulence equations (M-S equations). The capability

of the integrated analysis method to resolve ignition delay

and flame thickness is validated by solving hydrogen stream

and vitiated supersonic airstream mixing with and without

chemical reactions.

Accomplishment Description

The contour plot of the combustion trace species shows that

combustion is initiated approximately 18.5 cm downstream of

the inlet boundary; the data measured using ultraviolet radiation

show that combustion is initiated 18.0 cm downstream. The

excellent agreement between the measured and the calculated

flame locations indicates that the M-S equations can correctly

resolve the ignition delay. The ignition delay is caused by the

time required for the hydrogen species to be heated until it attains

the activation energy. The good agreement indicates that the

present analysis method accurately describes the molecular and

turbulent heat transfer occurring between the cold hydrogen

stream and the vitiated airstream. The calculated species concentrations obtained using the M-S equations are in closer agreement

with the measured data than those obtained using a probability

density function method and k--E turbulence models. Also, the

M-S equations successfully predict the increased shear layer

thickness caused by the chemical reaction–turbulence interaction. Calculation of the reacting flow case using 145 x 140 mesh

requires 1.5 Cray C-90 hours and 2.9 megawords of memory.

Significance

The present method can resolve such combustion physics as

ignition delay and flame thickness. The accomplishment is an

important step toward the development of a numerical method

that can correctly resolve local flame extinction in unsteady

turbulent flows and combustion instability for theoretical analysis

of advanced aeropropulsion systems.

Future Plans

The analysis method for reacting flows near flame extinction

conditions, simulations of unsteady, chemically reacting turbulent

flows, and applications of the instability analysis method will

be validated.

Keywords

Navier–Stokes, Combustion, Turbulence, Ignition delay,

Diffusion flame

Publications

1. Kim, S-W.: Numerical Investigation of the Influence of

Chemical Reaction on Turbulence Fields. AIAA

Paper 94-0650, 1994.

2. Kim, S-W.; and Benson, T. J.: Fluid Flow of a Row of Jets in

Crossflow—A Numerical Study. AIAA J., vol. 31, no. 5, 1993,

pp. 806-81 1 .

3. Kim, S-W.; and Benson, T. J.: Calculation of a Circular Jet in

Crossflow with a Multiple-Time-Scale Turbulence Model. Int.

J. Heat, vol. 35, no. 10, 1992, pp. 2,357-2,365.

Contour plot of hydroxyl combustion trace species for reacting case.

100 General/Chemistry
(c)

3 = o

E

ay

I

(d)

Implementation of NAS Parallel Benchmark SP.

Automated Instrumentation and Monitoring System

Jerry Yan, Principal Investigator

Co-investigators: Pankaj Mehra, Sekhar Sarukkai, Cathy Schulbach, Melisa Schmidt, and Brian Vanvoorst

Recom Technologies/NASA Ames Research Center

Research Objective

To develop scalable and portable performance evaluation tools

for parallel applications.

Approach

High-Performance Computing and Communication Program/

Computational Aerosciences grand challenge applications are

some of the most complex parallel programs ever developed.

Program developers and users need to know what computations

or communications are the dominant contributors to completion

time. The Automated Instrumentation and Monitoring System

(AIMS) measures and analyzes parallel program performance.

AIMS instruments and monitors message passing programs, animates execution traces, produces execution-time profiles, and

discovers and characterizes performance bottlenecks. The execution traces created by AIMS can be easily converted to formats

used by other analysis or visualization tools.

Accomplishment Description

The AIMS 3.0 instrumentor is more robust than previous versions

The newly designed monitor provides two levels of tracing

depending on the users need for detail and storage availability.

Communication costs for various data arrays are measured automatically. Prototypes of the AIMS index kernel and statistics

kernel rapidly characterize program performance.

Significance

AIMS is used for tuning parallel and distributed message passing

programs.

Future Plans

Scalable methodologies for data collection and analysis are

being developed. AIMS soon will be ported to the IBM SP-2

environment.

Keywords

Performance evaluation, Parallel processing, Instrumentation,

Performance monitoring and visualization

IL

= v

(a)

(b)

Traces gathered from an interdisciplinary application.

x

_ a

General/Computer Science 101
Computational Fluid Dynamics Solution for Maxwell's Equations

Ramesh K. Agarwal, Principal Investigator

Co-investigators: Dau-Sing Wang and Mark R. Axe

McDonnell Douglas Corporation

Research Objective

To develop computational codes capable of predicting the

electromagnetic signature of an aerospace vehicle. Such codes

would be of great value in the design and development of fighter

aircraft and missiles because a greater number of candidate configurations could be considered and their performance evaluated,

thus reducing the cost and time required in the design cycle.

Approach

The approach taken is to solve the three-dimensional (3-D)

Maxwell's equations. A central-differencing, finite-volume,

time domain/frequency domain, multizone method is used to

compute the electromagnetic (EM) field about scattering bodies.

A hybrid approach that combines a method of moments code

(CARLOS-3D) with the computational fluid dynamics-based EM

code is also developed to enhance the EM modeling capability.

Accomplishment Description

The EM code (CFDMAXES) was extended to calculate the scattering from geometrically and electrically complex 3-D bodies, in

both the time and the frequency domains. Accomplishments

include the calculation of the radar cross section of aircraft inlets,

nozzles and radomes, and radar range pylons. Modifications

to the CFDMAXES code enabled efficient coupling with the

CARLOS-31D code for the frequency domain scattering

calculation.

r

s

Color contour plots of the magnetic current magnitude on the surface of the one meter NASA almond at 2 GHz. Plots are for axial

incidence with vertical and horizontal polarization. Red = high

magnitude and blue = low magnitude.

Significance

The project represents the development of a leading-edge

technology in computational electromagnetics and it will have

substantial impact on the development of both fighter aircraft and

missile programs.

Future Plans

Planned extensions include implementing convergence enhancing techniques for the frequency domain formulation and applying broadband calculations for the time domain formulation.

Keywords

Computational fluid dynamics, Electromagnetic signature

Color contour plot of the magnetic current magnitude caused by

an axially incident electromagnetic field of frequency 1 GHz

within an untreated aircraft exhaust system. Red = high magnitude

and blue = low magnitude.

102 General/E lectromagnetics
(a} (b)

(c) (d) L^

Computational Fluid Dynamics Approach to Computational Electromagnetics

Vijaya Shankar, Principal Investigator

Co-investigators: William Hall, Chris Rowell, and Alireza Mohammadian

Rockwell International Science Center

Research Objective

To extend algorithmic advances in computational fluid dynamics

to the time-domain Maxwell's equations to study electromagnetic

problems of interest for defense and commercial applications.

Approach

The time-domain Maxwell's equations were written in differential

conservation form (for structured grid arrangement) or in integral

conservation form (for unstructured grid arrangement). An appropriate upwind, finite-volume time/space discretization was

employed using Riemann solvers for interface flux evaluation to

update the electric and magnetic fields at each time step. The

electromagnetic incident wave can be either a continuous wave

(single frequency) or a pulse (broadband frequency).

Accomplishment Description

Both the structured grid (RCS3D) and the unstructured grid

(RCSUM versions of the computational electromagnetics (CEM)

code run efficiently on the Cray C-90. Depending on the problem

size, the RCS3D code runs at 500-700 MFLOPS per processor

and the RCSUN code runs at 200+ MFLOPS. An MPP version

of the RCS3D code is also running on the Intel Paragon. The

structured/unstructured CEM capability was demonstrated on a

variety of electromagnetic problems including the prediction of

the radar cross section (RCS) of a complete fighter requiring

10 million grid points.

Significance

This CEM development represents the state of the art in largescale simulation and is a critical technology for design of low

observable platforms.

Future Plans

The CEM capability for RCS problems involving 100+ million

cells will be demonstrated using the Intel Paragon or the

Cray T3 D.

Keywords

Radar cross section, Maxwell's equations

CEM applications for defense and commercial use, (a) microstrip patch antenna, (b) low observables, (c) monopole antenna, and

(d) bioelectromagnetics.

Genera l/Electromagnetics 103
Electromagnetic Scattering by Airborne Structures

John L. Volakis, Principal Investigator

Co-investigator: Arindam Chatterjee

University of Michigan

Research Objective

To develop numerical methods and associated computer codes

for the analysis of electromagnetic scattering by airborne composite targets. Typically, a target is illuminated by a continuous

wave radar signal; however, this study focuses on the electromagnetic energy reflected from the target.

Approach

To model arbitrary composite targets and reduce memory

requirements, the finite element method (FEM) was used for electromagnetic field simulations. Even though FEM is used extensively in fluid mechanics and electrostatics, its implementation in

electromagnetics is fundamentally different and is complicated by

the need to model unbounded domain problems. Thus it has only

recently been considered for high-frequency electromagnetic

scattering computations.

Accomplishment Description

A three-dimensional (3-D) finite element code (FEMATS) for

electromagnetic scattering was developed. The code employs

edge-based elements for better simulation of electromagnetic

field characteristics and new absorbing boundary conditions

(ABCs) for more efficient truncation of the finite element mesh.

The mesh truncation is based on vector absorbing boundary conditions and can be applied on a boundary conformal to the scatterer or antenna system. More than 10 Electromagnetic Code

Consortium benchmark geometries were modeled and analyzed,

including metallic and dielectric plates, cavities, jet engine inlets,

and nonhomogeneous dielectrics. The largest problem simulated

scattering by a foam cylinder with collinear and off-axis wires.

The problem needed only 12.5 Cray C-90 minutes for each

observation angle to converge to the desired solution.

Significance

The development and validation of conformal ABCs leads to substantial reduction in the degrees of freedom required to model

targets. An 11 cubic wavelength dielectric scatterer was modeled

and larger problems can be modeled.

Future Plans

Further improvements on the ABCs and mesh refinement strategies will lead to a substantial reduction in memory and processing time. Numerical ABCs with iterative refinement and adaptive

finite elements for simulating large, complex 3-D structures with

minimal computational resources will be investigated.

Keywords

Three-dimensional finite elements, Absorbing boundary

conditions, Parallelization

—10

N —15

• •

—25 •

m

co —30 Measured U

Y ----- Body of revolution

v —35 c^ • FE-ABC (VV pol)

m

—40

0 30 60 90

Observation angle (deg)

Backscatter radar cross section of a metallic cylindrical inlet

(1.25.1 diameter,- 1.875 A height).

Mixed mesh termination scheme employing flat, cylindrical, and

spherical sections.

104 General/Electromagnetics
Computational Radar Cross Sections

Alex C. Woo, Principal Investigator

Co-investigators: Michael Simon and Michael Schuh

NASA Ames Research Center/Sterling Software Systems

Research Objective

To enhance the radar cross section (RCS) prediction of bodies

with metallic and treated surfaces.

Approach

A finite-volume time-domain program, RCS3D, was used to

propagate electromagnetic waves through a variety of structured,

body-fitted grids. Both time-harmonic and broadband-incident

fields were tested. Computed and experimentally measured RCS

data were compared to assess the effects of various developments

on the accuracy of the solution.

Accomplishment Description

A method was developed for terminating broadband (pulse

mode) calculations based on the energy of the electromagnetic

field. The energy is computed at successive instants of time. The

calculation ends when the energy has fallen to a specified fraction of the peak energy. A calculation for a square aperture cavity

required 18 megawords of memory and 7 CPU hours on the

Cray C-90. When using the pulse termination technique, the calculation had no effect on the required memory, but did reduce

the CPU time to less than 6 hours. The use of energy dissipation

as an automatic termination criterion is a valuable enhancement

to the code's usability. A second enhancement was to develop a

technique for generating bistatic inverse synthetic aperture radar

(ISAR) images. This technique uses the radar return for a range of

frequencies and bistatic look angles to determine where the areas

of intense scattering are located. The figure shows ISAR images

for a trapezoidal plate with the incident field coming from the

right side. The image for horizontal polarization shows that the

near edge has high return (dark grey), but for vertical polarization

the far edge has high return. The returns appearing before the

plate are wraparound ghost images from the ISAR algorithm.

Significance

A physical basis was implemented to automatically determine the

number of time steps required for a pulse mode calculation to

yield accurate RCS predictions. This criterion eliminates a degree

of uncertainty in the results and reduces the computational time.

Also, a tool is now available to produce bistatic ISAR images of

radar targets and to diagnose geometry and grid problems.

Future Plans

RCS31D will be validated on a radar test fixture with a variety of

detail attachments.

Keywords

Finite volume, Time domain, Bistatic inverse synthetic aperture

radar

4

Bistatic ISAR images from a single pulse calculation of RCS3D.

The incoming wave strikes the target (dashed lines) from the right.

Polarization is seen when the electric field is horizontal (top) and

vertical (bottom).

General/Electromagnetics 105
Arbitrary Lagrangian—Eulerian Three-Dimensional Simulations

Joseph D. Baum, Principal Investigator

Co-investigator: Rainald Lohner

Science Applications International Corporation/George Mason University

Research Objective

To investigate shock diffraction about moving and stationary

complex-geometry, three-dimensional (3-D) bodies.

Approach

A new 3-D arbitrary Lagrangian-Eulerian (ALE) hydro-solver

based on the finite-element method flux-corrected transport

(FEM-FCT) concept, FEFLO96, was used. The preferred approach

for grid adaptation is H-refinement. The high-order scheme used

was the consistent-mass Taylor—Galerkin algorithm. Combined

with a modified second-order Lapidus artificial viscosity scheme,

the resulting scheme is second-order accurate in space and

fourth-order accurate in phase. The spatio-temporal adaptation

was based on local H-refinement, where the refinement/deletion

criterion is a modified H2-seminorm, and density is used as the

critical parameter. Over the past year, the element-based data

structure was converted to an edge-based structure, thereby significantly reducing indirect addressing, floating point operations,

and memory overhead.

Accomplishment Description

The new 3-D, adaptive, finite-element, edge-based methodology

was applied to two simulations. The first simulation was a blast

diffraction within the passenger compartment of a civilian widebody jetliner (a Boeing 747). This study simultaneously evaluated

the ability of a wide-body jetliner to survive a terrorist attack and

the feasibility of applying an adaptive grid methodology to the

simulation of shock-wave diffraction about hundreds of complexgeometry structures. A large number of shock waves were produced because of shock diffraction about the seats, bulkheads,

and luggage compartments. Excellent shock adaptation and resolution were obtained. The second simulation modeled pilot/seat

ejection from an F-16 fighter at supersonic speed. The results

demonstrate the ability of the numerical methodology to aCCUrately predict both the steady flow around the plane and within

the cockpit after canopy ejection and the transient forces acting

on the pilot during ejection. The simulations identify a mechanism that may be responsible for the neck injury experienced by

many ejecting pilots. The results also demonstrate that, although

the ejection velocity is relatively low, the computational results

are significantly different from those obtained for a captive solution under identical conditions.

Significance

The results demonstrate the feasibility of applying 3-D, adaptive,

finite-element, edge-based schemes on unstructured grids to simulate shock-wave diffraction about multiple complex-geometry

moving or stationary structures.

Future Plans

The range of applications will be extended to model the structural

response to shock loading by incorporating state-of-the-art computational structural dynamics in the methodology.

Keywords

Shock-wave diffraction, Arbitrary Lagrangian—Eulerian methodology, Shock-capturing schemes, Unstructured grids

Blast diffraction in a Boeing 747. Internal pressure contours and

external Mach number contours at time = 14 milliseconds.

Pressure contours about the F-16 and the ejecting pilot at

initiation of ejection (time = Q milliseconds) and during ejection

(time = 27 milliseconds).

106 General/Fluid Mechanics
Unsteady Viscous Marine Propulsor Hydrodynamics

Fred Stern, Principal Investigator

Co-investigators: E. Paterson and B. Chen

University of Iowa

Research Objective

To develop computational fluid dynamics (CFD) simulation

methods applicable to naval hydrodynamics (marine propulsory),

which are unlike related applications because marine propulsors

are embedded in the thick stern boundary layer and wake and are

susceptible to cavitation.

Approach

The research is based on further developments and extensions of

current Reynolds-averaged Navier–Stokes (RANS) methods and a

new CFD capability for highly efficient unsteady RANS calculations, which also can be extended for large-eddy simulations

(LES) or direct numerical simulations (DNS). The current method

uses finite-analytic spatial discretization, pressure-implicit splitoperator, marker and cell (MAC), and semi-implicit pressurelinked equation pressure-velocity coupling algorithms, Chimera

domain decomposition, and a hierarchy of turbulence models

including Baldwin–Lomax, k--F- with either wall functions or a

one-equation model for the near-wall flow, and non-isotropic

models such as the nonlinear k--F and algebraic Reynolds-stress

models. The new method uses high-order finite differences,

fractional-step/MAC pressure-velocity coupling, fourth-order

artificial dissipation, and multigrid acceleration of the pressure

equation.

Accomplishment Description

Steady RANS solutions and validation were accomplished for

marine propulsors P411 9 and P4842 for design and off-design

conditions (1 Cray C-90 hour and 24 megawords of memory).

Zonal methods and advanced turbulence modeling for foils and

marine propulsor P4119 were completed. Unsteady-flow calculations and validation studies were done for comparison to the

Massachusetts Institute of Technology flapping-foil experiment

(3 Cray C-90 hours and 6 megawords of memory). Investigations

into interactions between natural and forced unsteady flows for

fixed and moving boundaries using Chimera domain decomposition were completed (20 Cray C-90 hours and 8 megawords of

memory). Finally, unsteady RANS solutions were extended for

marine propulsory P4119 and P4132 with temporally and

spatially varying inflows (28 Cray C-90 hours and 32 megawords

of memory).

Significance

This research supports the Office of Naval Research basic and

applied research programs focusing on the fundamental importance in the fluid mechanics of moving boundary problems and

interactions between natural and forced unsteady separated

flows. The research is also of strategic importance to the design of

marine propulsor–body configurations—especially signature

identification and reduction.

Future Plans

Development will continue on numerical methods with

application to unsteady propulsor-blade flow and complete

configurations. Also, traveling-wave external flows over simplified blade geometries will be studied using LES and DNS to

support the development of appropriate turbulence models for

unsteady propeller-blade boundary layers and wakes.

Keywords

Chimera, Propeller

Flapping-foil experiment overlaid grid system (top) and perturbation pressure contours (bottom) with instantaneous perturbation

velocity particle traces.

PAean 1 4 214T 34T 4r4T

3

-1 1

IF i

Surface pressure contours at different time periods (T) for marine

propulsor P4732 with sixth-harmonic screen-wake inflow for the

suction side (top) and pressure side (bottom).

General/Fluid Mechanics 107
Free-Surface Effects on Boundary Layers and Wakes

Fred Stern, Principal Investigator

Co-investigators: Y. Tahara and J. Choi

University of Iowa

Research Objective

To develop computational methods for the solution of the

unsteady, three-dimensional Navier—Stokes equations for surfacepiercing bodies with free-surface gravity waves.

Approach

The approach is based upon further developments and extensions

of current Reynolds-averaged Navier—Stokes (RANS) methods and

a new computational fluid dynamics capability for highly efficient unsteady calculations that can be extended for large-eddy

(LES) or direct-numerical (DNS) simulations. The current method

uses finite-analytic spatial discretization, marker and cell and

semi-implicit pressure-linked equation pressure-velocity coupling

algorithms, and a hierarchy of both isotropic and non-isotropic

turbulence models. The new method uses high-order finite differences, fractional-step/MAC pressure-velocity coupling, fourthorder artificial dissipation, and multigrid acceleration of the pressure equation. Both small-domain viscous-inviscid interactive

and viscous large-domain approaches are under development

with special emphasis on treatment of the free-surface boundary

conditions, turbulence modeling, and extensions for unsteady

flow. The computational studies are guided and supported by a

concurrent program of towing-tank experiments.

Accomplishment Description

A detailed study of the Navier—Stokes-wave/flat-plate flow geometry was completed (0.5 Cray C-90 hours and 10 megawords of

memory). Also completed was an analysis of wake bias through

Navier—Stokes, boundary-layer, and perturbation expansion

equation solutions for a flat-plate boundary layer and wake in

temporal, spatial, and traveling horizontal-wave external flows

(0.25 Cray C-90 hours and 6 megawords of memory). The interaction approach for the Series 60 C B = 0.6 ship model was validated (0.5 Cray C-90 hours and 10 megawords of memory).

Development of the large-domain approach was continued

(1 Cray C-90 hour and 10 megawords of memory), and development of LES/DNS simulation methods for the Navier—Stokes

wave/flat-plate flow geometry was initiated (25 Cray C-90 hours

and 50 megawords of memory).

Significance

This research is sponsored by the Office of Naval Research in

support of its basic research program focusing on 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. This research is of strategic

importance for the identification and reduction of surface

wake signatures.

Future Plans

Development of the numerical method with application to

unsteady free-surface flows and free-surface/propulsor interaction

will continue. Investigations concerning the physics of turbulence

and turbulence modeling for the solid—fluid juncture boundary

layer and wake, and studies of wave-induced separation for ships

at yaw will be completed.

Keyword

Ship hydrodynamics

0.094

— 0.081

0.068

0.055

0.042

0.030

0.017

r

0.004

z y+^`

^'^^ -0.009 j

-0.022

-0.034

-0.047

-0.066

Large-domain solution for a Series 60 CB = 0.6 ship model and Froude number = 0.316. The wave pattern (top),

pressure contours (bottom), and free-surface perturbation vectors (top and bottom) are shown.

108 General/Fluid Mechanics
Impeller Computational Fluid Dynamics Modeling

Matthew E. Thomas, Principal Investigator

Co-investigators: Mahesh M. Athavale and Mark L. Ratcliff

CFD Research Corporation

X

Research Objective

To develop, demonstrate, and validate an efficient and accurate

computational fluid dynamics (CFD) analysis procedure for the

complex turbulent flows in liquid rocket propulsion impellers.

Approach

CFD modeling of liquid rocket propulsion components requires

a combination of advanced geometric modeling procedures,

boundary condition implementation, robust CFD solution techniques, and highly interactive postprocessing capabilities. Two

geometric modeling packages, PATRAN and ICEM-CFD, were

examined. The CFD results were validated with benchmark quality laser Doppler velocimeter (LDV) impeller data.

Accomplishment Description

Geometric data used in this effort included initial graphics

exchange standard files associated with two distinct geometry

classes (a Rocketdyne test inducer and the Space Shuttle main

engine (SSME) high-pressure fuel turbopump (HPFTP) impeller).

Each package was examined in detail for user efficiency and

geometric modeling accuracy. The open-ended architecture of

PATRAN has the potential for mature interaction between CFD

and structural/thermal analysis packages. The structured grid

modeling procedures implemented in ICEM-CFD were suitable

for complex impeller geometries. Flow computations were done

using REFLEQS-3D, an advanced finite-volume Navier—Stokes

code with capabilities that include high-order differencing

schemes, different turbulence models, a range of boundary condition types, and treatment of rotating coordinate frames. Flow

calculations for the inducer were done on coarse and fine grids

(41,000 and 100,000 cells) built with PATRAN. The SSME HPFTP

impeller grid, with partial and full blades, was generated using

ICEM-CFD. Two grids (56,000 and 130,000 cells) were used on

the impeller solutions because they were more sensitive to the

grid size. The figures show the surface pressure plots for the two

components. The computed velocities and flow angles were

compared with LDV data. Detailed comparisons on the impeller

and limited comparisons on the inducer data showed very good

agreement with experimental measurements. This study demonstrates that CFD technology has sufficient efficiency and accurac}

for implementation into the impeller design process.

Significance

An efficient and accurate CFD analysis procedure that is

applicable within the preliminary and detailed design procedures

of liquid rocket propulsion turbomachinery components was

demonstrated.

Future Plans

The CFD analysis methodology is being extended to include

computations of cavitating flows. Attributes of PATRAN,

ICEM-CFD, and REFLEQS-3D identified during this effort are

being implemented to link the computer assisted design, structural, thermal, hydrodynamic, and results analysis (analytical and

test) disciplines. The package will be extended to other rotating

machines including turbines, fans, and propellers.

Keywords

Impeller, Turbomachinery, Computational fluid dynamics

Cp ,d

0.515

v

y`

- t

I ^ £ 

t ^r f

0.0

G

Pressure coefficient distribution on a Rocketdyne test inducer.

SSME high-pressure fuel turbopump impeller pressure loading.

General/Fluid Mechanics 109
Modeling of Pulmonary Fluid Dynamics

Jeffrey R. Hammersley, Principal Investigator

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

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

Research Objective

To study the complex fluid interactions within inaccessible small

respiratory airways and vessels. To determine the effects that

normal anatomic variations cause within airways and vessels,

the complexity of flow within confined bifurcating tubes will

be examined.

Approach

Multiblocked two-dimensional (2-D) and three-dimensional (3-D)

structured grids were constructed to mimic the anatomic features

of small airway and vascular bifurcation using the EAGLE grid

code. These computational bifurcations were studied over

a biologically realistic range of Reynolds numbers (50-1,400)

through the use of a Taylor—Whitfield ABRAINS flow solver modified to accommodate flow reversal. Evaluation of the computed

flow results was performed using FAST, PLOT3D, and secondary

programs for boundary layer and shear force calculations.

Accomplishment Description

Three-dimensional flow solutions in single bifurcation have been

produced along with multibranched symmetrical and asymmetrical 2-D solutions. These provide velocity, pressure, wall shear,

and particle trajectory information for a wide range of realistic

biologic flows. Work has been extended from steady flow to

the pulsatile and oscillatory flow within the 2-D geometric grids;

3-D multibranched grids and dynamic solutions are still

under development.

Significance

These data allow medical scientists to predict delivery of

aerosolized medications, quantify particulate deposition during

environmental exposures, and link changes in airway geometry to

the disability seen in asthma and emphysema. This work continues a collaborative effort to provide realistic biologic flow simula

tions of fluidic mixing, particle transport, and shear forces on

airway and vascular walls. It will be extended to pulsatile and

oscillatory conditions. When completed in 3-D, flow simulation

through biologic structures should produce a better understanding of the physical forces that limit transport in airways and

blood vessels.

Future Plans

The investigation of symmetrical and asymmetrical bifurcating

flows will be extended to dynamic flows in serially branching

3-D cascades. These efforts set the stage for future incorporation

of wall elasticity and particle-flow interactions.

Keywords

Lung, Incompressible flow, Oscillatory, Pulsatile

Inspiratory flow in a 2-D asymmetric network under oscillatory

conditions.

Expiratory phase flow in a 2-D asymmetric network.

110 General/Life Science
Structure and Functions of Membranes

Andrew Pohorille, Principal Investigator

Co-investigator: Michael A. Wilson

University of California, San Francisco

Research Objective

To explain how interfaces between protocellular membranes and

water could have promoted catalysis of the chemical reactions

needed for cellular functions in the absence of the complex

enzymes of contemporary cells.

Approach

The water—membrane system consists of a solute molecule and

72 membrane-forming molecules arranged in a bilayer between

two water lamellae of 1,200 molecules each. The behavior of this

system is simulated using the molecular dynamics method. The

results are the statistical and dynamic properties of the system

(which can be directly compared with experimental results) and a

detailed microscopic description of the system.

Accomplishment Description

Alanine dipeptide, a simple model for peptides, was studied at

the water—membrane interface and in other typical protobiological environments water, oil, and at the water—oil interface. It

was shown that the dipeptide's ability to form ordered structures

(conformations) strongly depends on the environment. It was further demonstrated that this molecule tends to accumulate at the

water—membrane and water—oil interfaces. By binding the substrate and lowering the activation energy the interface acts as a

catalyst to the isomerization reaction. The time required for a

single calculation varies from 60 Cray C-90 hours for the

water—membrane system to 5 hours for bulk water and oil

solvents. Central memory requirements vary from 10 to

50 megawords, depending on the data analysis performed.

Significance

The ability of simple peptides to form ordered structures and

adopt preferred orientations is unique to interfacial systems. Thus,

membrane surfaces of protocells may have provided an environment for primitive catalysis even before the development of the

complex enzymes of contemporary cells.

Future Plans

Studies of the transfer of small molecules across membranes and

the development of a general, molecular model of the transport of

nutrients and waste products through walls of protocells will continue. Also, proton transport across protocellular walls as an

essential mechanism of primitive bioenergetics will be studied.

Keywords

Origin of life, Protocells, Protein—membrane interactions

Alanine dipeptide at the interface between membrane and water. The oxygen and hydrogen

atoms of water are red and white, respectively; the lipid tail-group atoms are blue and the headgroup atoms are pink. In the alanine dipeptide molecule, the carbon atoms are green, the nitrogen atoms are yellow, the oxygen atoms are magenta, and the hydrogen atoms are gray.

General/Life Science 1 1 1
Finite Volume Neuronal Modeling

Muriel D. Ross, Principal Investigator

Co-investigators: David G. Doshay, Samuel W. Linton, and Timothy). Barth

NASA Ames Research Center/Sterling Software Systems

Research Objective

To investigate whether application of modern computational fluid

dynamics (CFD) tools can improve the accuracy of simulations

used to determine the influence of ultrastructural detail on neural

functioning. The one-dimensional (1-D) compartmental models

used in computational neuroscience were expanded to a model

based on finite volume analysis to determine whether this augmentation would produce more biologically relevant computer

models when communications between complex neurons

were simulated.

Approach

The morphology of the neurons was obtained from computer

assisted three-dimensional (3-D) reconstructions of neurons from

a vestibular gravity sensor (macula) sectioned serially and photographed in a transmission electron microscope. An unstructured

mesh describes neuronal surfaces in the reconstructions. The

neural structure used in the test is the calyx, a complex neuronal

ending that has inner and outer membranes separated by cytoplasm as well as twig-like extensions (processes) that often have

the geometric shape of frustums. The finite volume solver uses the

generalized minimum residual method to solve for voltage

changes in frustums (quasi-] -D)and triangular prisms (quasi2-D) of the calyx mesh. The connectivity scheme allows for an

unlimited number of quasi-] -Dprocesses and for their arbitrary

branching.

Accomplishment Description

Results show excellent agreement with available analytic

solutions (see figures). The effect produced on a calyx by activation of a synapse on a process was measured over a range of cell

membrane and cytoplasm resistances and synapse conductances.

Each parameter sweep required 50 simulations over a single

calyx, took about 3 Cray hours, and used less than 1 megaword

of memory.

Significance

The excellent agreement with analytic solutions obtained by

the solver gives credibility to the predictions it makes for more

complex geometries. The parameter sweeps demonstrate the

potential of the solver to provide information quickly when biologically realistic geometries are simulated. The speed of the Cray

and minimal memory requirements of the solver prove the

feasibility of accurate and biologically relevant simulations of

neuron networks.

Future Plans

Further simulations are required to quantify the effect of morphology on electrophysiological behavior. Complex simulations

involving interconnected calyces on single neurons and networks

of neurons are the next step. These simulations will begin with

quasi-2-D meshes, but will advance to fully 3-D meshes.

Keywords

Computational fluid dynamics, Neurons

10-3

10-210-1100101 102

Radius

There is excellent agreement between analytic (solid line) and

computed (dots) solutions for a circular flat plate with a constant

voltage at the origin.

Voltage distribution 0.025 msec after application of 20 mV to the

base of an abstract neural structure that has an analytic solution.

10

8

6

m

4

0

2

0

—2

10-4

112 General/Life Science
Solar Wind Interaction with Mars

Stephen H. Brecht, Principal Investigator

Co-investigator: John R. Ferrante

Berkeley Research Associates

r

Research Objective

To study the kinetic interaction of the solar wind plasma with

Mars and Venus using state-of-the-art three-dimensional (3-D)

plasma kinetic simulations. Of principal interest is the process of

shock formation where the effective viscous length is a significant

fraction of the planet radii. In addition, the kinetic processes that

lead to the plasma distribution functions measured by various

orbiting spacecraft are studied.

Approach

The study of the solar wind interaction with Mars and Venus is

performed with a 3-D hybrid particle code. A hybrid particle

code treats the ions as individual particles (kinetic) and the electrons as a neutralizing massless fluid. The code can treat an

unlimited number of ion species. The simulations study the effect

of mass introduction via charge exchange and photoionization of

the planets' atmospheres.

Accomplishment Description

The simulations show that the mass loading of the solar wind with

ionospheric oxygen affect the flow characteristics around Mars.

Under many circumstances the planet does not possess a "traditional" bow shock. In addition, the interaction between the solar

wind and the planet changes in a nonhydrodynamic way. Strong

solar wind pressure variations produce almost no effect on the

subsolar shock location, but make significant differences in the

flare angle and tail. The simulations were extended to include the

tail region of the planet. The results were compared to the data

returned to Earth from the Russian Phobos-2 spacecraft. Data

from both the elliptic and circular orbits (see the figure) were

tested against the simulation results by "flying" the Russian spacecraft through the simulation data. The consistency of our results

with the actual data provides evidence that the solar wind interaction with Mars does not produce a classic shock, although the

Alfven Mach number is approximately 13. Comparison with the

data taken in the tail of the planet demonstrates that Mars need

not have an intrinsic magnetic field to explain the data taken

by Phobos-2.

Significance

These results are gradually changing the view held by many

scientists concerning the solar wind interaction with Mars. They

clearly demonstrate that the interaction cannot be viewed in a

magnetohydrodynamic or gas-dynamic sense without missing

significant features of the interaction. The data as reported and

interpreted by some scientists do not provide evidence of the

presence of an intrinsic magnetic field within Mars.

Future Plans

The conditions under which Mars may have a shock structure

protecting it from the solar wind will be investigated. The role of

the planetary ionospheres will be investigated with the timedependent chemistry models developed for the HALFSHEL code.

Keywords

Shock formation, Mars, Solar wind

Publications

1. Brecht, S. H.; and Ferrante, J. R.: Examination of Mass Loading

Effects on the Martian Magnetosphere. EOS, Trans. Am.

Geophys. Union Fall Meeting, 1993, vol. 74, p. 538.

2. Brecht, S. H.; and Ferrante, J. R.: Hybrid Simulation of the

Martian Boundary Layer. EOS, Trans. Am. Geophys. Union

Spring Meeting, 1994, vol. 75, p. 284.

The magnetic field structure developing around Mars and the

actual Phobos-2 orbits through the system.

Genera l/Magnetospheric Physics 113
Burner-Stabilized Flames in Microgravity

K. Kailasanath, Principal Investigator

Co-investigator: G. Patnaik

Naval Research Laboratory/Berkeley Research Associates

Research Objective

To develop a better understanding of the effects of gravity and

heat loss on the structure and stability of premixed flames.

Approach

A detailed time-dependent, multidimensional, multispecies,

numerical model is used as a tool to simulate flames stabilized on

a burner on Earth and in zero gravity. The model includes finiterate, detailed chemical kinetics coupled to algorithms for convection, thermal conduction, viscosity, molecular diffusion, and

external forces such as gravity.

Accomplishment Description

Cellular structures are observed for the hydrogen—air flames in

both zero and Earth gravity. As the inflow velocity is reduced, the

flames move closer to the burner and the heat loss to the burner

increases. At low inflow velocities, the cellular structures are suppressed in both cases, suggesting that increased heat loss to the

burner is the primary mechanism stabilizing these flames. As

shown in the figure, Earth-gravity flames are smoother than the

zero-gravity flames, indicating that gravity still plays a small part

in further stabilizing the flames. A typical calculation takes

25 Cray C-90 hours.

Significance

These simulations clarify the relative importance of gravity and

heat loss to the burner on the structure and stability of lean

hydrogen—air flames.

Future Plans

The code will now be used to simulate and understand a number

of other interesting flame phenomena observed on Earth and

in microgravity.

Keywords

Premixed flames, Buoyancy, Heat losses

The effect of zero gravity (top) and Earth gravity (bottom) on a hydrogen—air flame stabilized on an isothermal burner at 10 cm/sec inflow

velocity (left) and 7 cm/sec inflow velocity (right).

114 General/Reactive Flow
The schematic of the disk-shaped protoplanetary nebula is composed of a gas—

particle mix with local vertical thickness

about one-tenth the local radius. The

computational grid is envisaged as a tiny

representative volume in the overall nebula. The boxes represent the computational grid in a sequence of time steps

showing the evolution of several different

sets of particles, which come together as

clumps at different times.

Protoplanetary Nebula Particle-Gas Dynamics

Jeffrey N. Cuzzi, Principal Investigator

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

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

Research Objective

To study conditions prevailing in the nebula from which the solar

system formed at a time when the growing particles were small

enough to have important dynamic interactions with the predominant gas, and to relate these studies to evidence seen in primitive

meteorites.

Approach

Direct numerical simulations of three-dimensional turbulence

were used to trace the evolution of large numbers of test particles

of varying stopping time in the gas. Particles of stopping time

comparable to the turn-over time of the smallest eddy are preferentially concentrated into clumps that are at least an order of

magnitude more dense than average clumps. Turbulence particle

damping may also be incorporated.

Accomplishment Description

Calculations were extended to higher Taylor microscale Reynolds

numbers (80 completed, 150 running) and several aspects of the

resulting particle density distribution were studied. Several physically based relationships were developed to allow scaling the

results to the much larger Reynolds numbers of the nebula. The

scaling relations are being validated with the calculations now

running. The fluid regimes that are depleted or populated by partitles and the relative velocities of the particles to the gas and to

each other are described.

Significance

The study implies that dense concentrations of particles appear

on spatial scales comparable to the Kolmogorov scale, and that

particles of millimeter and submillimeter diameter will be preferentially concentrated by a factor between 102 and 105 in nebula

turbulence. These sizes are in excellent agreement with the sizes

of chondrules and other meteorite fragments that are dominant

components of primitive meteorites. Previously, there was no

explanation for the makeup or production of chondritic meteorites; this concentration process may be the cause of the sorting,

and thus a critical step in the accumulation of the earliest cometand asteroid-sized planetesimals.

Future Plans

More accurate calibration of the concentration dependence and

preferred size on Reynolds number must be obtained. The effect

of particle mass loading on turbulence damping at high concentrations will be studied.

Keywords

Planet formation, Turbulence, Two-phase flow

General/Space Science 115
Page intentionally left blank 
Index
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Index by Research Sites

Page

Adaptive Research Corporation

Robert R. Chamberlain Jet Interaction Effects on Aerodynamic Control of Missiles

Allison Engine Company

Edward J. Hall Advanced Ducted Propfan Analysis Code Certification ...

Edward J. Hall Fan Rotor/Endwall Treatment Interaction .........................

Edward J. Hall Transonic Core Compressor Rotor/Stator Interaction .......

Applied Research Associates, Inc.

Robert H. Sues Parallel Multidisciplinary Stochastic Optimization ..........

AS&M, Inc.

Michael C. Fischer Supersonic Laminar Flow Control Experiment .................

(with NASA Lewis Research Center)

Berkeley Research Associates

25

.............. 53

.............. 54

.............. 52

... 63

32

Stephen H. Brecht Solar Wind Interaction with Mars ................................................................................ 113

K. Kailasanath Burner-Stabilized Flames in Microgravity .................................................................... 114

(with Naval Research Laboratory)

Boeing Commercial Airplane Group

David E. Ashpis Transition in a Highly Disturbed Environment ............................................................. 75

(with NASA Lewis Research Center)

N. J. Yu Evaluation of Navier—Stokes Codes for Transport Configuration Analysis ..................... 31

Boeing Company, The

Forrester T. Johnson Multipoint Aerodynamic Design ................................................................................. 10

Brown University

George E. Karniadakis Direct and Large-Eddy Simulation of Wake Flows ....................................................... 78

George F. Karniadakis Simulation of Complex-Geometry WalJ-Bounded Flows ............................................. 79

Cambridge Hydrodynamics, Inc.

Steven A. Orszag Unsteady Turbulence Model Simulations .................................................................... 86

119
Page

Carnegie Mellon University

Wei J. Chyu Airframe/Inlet Aerodynamics ...................................................................................... 47

(with NASA Ames Research Center)

Toni I-P. Shih Flow in Turbine-Blade Cooling Passages ..................................................................... 60

(with NASA Lewis Research Center)

CFD Research Corporation

Clifford E. Smith Jets-in-Cross-Flow Mixing ........................................................................................... 62

(with NASA Lewis Research Center)

Matthew E. Thomas Impeller Computational Fluid Dynamics Modeling .................................................... 109

European Space Agency

John Rakiewicz Titan Entry Flow Field of the Cassini—Huygens Probe ................... 27

(with Jet Propulsion Laboratory and McDonnell Douglas Aerospace)

General Motors Corporation, Allison Gas Turbine Division

Kurt F. Weber Chimera Domain Decomposition Applied to Turbomachinery Flow ........................... 67

George Mason University

Joseph D. Baum Arbitrary Lagrangian—Eulerian Three-Dimensional Simulations ................................... 106

(with Science Applications International Corporation)

George Washington University

Jassim A. AI-Saadi Transonic Flows Around Complex Geometries ........................................................... 35

(with NASA Langley Research Center)

Grumman Corporate Research Center

Frank Marconi Comparison ofTLNS3D Computations with Test Data ................................................ 13

High Technology Corporation

Bart A. Singer Origins of Near-Wall Turbulence ................................................................................ 90

(with NASA Langley Research Center)

120
Page

ICASE

Dimitri J. Mavriplis Navier—Stokes Computations on Unstructured Grids ................................................... 14

R. C. Swanson Simulation of Scramjet Flow Fields ............................................................................. 64

(with NASA Langley Research Center)

ICOMP

Chunill Hah Unsteady Three-Dimensional Flow in a Transonic Compressor ................................... 51

(with NASA Lewis Research Center and Wright Patterson Air Force Base)

Shaye Yungster National Aero-Space Plane Nozzles with External Burning ......................................... 68

(with NASA Lewis Research Center)

Iowa State University

Richard H. Pletcher Three-Dimensional Liquid Sloshing Flow Simulation .................................................. 87

Jet Propulsion Laboratory

Christian L. Keppenne Intraseasonal Atmospheric Oscillations ....................................................................... 97

John Rakiewicz Titan Entry Flow Field of the Cassini—Huygens Probe .................................................. 27

(with McDonnell Douglas Aerospace and European Space Agency)

Lockheed Aeronautical Systems Company

Pradeep Raj CFL3D Code Evaluation .............................................................................................. 15

Lockheed Engineering and Sciences Company

Kenneth M. Jones Subsonic High-Lift Analysis ........................................................................................ 29

(with NASA Langley Research Center and ViGYAN, Inc.)

Steve L. Karman, Jr. Unstructured Grid/Flow Solver Calibration ................................................................. 33

Christopher L. Reed High-Angle-of-Attack Aerodynamics ........................................................................... 41

David M. Schuster Analysis of High-Speed Civil Transport Transient Response ......................................... 17

MCAT Institute

Guru P. Guruswamy Aeroelasticity of Wing—Body—Control Configurations .................................................. 8

(with NASA Ames Research Center)

Guru P. Guruswamy Wing—Body Aeroelasticity ........................................................................................... 9

(with NASA Ames Research Center)

Nateri K. Madavan Boundary-Layer Transition on a Heated Flat Plate ....................................................... 82

Lewis B. Schiff Investigation of Forebody Flow Control ....................................................................... 30

(with NASA Ames Research Center)

121
Page

E. Tu Laminar Flow Supersonic Wind Tunnel ...................................................................... 34

(with NASA Ames Research Center)

McDonnell Douglas Aerospace

John Rakiewicz Titan Entry Flow Field of the Cassini—Huygens Probe .................................................. 27

(with Jet Propulsion Laboratory and European Space Agency)

Arvin Shmilovich Analysis of Nacelle/Pylon/Wing Installations .............................................................. 61

Arvin Shmilovich Computational Studies of Advanced Wings at High Incidence .................................... 42

Scott Ward Reusable Launch Vehicle Flow Simulations ................................................................ 66

McDonnell Douglas Corporation

Ramesh K. Agarwal Computational Fluid Dynamics Solution for Maxwell's Equations .............................. 102

Ramesh K. Agarwal Unstructured Grid Finite-Element Solutions ................................................................ 3

(with Mississippi State University)

Alan B. Cain Simulation of Supersonic Jet Screech .......................................................................... 45

Raymond R. Cosner F/A-18 Aerodynamic Assessment ................................................................................ 37

Jerry E. Deese Launch Vehicle Flow-Field Simulation ........................................................................ 38

Linda D. Kral Turbulence Modeling for Three-DimensionalFlow Fields ........................................... 80

Medical College of Ohio

Jeffrey R. Hammersley Modeling of Pulmonary Fluid Dynamics .................................................................... 110

(with University of Arkansas and Mississippi State University)

Mississippi State University

Ramesh K. Agarwal Unstructured Grid Finite-Element Solutions ................................................................ 3

(with McDonnell Douglas Corporation)

Jeffrey R. Hammersley Modeling of Pulmonary Fluid Dynamics .................................................................... 1 10

(with University of Arkansas and Medical College of Ohio)

NASA Ames Research Center

Timothy J. Barth Parallel Implementation of an Unstructured Mesh Navier—Stokes Solver ..................... 4

(with Sterling Software Systems)

Brian J. Cantwell Incompressible Plane Wake Computation Comparisons ............................................. 76

(with Stanford University)

Neal M. Chaderjian Effects of Roll Angle on Vortex Breakdown ................................................................. 24

(with Naval Post Graduate School and Stanford University)

122
Page

Haecheon Choi Numerical Simulation of Riblets ................................................................................. 77

(with Stanford University)

Wei J. Chyu Airframe/Inlet Aerodynamics ....................................................................................... 47

(with Carnegie Mellon University)

Jeffrey N. Cuzzi Protoplanetary Nebula Particle-Gas Dynamics ............................................................ 115

(with University of California, Santa Cruz and Synernet, Inc.)

Karen L. Gundy-Burlet Unsteady Turbomachinery Computations .................................................................... 50

Guru P. Guruswamy Aeroelasticity of Wing—Body—Control Configurations .................................................. 8

(with MCAT Institute)

Guru P. Guruswamy Wing—Body Aeroelasticity ........................................................................................... 9

(with MCAT Institute)

Creon Levit Evolution of Planetary Rings ....................................................................................... 94

Thomas S. Lund Large-Eddy Simulation of Jet Combustors .................................................................... 81

(with Stanford University)

W. J. McCroskey Calculations of High Performance Rotorcraft .............................................................. 71

(with U.S. Army Aeroflightdynamics Directorate, Sterling Software Systems,

and University of California, Davis)

Parviz Moin Large-Eddy Simulation of Complex Flows ................................................................... 85

(with Stanford University)

Parviz Moin Large-Eddy Simulation of Separated External Flows .................................................... 84

(with Stanford University)

James B. Pollack Dynamics of the Martian Atmosphere ......................................................................... 98

Yehia Rizk Prediction of the Unsteady F-18 Flow Field ................................................................. 16

Michael M. Rogers Self-Similar Turbulent Plane Wakes ............................................................................. 89

(with Stanford University and University of Wisconsin, Madison)

Muriel D. Ross Finite Volume Neuronal Modeling .............................................................................. 112

(with Sterling Software Systems)

Lewis B. Schiff Investigation of Forebody Flow Control ....................................................................... 30

(with MCAT Institute)

Bruce F. Smith Studies of Galaxy Formation and Evolution ................................................................. 96

(with University of Chicago)

Roger C. Strawn Solution-Adaptive Computations of Helicopter Acoustics ........................................... 72

(with U.S. Army Aeroflightdynamics Directorate, RIACS, and

Stanford University )

E. Tu Laminar Flow Supersonic Wind Tunnel ....................................................................... 34

(with MCAT Institute)

Alex C. Woo Computational Radar Cross Sections ........................................................................... 105

(with Sterling Software Systems)

123
Page

Jerry Yan Automated Instrumentation and Monitoring System ................................................... 101

(with Recom Technologies)

Richard E. Young Simulation of Volcanic Aerosol Clouds ....................................................................... 99

S. H. Konrad Zhu Simulation of Rarefied Jet Interaction ......................................................................... 28

(with Rockwell International, Rocketdyne Division)

NASA Langley Research Center

Jassim A. AI-Saadi Transonic Flows Around Complex Geometries ........................................................... 35

(with George Washington University)

Farhad Ghaffari Applied Computational Fluid Dynamics ..................................................................... 5

(with ViGYAN, Inc.)

Kenneth M. Jones Subsonic High-Lift Analysis ........................................................................................ 29

(with ViGYAN, Inc. and Lockheed Engineering and Sciences Company)

Man Mohan Rai Direct Simulation of Turbulent/Transitional Airfoil Flow .............................................. 88

Bart A. Singer Origins of Near-Wall Turbulence ................................................................................ 90

(with High Technology Corporation)

R. C. Swanson Simulation of Scramjet Flow Fields ............................................................................. 64

(with ICASE)

James L. Thomas Complete F-18 Hybrid Solution .................................................................................. 18

NASA Lewis Research Center

David E. Ashpis Transition in a Highly Disturbed Environment ............................................................ 75

(with Boeing Commercial Airplane Group)

Rodrick V. Chima Multiblock Solver for Turbomachinery Flows .............................................................. 46

Michael C. Fischer Supersonic Laminar Flow Control Experiment ............................................................ 32

(with AS&M, Inc.)

Chunill Hah Unsteady Three-Dimensional Flow in aTransonic Compressor ................................... 51

(with ICOMP and Wright Patterson Air Force Base)

Toni I-P. Shih Flow in Turbine-Blade Cooling Passages ..................................................................... 60

(with Carnegie Mellon University)

Clifford E. Smith Jets-in-Cross-Flow Mixing ........................................................................................... 62

(with CFD Research Corporation)

ShayeYungster National Aero-Space Plane Nozzles with External Burning ......................................... 68

(with ICOMP)

124
Page

NASA Marshall Space Flight Center

James Caristi Simulating the Cosmic X-Ray Experiment .................................................

(with Valparaiso University)

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

Naval Air Warfare Center

Steven B. Kern Vortex-Flow Control .................................................................................

Naval Post Graduate School

Neal M. Chaderjian Effects of Roll Angle on Vortex Breakdown ................................................

(with NASA Ames Research Center and Stanford University)

........... 95

57

.... 1 1

.... 24

Naval Research Laboratory

Fernando F. Grinstein Simulation of Spatially Evolving Jets ............................................................................ 7

K. Kailasanath Burner-Stabilized Flames in Microgravity .................................................................... 114

(with Berkeley Research Associates )

North Carolina State University

Hassan A. Hassan Transition Model for High-Speed Flow ........................................................................ 26

Overset Methods, Inc.

Christopher A. Atwood Simulation of Coupled Fluid and Flightdynamic Systems ............................................ 36

Pennsylvania State University

Budugur Lakshminarayana Three-Dimensional Steady and Unsteady Turbulent Flows in Turbomachines .............. 55

Lyle N. Long Hybrid Massively Parallel Aeroacoustics Scheme ........................................................ 93

C. L. Merkle Computational Propulsion Applications ...................................................................... 56

Recom Technologies

Jerry Yan Automated Instrumentation and Monitoring System .................................................... 101

(with NASA Ames Research Center)

RIACS

Roger C. Strawn Solution-Adaptive Computations of Helicopter Acoustics ........................................... 72

(with U.S. Army Aeroflightdynamics Directorate, NASA Ames Research

Center, and Stanford University)

125
Page

Rockwell International Science Center

W. W. Follett Vortex Induced Mixing Behind Ramp Injectors ........................................................... 49

(with Rockwell International, Rocketdyne Division)

Vijaya Shankar Computational Fluid Dynamics Approach to Computational Flectromagnetics ........... 103

Rockwell International, North American Aircraft Division

Philip B. Gingrich Design-by-Optimization Method ................................................................................ 6

Jong H. Wang Hydrocarbon Scramjet Combustor ............................................................................. 65

Chung-Jin Woan Flow Simulation about Attack Aircraft ........................................................................ 19

David T. Yeh Multiple-Body Aerodynamics ..................................................................................... 20

Rockwell International, Rocketdyne Division

W. W. Follett Vortex Induced Mixing Behind Ramp Injectors ........................................................... 49

(with Rockwell International Science Center)

S. H. Konrad Zhu Simulation of Rarefied Jet Interaction ......................................................................... 28

(with NASA Ames Research Center)

Rockwell International, Space Systems Division

Daniel F. Dominik Space Shuttle Flow Fields with Plume Simulation ....................................................... 39

Rose Engineering and Research, Inc.

William C. Rose High-Speed Inlet Design and Analysis ........................................................................ 58

Science Applications International Corporation

Joseph D. Baum Arbitrary Lagrangian—Eulerian Three-Dimensional Simulations ................................... 106

(with George Mason University)

Scientific Research Associates, Inc.

Frederik J. de Jong Turbine Blade Tip Clearance Flows ............................................................................. 48

Stanford University

Andrew Anagnost Time-Accurate Simulation of Hypervelocity Wakes .................................................... 23

Brian J. Cantwell Incompressible Plane Wake Computation Comparisons ............................................. 76

(with NASA Ames Research Center)

Neal M. Chaderjian Effects of Roll Angle on Vortex Breakdown ................................................................. 24

(with NASA Ames Research Center and Naval Post Graduate School)

126
Page

Haecheon Choi Numerical Simulation of Riblets ................................................................................. 77

(with NASA Ames Research Center)

Sanjiva K. Lele Skewed Compressible Mixing Layers .......................................................................... 12

Thomas S. Lund Large-Eddy Simulation of Jet Combustors .................................................................... 81

(with NASA Ames Research Center)

Parviz Moin Direct Numerical Simulation of Shock/Turbulence Interaction .................................... 83

Parviz Moin Large-Eddy Simulation of Complex Flows ................................................................... 85

(with NASA Ames Research Center)

Parviz Moin Large-Eddy Simulation of Separated External Flows .................................................... 84

(with NASA Ames Research Center)

Michael M. Rogers Self-Simi larTurbulent Plane Wakes ............................................................................. 89

(with NASA Ames Research Center and University of Wisconsin, Madison)

Roger C. Strawn Solution-Adaptive Computations of Helicopter Acoustics ........................................... 72

(with U.S. Army Aeroflightdynamics Directorate, NASA Ames Research

Center, and RIACS)

Sterling Software Systems

Timothy J. Barth Parallel Implementation of an Unstructured Mesh Navier—Stokes Solver ..................... 4

(with NASA Ames Research Center)

W. J. McCroskey Calculations of High Performance Rotorcraft .............................................................. 71

(with NASA Ames Research Center, U.S. Army Aeroflightdynamics

Directorate, and University of California, Davis)

Muriel D. Ross Finite Volume Neuronal Modeling .............................................................................. 112

(with NASA Ames Research Center)

Alex C. Woo Computational Radar Cross Sections ........................................................................... 105

(with NASA Ames Research Center)

Synernet, Inc.

Jeffrey N. Cuzzi Protoplanetary Nebula Particle-Gas Dynamics ............................................................ 115

(with NASA Ames Research Center and University of California, Santa Cruz)

University of Arkansas

Jeffrey R. Hammersley Modeling of Pulmonary Fluid Dynamics ..................................................................... 110

(with Medical College of Ohio and Mississippi State University)

University of California, Davis

W. J. McCroskey Calculations of High Performance Rotorcraft .............................................................. 71

(with NASA Ames Research Center, U.S. Army Aeroflightdynamics Directorate,

and Sterling Software Systems)

127
Page

University of California, San Francisco

Andrew Pohorille Structure and Functions of Membranes .......................................................

University of California, Santa Cruz

Jeffrey N. Cuzzi Protoplanetary Nebula Particle-Gas Dynamics ............................................ 115

(with NASA Ames Research Center and Synernet, Inc.)

University of Chicago

Bruce F. Smith Studies of Galaxy Formation and Evolution ...........

(with NASA Ames Research Center)

University of Iowa

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

Fred Stern Unsteady Viscous Marine Propulsor Hydrodynamics ...

University of Michigan

John L. Volakis Electromagnetic Scattering by Airborne Structures

U n iversity of Toledo

Sang-Wook Kim Chemically Reacting Compressible Shear Layer .,

96

...... 108

...... 107

104

100

University of Wisconsin, Madison

Michael M. Rogers Self-Si mllarTUrbulent Plane Wakes ...................................................................... 89

(with NASA Ames Research Center and Stanford University)

U.S. Army Aeroflightdynamics Directorate

W. J. McCroskey Calculations of High Performance Rotorcraft ............................................................. 71

(with NASA Ames Research Center, Sterling Software Systems, and University

of California, Davis)

Roger C. Strawn Solution-Adaptive Computations of Helicopter Acoustics .......................................... 72

(with NASA Ames Research Center, RIACS, and Stanford University)

Valparaiso University

James Caristi Simulating the Cosmic X-Ray Experiment ................................................................. 95

(with NASA Marshall Space Flight Center)

128
Page

VIGYAN, Inc.

Farhad Ghaffari Applied Computational Fluid Dynamics .................................................

(with NASA Langley Research Center)

Kenneth M. Jones Subsonic High-Lift Analysis ....................................................................

(with NASA Langley Research Center and Lockheed Engineering and

Sciences Company)

W K II Ld b r . e y on en e g Reyno Id N ber Effects on a Delta Win s um g ...............................................................

Wright Patterson Air Force Base

Chunill Hah Unsteady Three-Dimensional Flow in a Transonic Compressor ................................... 51

(with NASA Lewis Research Center and (COMP)

Balu Sekar Three-Dimensional Mixing Flows in High-Speed Combustors ..................................... 59

.................... 5

.............. 29

40

129
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NAS Technical Summaries, March 1993–February 1994

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Numerical Aerodynamic Simulation Program

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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 i ntegrati ng leadi ng-edge supercomputing technologies, thus benefitti ng other supercomputer

centers in government and industry. The 1993-94 operational year concluded with 448 high-speed

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