About the cover:

Top: NASA Lewis has been responsible for developing the propulsion technologies required for a

next-generation supersonic commercial transport, or High-Speed Civil Transport (HSCT). NASA

personnel have made significant contributions to propulsion component and high-temperature

engine materials technologies that will not only contribute to future HSCT aircraft and engine

designs but will be utilized in a wide variety of other government and industry programs, both

aerospace and nonaerospace.

Bottom: The agreement between Lewis and America True calls for detailed analysis of different sail and

mast designs.
National Aeronautics and

Space Administration

Lewis Research Center

Cleveland, Ohio 44135

NASA/TM—1999-208815

Research &

Technology

1998
ii

Trade names or manufacturers' names are used in this report for identification

only. This usage does not constitute an official endorsement, either expressed

or implied, by the National Aeronautics and Space Administration.

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Price Code: A10

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Price Code: A10

Available from

This report is available in electronic form at http://www.grc.nasa.gov/WWW/RT/.

Note that at the time of printing, the NASA Lewis Research Center was undergoing a name

change to the NASA John H. Glenn Research Center at Lewis Field. URL's for the Center are

gradually being changed from lerc.nasa.gov to grc.nasa.gov. This process was just starting at

the time of this printing. If a URL given in this document does not work, try substituting

"grc" for "lerc" or "lerc" for "grc", wherever they appear.
iii

NASA Lewis Research Center is responsible for developing

and transferring critical technologies that address national

priorities in aerospace propulsion and space applications

in partnership with U.S. industries, universities, and other

Government institutions.

As NASA’s premier Center for aerospace propulsion, our

role is to develop, demonstrate, and transfer relevant

technologies to U.S. industry. As NASA’s designated Center

of Excellence in Turbomachinery, we develop new and

innovative technologies that improve the reliability, performance, efficiency, affordability, capacity, and environmental compatibility of turbomachinery-based propulsion

systems. We also maintain a science and technology

development role in communications, space power,

onboard propulsion, and microgravity fluid physics and combustion. We are committed to

enabling U.S.-based aerospace and nonaerospace industries to benefit directly from the technologies developed through our programs. Our goal is to maximize the benefit of our efforts

to the Nation and to optimize the return on each taxpayer’s investment.

Over 3600 civil service employees and support service contractor personnel staff Lewis. Scientists

and engineers comprise more than half of our workforce, with technical specialists, skilled workers, and an administrative staff supporting them. We aggressively strive for technical excellence

through continuing education, diversity in our workforce, and continuous improvement in our

management and business practices so we can extend the edge of technology in space and

aeronautics.

The Lewis Research Center is a unique facility located in the southwest corner of Cleveland,

Ohio. Situated on 350 acres of land adjacent to the Cleveland Hopkins International Airport,

Lewis comprises more than 140 buildings, including 24 major facilities and over 500 specialized

research and test facilities. Additional facilities are located at the Plum Brook Station, which is

about 50 miles west of Cleveland.

Our end product is knowledge. This report is designed to help make this knowledge fully available to potential users—the aircraft engine industry, the space industry, the energy industry, the

automotive industry, the aerospace industry, and others. It is organized so that a broad cross

section of the community can readily use it. Each article begins with a short introductory paragraph that should prove to be a valuable tool for the layperson. These articles summarize the

progress made during the year in various technical areas and portray the technical and administrative support associated with Lewis’ technology programs. We hope that the information is

useful to all.

At the time of publication, NASA Lewis was undergoing a name change to the NASA John H.

Glenn Research Center at Lewis Field. If additional information is desired, readers are encouraged to contact the researchers identified in the articles and to visit NASA Glenn on the World

Wide Web at http://www.grc.nasa.gov. This document is available on the World Wide Web

(http://www.grc.nasa.gov/WWW/RT/).

Donald J. Campbell

Director

Introduction
Deputy Director Deputy Director

for Operations

Julian M. Earls Marvin E. Goldstein John W. Gaff

Assistant Deputy

Director for Policy

Office of the

Chief Counsel

J. William Sikora

Office of Safety, Environmental,

and Mission Assurance

Vernon W. Wessel

Aeropropulsion Research

Program Office

Frank D. Berkopec

Chief

Scientist

Plans and 

Programs Office

Office of Equal

Opportunity Programs

Belinda M. Roberts Olga D. Gonzalez-Sanabria

Office of the Chief

Financial Officer

Robert E. Fails

Office of Human

Resources

Maury L. Blanton

Office of

Acquisition

Julian M. Earls*

External

Programs

John M. Hairston, Jr.

Engineering and 

Technical Services

Randall B. Furnas

Space

Gerald J. Barna

Research and 

Technology

Woodrow Whitlow, Jr.

Aeronautics

Carol J. Russo

CD-48534February 1999

Director

Donald J. Campbell Martin P. Kress

*Acting

ISO Senior Project

Manager

Jeffrey E. Haas

NASA Lewis Research Center Senior Management

iv
v

Est. 10-1-96CD-44972Rev. 2-16-99

Resources

Analysis and

Management

Office

Financial

Management

Division

Special

 Projects

Office

Hypersonics

Projects

Office

Office of the

Chief Financial Officer

Office of

Equal Opportunity

Plans and Programs

Office

Aeropropulsion Research

Program Office

Director

Deputy Director

Deputy Director for Operations

Assistant Deputy Director for Policy

Chief Scientist

0141 0142 0210 0220

0100

Quality

Management

Office

Institutional

Safety

Office

Environmental

Management

Office

Security

Management

Office

Office of Safety, Environmental 

and Mission Assurance

0510 0520 0530 0540 0550

0500

Office of

the Chief Counsel

0120

0140 0170 0180 0200

NASA Office of

Inspector General

0160

Army Office

0300

Procurement

Division

0610

Logistics 

and Technical

Information

Division

0620

Office of

Acquisition

0600

Office of

Human Resources

0400

Power and On-Board

Propulsion

Technology

Division

Instrumentation

and Controls

Division

Research and Technology 

Directorate

5100 5400 5500

Communications

Technology

Division

5600

Materials

Division

Structures

and Acoustics

Division

5800 5900

Turbomachinery

and Propulsion

Systems

Division

5000 5010

Office

of

UniversityPrograms

Plum Brook

Management

Office

Aircraft

Operations

Office

Computer

Services

Division

7030 7040 7100

Engineering and Technical 

Services Directorate

7000

Test

Installations

Division

7200

Manufacturing

Engineering

Division

Facilities

and Test

Engineering

Division

7400 7500

Engineering

Design

and Analysis

Division

Systems

Engineering

Division

7700 7800

Subsonic

Systems

Office

High-Speed

Systems

Office

Aeronautics Directorate

2200 2300

Propulsion

Systems

Analysis

Office

Computing

and Interdisciplinary

Systems

Office

2400 2900

2000

Space

Communications

Office

Space 

Transportation

Project

Office

Space Directorate

6100 6500

Microgravity

Science

Division

Power

and Propulsion

Office

6700 6900

6000

Educational

Program

Office

Community

and Media

Relations

Office

Commercial

Technology

Office

External Programs

Directorate

9200 9300 9400

9000

Aviation Safety

Program

Office

2500

Project

Assurance

Office

0510

0106

ISO

Project

Office

NASA Lewis Research Center
vi

Carol J. Russo

Director of Aeronautics

Virginia A. Cantwell

President LESA

Julian M. Earls

Deputy Director for

Operations

Robert E. Fails

Chief Financial Officer

Gerald J. Barna

Director of Space

John M. Hairston, Jr.

Director of External Programs

Donald J. Campbell

Director

Quality Council Dennis M. Sender Special Assistant to the Quality Council

Martin P. Kress

Deputy Director The Quality Council was established in October 1992 to adopt

and implement a Total Quality (TQ) plan for Lewis. It is composed of Executive Council members as well as the presidentof the American Federation of Government Employees (AFGE),Local 2812, and the president of the Lewis Engineers andScientists Association (LESA), IFPTE Local 28. A representative of major onsite support service contractors serves as a liaison. February 1999

Marvin E. Goldstein

Chief Scientist

Donald R. Streibing

President AFGE

Randall B. Furnas

Director of Engineering

and Technical Services

Christopher J. Pestak

Support Service

Contractor Liaison

Woodrow Whitlow, Jr.

Director of Research

and Technology
vii

Contents

Aeronautics

High-Speed Systems

High-Speed Research—Propulsion Project Accomplishments . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

Propulsion Systems Analysis

Neural Network and Regression Approximations Used in Aircraft Design . . . . . . . . . . . . . . . . . . 3

Computing and Interdisciplinary Systems

Numerical System Solver Developed for the National Cycle Program . . . . . . . . . . . . . . . . . . . . . . 4

Inlet-Compressor Analysis Performed Using Coupled Computational Fluid

Dynamics Codes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

FoilSim—Basic Aerodynamics Software Created . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

Common Analysis Tool Being Developed for Aeropropulsion: the National Cycle Program

Within the Numerical Propulsion System Simulation Environment . . . . . . . . . . . . . . . . . . . . . 9

Flow of GE90 Turbofan Engine Simulated . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

Research and Technology

Materials

MoSi2-Base Structural Composite Passed Engine Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

Deleterious Phase Formation in Next-Generation Nickel-Base Superalloys Predicted . . . . . . . 15

Affordable Manufacturing Technologies Being Developed for Actively Cooled

Ceramic Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

Sensor Lead Wires Positioned on SiC-Based Monolithic Ceramic and Fiber-Reinforced

Ceramic Matrix Composite Subcomponents With Flat and Curved Surfaces . . . . . . . . . . . . . 17

Modal Acoustic Emission Used at Elevated Temperatures to Detect Damage and

Failure Location in Ceramic Matrix Composites . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

High-Temperature Solid Lubricants Developed by NASA Lewis Offer Virtually

“Unlimited Life” for Oil-Free Turbomachinery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

New Vapor/Mist Phase Lubricant Formulated . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

Monte Carlo Simulation of Alloy Design Techniques: Fracture and Welding

Studied Using the BFS Method for Alloys . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

Low-Cost, High Glass-Transition Temperature, Thermosetting Polyimide Developed . . . . . . . . 25

Novel Ultraviolet-Light-Curable Polyimides . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

10 000-hr Cyclic Oxidation Behavior of 68 High-Temperature Co-, Fe-, and Ni-Base

Alloys Evaluated at 982 °C (1800 °F) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

Single-Crystal NiAl-X Alloys Tested for Hot Corrosion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

Thermal Mechanical Stability of Single-Crystal-Oxide Refractive Concentrators

Evaluated for High-Temperature Solar-Thermal Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

Power and On-Board Propulsion Technology

Solar Array Verification Analysis Tool (SAVANT) Developed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

Space Plasma Shown to Make Satellite Solar Arrays Fail . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

Lithium-Ion Polymer Rechargeable Battery Developed for Aerospace and

Military Applications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

Cathodes Delivered for Space Station Plasma Contactor System . . . . . . . . . . . . . . . . . . . . . . . . 37

High-Performance Bipropellant Engine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37

Liquid Bipropellant Microrocket Concept Under Development . . . . . . . . . . . . . . . . . . . . . . . . . . 38

Ion Thruster and Power Processor Developed for the Deep Space 1 Mission . . . . . . . . . . . . . . 39

Advanced Refrigerator/Freezer Technology Development Project . . . . . . . . . . . . . . . . . . . . . . . . 40

Power System Optimization Codes Modified . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
viii

Lightweight Battery Charge Regulator Used to Track Solar Array Peak Power . . . . . . . . . . . . . . 42

International Low-Earth-Orbit Spacecraft Materials Test Program Initiated for

Better Prediction of Durability and Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

Retrieval of Mir Solar Array . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

Thermal Control Replacement Materials Evaluated for Durability and Selected for the

Hubble Space Telescope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

Effects Investigated of Ambient High-Temperature Exposure on Alumina-Titania

High-Emittance Surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46

Polymers Erosion and Contamination Experiment Being Developed . . . . . . . . . . . . . . . . . . . . . 47

Synchrotron Vacuum Ultraviolet Light and Soft X-Ray Radiation Effects on

Aluminized Teflon FEP Investigated . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48

Teflon FEP Analyzed After Retrieval From the Hubble Space Telescope . . . . . . . . . . . . . . . . . . . 50

Highly Conducting Graphite Epoxy Composite Demonstrated . . . . . . . . . . . . . . . . . . . . . . . . . . . 51

Changes in the Optical Properties of Materials Are Observed After 18 Months

in Low Earth Orbit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52

Low-Temperature Electronic Components Being Developed . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53

Atomic Oxygen Lamp Cleaning Facility Fabricated and Tested . . . . . . . . . . . . . . . . . . . . . . . . . . . 54

Parallel Stirling Converters Being Developed for Spacecraft Onboard Power . . . . . . . . . . . . . . . 55

Inflatable Solar Thermal Concentrator Delivered . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56

Instrumentation and Controls

Acoustic Pyrometry Applied to Gas Turbines and Jet Engines . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

Scanning Mode Shock Position Sensor Invented and Demonstrated . . . . . . . . . . . . . . . . . . . . . 58

Pressure-Sensitive Paint Applied to Ice Accretions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

Neural Networks Used to Compare Designed and Measured Time-Average Patterns . . . . . . . . 60

Nondestructive Strain Measurement System Used to Determine Surface Strain on Fibers . . . . 61

High Reliability Engine Control Demonstrated for Aircraft Engines . . . . . . . . . . . . . . . . . . . . . . . 62

Distortion Tolerant Control Flight Demonstration Shown to Be Successful . . . . . . . . . . . . . . . . 64

Combustion Instabilities Modeled . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66

Communications Technology

Web Transfer Over Satellites Being Improved . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67

Advances Made in the Next Generation of Satellite Networks . . . . . . . . . . . . . . . . . . . . . . . . . . . 67

Internet Protocol Enhanced Over Satellite Networks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68

Satellite ATM Networks: Architectures and Guidelines Developed . . . . . . . . . . . . . . . . . . . . . . . 69

Broadband Satellite Technologies and Markets Assessed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

A More Accurate and Efficient Technique to Obtain Helical Traveling-Wave Tube Interaction

Impedance Using Computational Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

Diamond Analyzed by Secondary Electron Emission Spectroscopy . . . . . . . . . . . . . . . . . . . . . . . 73

Finite Ground Coplanar (FGC) Waveguide: Characteristics and Advantages Evaluated for

Radiofrequency and Wireless Communication Circuits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74

Angular Distribution of Elastically Scattered Electrons Determined and Its Effect on

Collector Performance Computed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76

Thin-Film Ferroelectric Tunable Microwave Devices Being Developed . . . . . . . . . . . . . . . . . . . . 78

Technique Developed for Optimizing Traveling-Wave Tubes . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80

Cassini Radio Science Experiments on Saturn and Titan Preserved Because of

 Lewis Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81

Novel Low-Cost, Low-Power Miniature Thermionic Cathode Developed for

Microwave/Millimeter Wave Tube and Cathode Ray Tube Applications . . . . . . . . . . . . . . . . . 82

K-Band Phased Array Developed for Low-Earth-Orbit Satellite Communications . . . . . . . . . . . 84

Conformal, Transparent Printed Antenna Developed for Communication and

Navigation Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86

Innovative, Inexpensive Etching Technique Developed for Polymer ElectroOptical Structures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88

High-Performance, Low-Complexity Codes Researched for Communication Channels . . . . . . 88
ix

Reconfigurable Data Communications Packet-Switch Emulation Test Bed Demonstrated . . . . 90

Bandwidth-Efficient Wireless Digital Modem Developed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 91

622-Mbps Orthogonal Frequency Division Multiplexing Modulator Developed . . . . . . . . . . . . . 92

Silicon-Germanium Fast Packet Switch Developed for Communications Satellites . . . . . . . . . . 93

Turbomachinery and Propulsion Systems

Multistage Turbomachinery Flows Simulated Numerically . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 94

Optical Flow-Field Techniques Used for Measurements in High-Speed

Centrifugal Compressors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 95

LeRC–HT: NASA Lewis Research Center General Multiblock Navier-Stokes

Heat Transfer Code Developed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 97

Research Data Acquired in World-Class, 60-atm Subsonic Combustion Rig . . . . . . . . . . . . . . . . 98

X–33 Combustion-Wave Ignition System Tested . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99

Use of Atomic Fuels for Rocket-Powered Launch Vehicles Analyzed . . . . . . . . . . . . . . . . . . . . . 101

NASA/FAA Tailplane Icing Program Completed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102

National Transportation Safety Board Aircraft Accident Investigation Supported . . . . . . . . . . . 104

Microelectrical Mechanical Systems Flow Control Used to Manage Engine Face

Distortion in Compact Inlet Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105

Two-Dimensional Bifurcated Inlet/Engine Tests Completed in 10- by 10-Foot

Supersonic Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 107

WIND Flow Solver Released . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 108

Results of Liquid Motion Experiment Analyzed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 109

Rapid Chill and Fill of a Liquid Hydrogen Tank Demonstrated . . . . . . . . . . . . . . . . . . . . . . . . . . 110

Zero Boiloff Storage of Cryogenic Propellants Achieved at Lewis’ Supplemental

Multilayer Insulation Research Facility . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 112

Engineering Application Integration Architecture Devised . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 113

Space-Time Conservation Element and Solution Element Method Being Developed . . . . . . . 114

Structures and Acoustics Division

Nondestructive Evaluation Correlated With Finite Element Analysis . . . . . . . . . . . . . . . . . . . . . 117

Deformation Behaviors of HIPped Foil Compared With Those of Sheet Titanium Alloys . . . . 119

Thermomechanical Fatigue Durability of T650–35/PMR–15 Sheet-Molding

Compound Evaluated . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 120

“Ultra”-Fast Fracture Strength of Advanced Structural Ceramic Materials Studied at

Elevated Temperatures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 122

Viscoplastic Constitutive Theory Demonstrated for Monolithic Ceramic Materials . . . . . . . . . . 123

Resistance of Titanium Aluminide to Domestic Object Damage Assessed . . . . . . . . . . . . . . . . 124

Multiaxial Experiments Conducted to Aid in the Development of Viscoplastic Models . . . . . . 125

Multidisciplinary Probabilistic Heat Transfer/Structural Analysis Code

Developed—NESTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 126

Design Tool Developed for Probabilistic Modeling of Ceramic Matrix Composite Strength . . 127

CARES/Life Ceramics Durability Evaluation Software Enhanced for Cyclic Fatigue . . . . . . . . . 129

Novel Method Used to Inspect Curved and Tubular Structural Materials . . . . . . . . . . . . . . . . . 130

Composite Nozzle/Thrust Chambers Analyzed for Low-Cost Boosters . . . . . . . . . . . . . . . . . . . 131

Flutter Version of Propulsion Aeroelasticity Code Completed . . . . . . . . . . . . . . . . . . . . . . . . . . 132

Stability of the Tilt Modes of an Actively Controlled Flywheel Analyzed . . . . . . . . . . . . . . . . . . 133

Dual-Laser Probe for Measuring Blade-Tip Clearance Tested . . . . . . . . . . . . . . . . . . . . . . . . . . . 135

Time-Shifted Boundary Conditions Used for Navier-Stokes Aeroelastic Solver . . . . . . . . . . . . 135

Low-Noise Potential of Advanced Fan Stage Stator Vane Designs Verified in NASA Lewis

Wind Tunnel Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 137

NASA Space Mechanisms Handbook—Lessons Learned Documented . . . . . . . . . . . . . . . . . . . 138

Metrology Evaluation of Superfinished Gears Completed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140

Three-Dimensional Gear Crack Propagation Studied . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 141

NASA Lewis Thermal Barrier Feasibility Investigated for Use in Space Shuttle Solid-Rocket

Motor Nozzle-to-Case Joints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 143
x

Space

Space Communications

Advanced Communications Technology Satellite Now Operating in an Inclined Orbit . . . . . . 148

Microgravity Science

DARTFire Sees Microgravity Fires in a New Light—Large Data Base of Images Obtained . . . . 149

Solid Surface Combustion Experiment Yields Significant Observations . . . . . . . . . . . . . . . . . . 151

Stability of Enclosed Laminar Flames Studied in Microgravity . . . . . . . . . . . . . . . . . . . . . . . . . . 153

Unlocking the Keys to Vortex/Flame Interactions in Turbulent Gas-Jet Diffusion

Flames—Dynamic Behavior Explored on the Space Shuttle . . . . . . . . . . . . . . . . . . . . . . . . . . 154

Crystal Growth Using MEPHISTO . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 156

Effects of g-Jitter on Diffusion in Binary Liquids . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 158

Viscosity of Xenon Examined in Microgravity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 160

Laser Light Scattering With Multiple Scattering Suppression Used to Measure

Particle Sizes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 161

Phase-Shifted Laser Feedback Interferometry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 163

Microgravity Experiments Being Developed for Microscopic Study of Colloidal Spheres . . . . 164

Onset of Convection Due to Surface Tension Variations in Multicomponent and

Binary Fluid Layers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 166

Space Acceleration Measurement System for Free Flyers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 168

Binary Colloidal Alloy Test Conducted on Mir . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 170

Coarsening in Solid-Liquid Mixtures Studied on the Space Shuttle . . . . . . . . . . . . . . . . . . . . . . 171

Third and Final Shuttle Mission of the Isothermal Dendritic Growth Experiment

Conducted—Highest Supercooling Ever Recorded Achieved . . . . . . . . . . . . . . . . . . . . . . . . . 172

Tracker—Image-Processing and Object-Tracking System Developed . . . . . . . . . . . . . . . . . . . . . 174

Engineering and Technical Services

Computer Services

NASA Lewis Launch Collision Probability Model Developed and Analyzed . . . . . . . . . . . . . . . 178

Manufacturing Engineering

Customized Hermetic Feedthrough Developed to Isolate Fluids . . . . . . . . . . . . . . . . . . . . . . . . 179

Engineering Design and Analysis

Pyroshock Environments Characterized for Spacecraft Missions . . . . . . . . . . . . . . . . . . . . . . . . 180

Traversing Microphone Track Installed in NASA Lewis’ Aero-Acoustic Propulsion

Laboratory Dome . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 182

Virtual Interactive Classroom: A New Technology Developed for Distance Learning . . . . . . . . 183

Commercial Technology

Silicon Carbide Being Developed for High-Definition Television Transmitter Modules . . . . . . 186

NASA Has Joined America True’s Design Mission for 2000 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 187

National Combustion Code, a Multidisciplinary Combustor Design System, Will Be

Transferred to the Commercial Sector . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 187

Appendixes

Definitions of NASA Headquarters Program Offices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 189

Definitions of Programs and Projects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 190

Index of Authors and Contacts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 192
1
2

This past year has been one of great technical accomplishment for the

propulsion element of NASA’s High-Speed Research (HSR) Program. The

HSR Program is a NASA/industry partnership to develop the high-risk/

high-payoff airframe and propulsion technologies applicable to a secondgeneration supersonic commercial transport, or High-Speed Civil Transport (HSCT). The NASA Lewis Research Center manages the propulsion

element, which also includes industry participation. These technologies

will contribute greatly to U.S. industry’s ability to make an informed product launch decision for an HSCT vehicle. Specific NASA Lewis accomplishments in 1998 include

· The lean premixed prevaporized combustor concept was selected for

follow-on full-scale development. The combustor concept has demonstrated the ability to produce ultra-low emissions of less than 5 g of

nitrogen oxides (NOx) per kg of fuel burned while maintaining a combustion efficiency of greater than 99.9 percent with minimal technical or

developmental risks.

· Combustor sector tests were initiated in CE–9 of the Engine Research

Building to evaluate ceramic matrix composite parts in a realistic combustor environment.

· An alternative metal combustor liner was assessed. These tests showed

that metal liners made from nickel-based superalloys and coated with a

thick thermal barrier coating can meet HSR emission and life goals.

High-Speed Systems

· The Large Scale Model 1 (LSM–1)

test, which was conducted to test

a nozzle design in a large-scale,

more realistic engine environment,

was completed. The goal was to

improve confidence in the ability

of the nozzle technologies being

developed to meet the noise goals

of the HSR Program. The test not

only showed that the technologies are on track but successfully

demonstrated the use of ceramic

matrix composite tiles in HSCT

engine applications.

· Over 18 000 hours of durability

testing were completed for all four

advanced materials systems in

Lewis’ materials laboratories.

Superalloy, gamma-titanium aluminide, ceramic matrix composite,

and thermal protection system

performance were evaluated for

HSCT nozzle applications.

· A turbomachinery disk alloy was

selected for a large-scale processing effort. This alloy has mechanical

properties far superior to current

production materials and is projected to meet HSCT requirements.

· A concept for a second generation, two-dimensional, mixedcompression supersonic inlet was

coupled to a turbojet engine and

tested in Lewis’ 10- by 10-Foot

Supersonic Wind Tunnel. The

extensive data base of results for

inlet/engine operability and unstart

conditions contributed to the

maturation of inlet technologies

for future commercial supersonic

transport vehicles.

· Proactive transfer of HSR propulsion technologies to other NASA

programs has been initiated.

Specifically, an economic fabrication technique for sheet titanium

High-Speed Civil Transport (HSCT).

High-Speed Research—Propulsion Project Accomplishments
3

aluminide developed for the HSCT exhaust nozzle is being considered for

use in the thermal protection system for both the reusable launch vehicle

and X–33 programs.

For more information, visit Lewis’ HSR and aeropropulsion facility sites on the

World Wide Web:

HSR: http://www.grc.nasa.gov/WWW/HSR

Engine Research Building:

http://www.grc.nasa.gov/WWW/AFED/facilities/erb.html

Propulsion Systems Laboratory:

http://www.grc.nasa.gov/WWW/AFED/facilities/psl.html

Aero-Acoustic Propulsion Laboratory:

http://www.grc.nasa.gov/WWW/AFED/facilities/aapl.html

10- by 10-Foot Supersonic Wind Tunnel:

http://www.grc.nasa.gov/WWW/AFED/facilities/10x10.html

NASA Lewis Research Center’s CometBoards Test Bed was used to create

regression and neural network models for a High-Speed Civil Transport

(HSCT) aircraft. Both approximation models that replaced the actual

analysis tool predicted the aircraft response in a trivial computational effort.

The models allow engineers to quickly study the effects of design variables

on constraint and objective values for a given aircraft configuration. For

example, an engineer can change the engine size by 1000 lb of thrust and

quickly see how this change affects all the output values without rerunning

the entire simulation. In addition, an engineer can change a constraint and

Neural Network and Regression Approximations Used in

Aircraft Design

Aeronautics

use the approximation models to

quickly reoptimize the configuration. Generating the neural network

and the regression models is a

time-consuming process, but this

exercise has to be carried out only

once. Furthermore, an automated

process can reduce calculations

substantially.

Actual

Neural network

Regression method

Analyzer

(a) 0.9

1.0

1.1

1.2

Engine thrust, lb

Normalized weight

35 37 39 43 41 45x103

0.9

1.0

1.1

1.2

Wing area, ft2

Normalized weight

(b)

75 80 85 90x102

Actual

Neural network

Regression method

Analyzer

Comparison of aircraft weight obtained by using approximate methods and actual analysis. (a) As a function of engine thrust.

(b) As a function of wing area.

Lewis contacts:

Dr. Robert J. Shaw, (216) 977–7135,

Robert.J.Shaw@grc.nasa.gov;

and Lori A. Manthey, (216) 433–2484,

Lori.A.Manthey@grc.nasa.gov

Author: Lori A. Manthey

Headquarters program office: OAT

Programs/Projects: HSR, RLV, X–33

Propulsion Systems Analysis
4

As part of the National Cycle Program (NCP), a powerful new numerical

solver has been developed to support the simulation of aeropropulsion

systems. This software uses a hierarchical object-oriented design. It can

provide steady-state and time-dependent solutions to nonlinear and even

discontinuous problems typically encountered when aircraft and spacecraft

propulsion systems are simulated. It also can handle constrained solutions,

in which one or more factors may limit the behavior of the engine system.

Time-dependent simulation capabilities include adaptive time-stepping and

synchronization with digital control elements. The NCP solver is playing an

important role in making the NCP a flexible, powerful, and reliable simulation package.

The NCP solver uses a modified Newton-Raphson method. For any given

system model of interest, a set of independent and dependent parameters

are defined, either by the user or automatically by the model elements.

The solver characterizes the effects of each independent parameter on all

dependent variables through direct perturbation of the model. It uses this

information as a guide as it attempts to find the set of independent values

that will render all dependent errors equal to zero (the solution). This is

typically achieved through a number of successive iterations. The NCP

solver employs a sophisticated set of algorithms to enhance the model’s

convergence toward the solution, and it has a number of attributes that

allow expert users to fine tune the solution process.

The NCP solver has a flexible, object-oriented design that allows appropriate solver information to be embedded in the model, even though the

Numerical System Solver Developed for the National

Cycle Program

solver itself is part of the NCP

infrastructure and is not model

specific. The functions of the solver

are also hierarchical, as illustrated in the figure. The Transient

Executive controls the overall

performance of time-dependent

simulations, which consist of one

or more individual solutions to the

model at different points in time.

The Transient Executive sends

messages to the Solver Executive,

requesting a solution at each individual point in time. The Solver

Executive, in turn, sends messages

to the independents and dependents to perform low-level functions as required (including integration with respect to time). This

hierarchical approach provides a

great deal of design flexibility

and minimizes the need to keep

multiple copies of data during

execution.

One issue that needs to be addressed is the generation of output for a

given set of design variables. The analysis tools must be smart enough to

respond completely to changes in design values. For example, a change

in the overall pressure ratio of an engine will affect several behavior

parameters besides an increase in the pressure. For example, the overall

efficiency of the compressor will change, the cooling flow temperature will

rise, more cooling flow will be required, and the compressor will increase

in weight. These problems, however, can be overcome with more capable

tools and a little planning.

The figure shows how closely the neural network and regression techniques track the output value of aircraft gross weight from the actual

analysis. Both approximation techniques have been shown to produce

good agreement for all desired response parameters.

Lewis contacts:

Dr. Surya N. Patnaik, (216) 433–5916, Surya.N.Patnaik@grc.nasa.gov;

Dale A. Hopkins, (216) 433–3260, Dale.A.Hopkins@grc.nasa.gov; and

Thomas M. Lavelle, (216) 977–7042, Thomas.M.Lavelle@grc.nasa.gov

Authors: Dr. Surya N. Patnaik, Dale A.

Hopkins, and Thomas M. Lavelle

Headquarters program office: OAT

Programs/Projects: HSCT, AST, HSR

Computing and Interdisciplinary Systems
5

The NCP solver was developed at the NASA Lewis Research Center in

cooperation with experts at each of the industry partner companies. It

includes the best practices of each partner company in a single package.

The original version of the solver included steady-state analysis capabilities

only. In fiscal year 1998, time-dependent simulation capabilities were

added, including support for simulation of the digital control elements.

In fiscal year 1999, the NCP solver will be enhanced to include special

handling for high-frequency, spatially distributed models (dynamics). A

framework for optimization will also be incorporated in the solver in 1999.

Bibliography

Evans, A.L.; Follen, G.; and Naiman, C.: Numerical Propulsion System Simulation’s

National Cycle Program. AIAA Paper 98–3113, 1998.

Ashleman, R.H.; Lavelle, T.; and Parsons, F.: The National Cycle Program: A Flexible

System Modeling Architecture for Aircraft Engine Simulation. AIAA Paper 98–3114,

1998.

Independent 1

Independent 2

Independent 3

Independent 4

Independent 5

Dependent 1

Dependent 2

Dependent 3

Dependent 4

Dependent 5

Solver Executive

Transient Executive

Set inputs

Get outputs

Set

independents Set time Execute once Get errors

User command or external code interface

Run a transient 

Run a single simulation over time

steady-state point

Run to convergence

at a single point in time

Engine

model

NCP solver subsystem hierarchy.

Dynacs Engineering Co., Inc. contact:

Michael P. Binder, (216) 977–1454,

Michael.P.Binder@grc.nasa.gov

Lewis contacts:

James L. Felder, (216) 433–3956,

James.L.Felder@grc.nasa.gov; and

Cynthia G. Naiman, (216) 433–5238,

Cynthia.G.Naiman@grc.nasa.gov

Author: Michael P. Binder

Headquarters program office: OAT

Programs/Projects: NCP, HPCCP, HSR

Aeronautics
6

Inlet-Compressor Analysis Performed Using Coupled

Computational Fluid Dynamics Codes

References

1. Freund, D.; and Sajben, M.: Reflection of Large Amplitude Acoustic

Pulses From an Axial Flow Compressor. AIAA Paper 97–2879, 1997.

2. Suresh, A., et al.: Analysis of Compressor-Inlet Acoustic Interactions

Using Coupled Component Codes.

AIAA Paper 99–0749, 1999.

Lewis contact:

Gary L. Cole, (216) 433–3655,

Gary.L.Cole@grc.nasa.gov

Dynacs Engineering Co., Inc. contacts:

Ambady Suresh, (216) 977–1384,

Ambady.Suresh@grc.nasa.gov; and

Scott Townsend, (216) 977–1080,

Scott.Townsend@grc.nasa.gov

Authors: Gary L. Cole, Ambady Suresh,

and Scott Townsend

Headquarters program office: OAT

Programs/Projects: HPCCP, NPSS

A thorough understanding of dynamic interactions between inlets and

compressors is extremely important to the design and development of

propulsion control systems, particularly for supersonic aircraft such as the

High-Speed Civil Transport (HSCT). Computational fluid dynamics (CFD)

codes are routinely used to analyze individual propulsion components. By

coupling the appropriate CFD component codes, it is possible to investigate inlet-compressor interactions. The objectives of this work were to gain

a better understanding of inlet-compressor interaction physics, formulate a

more realistic compressor-face boundary condition for time-accurate CFD

simulations of inlets, and to take a first step toward the CFD simulation of

an entire engine by coupling multidimensional component codes. This

work was conducted at the NASA Lewis Research Center by a team of civil

servants and support service contractors as part of the High Performance

Computing and Communications Program (HPCCP).

An inlet-compressor experiment (ref.1) conducted at the University of

Cincinnati was chosen as the application for this CFD analysis because

of the availability of experimental data for validation. A schematic of the

experiment is shown in the figure. It consists of a constant-area annular

inlet coupled to a GE T–58 engine (General Electric) that was modified for

cold operation. The collapse of a flexible bump on the hub of the inlet

produced a well-defined acoustic pulse that traveled downstream and

interacted with the engine. Static-pressure time histories were measured

at the four axial locations shown in the figure. NPARC, a general purpose

CFD code, was used to simulate the inlet portion of the experiment. The

engine was approximated by the first-stage rotor and was simulated with

ADPAC, a turbomachinery code. Details concerning the code coupling

approach and the issues involved are given in reference 2.

The NPARC domain was solved using a single grid block, whereas the

ADPAC domain was divided into multiple blocks, with all the blocks

computed in parallel to achieve faster execution. Simulation of the experiment for a 10.4-msec transient took about 32.5 hours for a seven-block

ADPAC case and 12 hours for a 21-block case. A snapshot in time of flowfield pressure contour plots and the four pressure-sensor time histories

are shown in the figure. (Only part of the NPARC domain is shown.)

Although not shown in the figure, the computed pressure time histories

were found to give reasonable agreement with the experimental data.

The investigation indicates that coupling inlet and turbomachinery CFD

codes is a feasible approach to studying inlet-engine interaction problems.

Additional investigation of code coupling issues and the means for reducing execution time is planned. This approach offers the possibility of

including other specialized codes (e.g., combustor) to provide a full engine

simulation. The coupled NPARC–ADPAC codes could also serve as a test

bed for exploring other flow perturbations of interest and for validating

simplified boundary conditions.
7

Snapshot of (partial) NPARC–ADPAC pressure contour plots and pressure time histories at sensor locations during

simulation of the University of Cincinnati experiment. All dimensions in diagram at top are in centimeters.

Flexible

bump

Station

number

Center of first stage rotor

Outlet guide Transducer Case vane

Hub

Variable inlet 

guide vane

45.1 22.9 20.3 17.8 8.3

1 2 3 4

C L

Aeronautics
8

FoilSim—Basic Aerodynamics Software Created

Airfoil view panel within FoilSim.

FoilSim is interactive software that simulates the airflow around various

shapes of airfoils. The graphical user interface, which looks more like a

video game than a learning tool, captures and holds the students’ interest.

The software is a product of NASA Lewis Research Center’s Learning

Technologies Project, an educational outreach initiative within the High

Performance Computing and Communications Program (HPCCP).

Originally, the code was written for

college level engineering students.

Adjustment of the code to the high

school level resulted in FoilSim,

which was created and tested by a

diverse team composed of NASA

employees, contractors, educators,

and students. Lewis’ Learning

Technologies Project is a part of

NASA’s agencywide Learning

Technologies Project, which is

managed by the NASA Ames

Research Center.

Find out more about the Learning

Technologies Project:

http://learn.ivv.nasa.gov/

Download FoilSim from the Web:

http://www.grc.nasa.gov/WWW/K-12/

aerosim

Find supporting resources:

http://www.grc.nasa.gov/WWW/K-12

Lewis contact:

Beth Lewandowski, (216) 433–8873,

Beth.E.Lewandowski@grc.nasa.gov

Author: Ruth A. Petersen

Headquarters program office: OAT

Programs/Projects: HPCCP

This airfoil view panel is a simulated view of a wing being tested in a wind

tunnel. As students create new wing shapes by moving slider controls that

change parameters, the software calculates their lift. FoilSim also displays

plots of pressure or airspeed above and below the airfoil surface. A satisfied user comments, “To be able to change the parameters of the experiment with a click of a button is fantastic, as is the ability to observe the

results instantaneously.” An additional feature of the program is “Play Ball,”

where students learn more about aerodynamics through controlling the

conditions of a baseball pitch, including altitude, speed of pitch, and spin

of pitch.

Interactive lessons that accompany the package prompt students to engage

in problem solving and discovery. Teachers are impressed at how well the

software and lessons inspire students to explore their intellectual potential.

One teacher says, “FoilSim allows students to hypothesize and provides

them with immediate, nonthreatening feedback. They really seem to begin

to appreciate the process of experimenting rather than just getting the

correct answer.”
9

Common Analysis Tool Being Developed for Aeropropulsion:

the National Cycle Program Within the Numerical Propulsion

System Simulation Environment

The NPSS/NCP team consists of

propulsion experts and software

engineers from GE Aircraft Engines,

Pratt & Whitney, the Boeing Company, AlliedSignal Engines, Allison,

Williams International, Teledyne

Ryan Aeronautical, Arnold Engineering Development Center,

Wright Patterson Air Force Base,

and NASA Lewis. Formal software

development processes are followed to facilitate technology

transfer.

In fiscal year 1998, the major

accomplishment of the NCP team

was the completion and distribution of NCP Version 1. New capabilities include transient analysis,

test data reduction, a controls

toolbox, and initial zooming. NCP

applications to the NPSS architecture successfully demonstrated

distributed simulations. Workshops

were conducted throughout the

year, training over 60 NCP users

and component developers. NCP is

being used as the modeling tool

for in-house, High-Speed Research

(HSR) at Lewis and GE Aircraft

Engines.

In fiscal year 1999, the NCP team

will complete the major pieces

of the cycle system, including

customer deck generation and

dynamic solver capability, and

will finish off the complete set of

engine components. At the same

time, NCP’s development will shift

to a major emphasis on zooming.

As part of NPSS, the NCP is supported under the NASA High

Performance Computing and

Communications Program.

National Cycle Program makes it possible to zoom to a more detailed analysis.

The NASA Lewis Research Center is developing an environment for analyzing and designing aircraft engines—the Numerical Propulsion System

Simulation (NPSS). NPSS will integrate multiple disciplines, such as aerodynamics, structure, and heat transfer, and will make use of numerical

“zooming” on component codes. Zooming is the coupling of analyses at

various levels of detail. NPSS uses the latest computing and communication technologies to capture complex physical processes in a timely, costeffective manner. The vision of NPSS is to create a “numerical test cell”

enabling full engine simulations overnight on cost-effective computing

platforms.

Through the NASA/Industry Cooperative Effort agreement, NASA Lewis and

industry partners are developing a new engine simulation called the

National Cycle Program (NCP). NCP, which is the first step toward NPSS

and is its initial framework, supports the aerothermodynamic system

simulation process for the full life cycle of an engine. U.S. aircraft and

airframe companies recognize NCP as the future industry standard common analysis tool for aeropropulsion system modeling. The estimated

potential payoff for NCP is a $50 million/yr savings to industry through

improved engineering productivity.

NCP is being developed to provide all the features that currently exist

in cycle simulations as well as the framework needed to enable NPSS

concepts of coupling multidisciplinary tools at various levels of detail in a

collaborative distributed environment while preserving an overall engine

system view. What NCP now allows is illustrated by the following illustration.

Aeronautics

National

Cycle

Program

High-fidelity 

analysis

Axisymmetric engine

High-fidelity 

analysis Three-dimensional

low-pressure 

subsystem Three-dimensional

high-pressure CD-98-77860 compressor

Flowpath

analysis

Mixer 

nozzle Fan/booster

Fan/nacelle
10

Bibliography

Evans, A.L.; Follen, G.; and Naiman, C.: Numerical Propulsion System Simulation’s

National Cycle Program. AIAA Paper 98–3113, 1998.

Ashleman, R.H.; Lavelle, T.; and Parsons, F.: The National Cycle Program: A Flexible

System Modeling Architecture for Aircraft Engine Simulations. AIAA Paper 98–3114,

1998.

Lewis contacts:

Gregory J. Follen, (216) 433–5193, Gregory.J.Follen@grc.nasa.gov; and

Cynthia G. Naiman, (216) 433–5238, Cynthia.G.Naiman@grc.nasa.gov

Authors: Gregory J. Follen and Cynthia G. Naiman

Headquarters program office: OAT

Programs/Projects: HPCCP, NCP

Flow of GE90 Turbofan Engine Simulated

The primary flow path of the GE90

turbofan consists of a nacelle and

inlet, 49 blade rows of turbomachinery, and an exhaust nozzle.

Secondary flows entering and

exiting the primary flow path—

such as bleed, purge, and cooling

flows—were modeled macroscopically as source terms to accurately

simulate the engine. The information on these source terms came

from detailed descriptions of the

cooling flow and from thermodynamic cycle system simulations.

These provided boundary condition

data to the three-dimensional

analysis. A simplified combustor

was used to feed boundary conditions to the turbomachinery.

Flow simulations of the fan, highpressure compressor, and highand low-pressure turbines were

completed with the APNASA code.

The objective of this task was to create and validate a three-dimensional

model of the GE90 turbofan engine (General Electric) using the APNASA

(average passage) flow code. This was a joint effort between GE Aircraft

Engines and the NASA Lewis Research Center. The goal was to perform an

aerodynamic analysis of the engine primary flow path, in under 24 hours

of CPU time, on a parallel distributed workstation system. Enhancements

were made to the APNASA Navier-Stokes code to make it faster and more

robust and to allow for the analysis of more arbitrary geometry. The resulting simulation exploited the use of parallel computations by using two

levels of parallelism, with extremely high efficiency.

GE90 high-pressure compressor surface pressures. The figures

are shown in color in the online version of this article (http://

www.grc.nasa.gov/WWW/RT1998/2000/2900veres.htm).
11

Two high-pressure compressors were modeled, the original and a new,

improved design. The APNASA flow simulation of the new geometry

predicted a significant performance improvement over the baseline design.

APNASA has been further validated by these rig tests. The 128-processor

SGI Origin at the NASA Ames Research Center was used to simulate the

compression system, which comprises 31 blade rows, in 38 hours of elapsed time. The computer simulation time to convergence on the 18-bladerow, cooled, high- and low-pressure turbine simulation was 15 hours.

GE90 high-pressure cooled turbine flow simulation.

Full GE90 turbofan engine flow simulation (with simple combustor).

Because of its higher efficiency and reduced specific fuel consumption, the

improved high-pressure compressor will significantly reduce the operating

cost of the GE90 engine. The accurate predictive simulation capability of

APNASA will increase design confidence in future versions of the baseline

GE90 engine featuring higher levels of thrust. Increased design confidence

will result in fewer test rigs and lower costs during development and

certification.

All the major engine components

have been simulated with APNASA

for 49 blade rows (see the final

figure). This high-fidelity simulation

will be used by GE to evaluate

modifications for future versions

of the GE90 to reduce derivative

engine development time and cost.

This work, which is a major element

of the Numerical Propulsion System

Simulation (NPSS), supports the

High Performance Computing and

Communication Program’s (HPCCP)

Grand Challenge milestone to

“Demonstrate end-to-end reductions in cost and time to solution

for aerospace propulsion design

applications on heterogenous

computing systems.”

More information about this research

is available on the World Wide Web:

http://hpcc/ge90/

Lewis contact:

Joseph P. Veres, (216) 433–2436,

Joseph.P.Veres@grc.nasa.gov

Author: Joseph P. Veres

Headquarters program office: OAT

Programs/Projects: HPCCP, NPSS

Aeronautics
13

Research and Technology
14

Materials

MoSi2-Base Structural Composite Passed Engine Test

The intermetallic compound molybdenum disilicide (MoSi2) is an attractive

high-temperature structural material for advanced engine applications. It

has excellent oxidation resistance, a high melting point, relatively low

density, and high thermal conductivity; and it is easily machined. Past

research at the NASA Lewis Research Center has resulted in the development of a hybrid composite consisting of a MoSi2 matrix reinforced with

silicon nitride (Si3N4) particulate and silicon carbide (SiC) fibers. This

composite has demonstrated attractive strength, toughness, thermal

fatigue, and oxidation resistance, including resistance to “pest” oxidation.

These properties attracted the interest of the Office of Naval Research and

Pratt & Whitney, and a joint NASA/Navy/Pratt & Whitney effort was developed to continue to mature the MoSi2 composite technology. A turbine

blade outer air seal, which was part of the Integrated High Performance

Turbine Engine Technology (IHPTET) program, was chosen as a first component on which to focus.

The first tasks of the materials development effort were to develop improved processing methods to reduce costs and to use fine-diameter fibers

that enable the manufacturing of complex shapes. Tape-casting methods

were developed to fully infiltrate the fine SiC fibers with matrix powders.

The resulting composites were hot pressed to 100-percent density. Composites with cross-plied fiber architectures with 30 vol % Hi-Nicalon SiC

fibers (Nippon Carbon Company, Japan) and 30 vol % nitride particles

are now made routinely and demonstrate a good balance of properties.

Time, msec

0.0 0.2 0.4 0.6 0.8 1.0

Force, N

0

100

10–1

10–2

101

102

5

4

3

2

1

0

6x103

–300 300 600 900 1200 1500

Temperature, °C

CVN Energy, J

SiC

AS–800 (b Si3N4)

Hot pressed Si3N4

MoSi2–50Si3N4

NiAl

Hastelloy X (Wrought superalloy)

B 1900 

(Cast superalloy)

[0°/90]

SCS–6[0°]/MoSi2–50Si3N4

The next task entailed the measurement of a wide variety of

mechanical properties to confirm

the suitability of this composite in

engines. In particular, participants

in this effort demonstrated that

composites made with Hi-Nicalon

fibers had strength and toughness

properties equal to or better than

those of the composites made with

the large-diameter fibers that had

been used previously. Another

critically important property meassured was impact resistance. Aircraft engine components require

sufficient toughness to resist manufacturing defects, assembly damage, stress concentrations at

notches, and foreign object damage. Engine company designers

indicated that impact resistance

would have to be measured before

they would seriously consider

these types of composites. The

Charpy impact properties of MoSi2 base composites compared with several competitive materials. Left: Force-time

curves. Right: Charpy-V-Notch (CVN) energy plot.
15

Research and Technology

Charpy V-Notch test was chosen to assess impact resistance, and both

monolithic and composite versions of MoSi2 were tested from –300

to 1400 °C. The results (see the preceding graphs) show that nitrideparticulate-reinforced MoSi2 exhibited impact resistance higher than that of

many monolithic ceramics and intermetallics and that the fiber-reinforced

composites had even higher values, approaching that of cast superalloys.

These and other results led to the decision to test the hybrid composite

in the aggressive environment of a gas turbine engine. Test coupons were

fabricated and tested in Pratt & Whitney’s IHPTET XTC/66 demonstrator

engine to simulate the thermal cycling conditions of a blade outer air seal.

The composite was exposed to thermal cycles to 1200 °C, with a 600 °C

gradient through the thickness of the coupon. It survived all 15 engine

cycles without any appearance of distress (see the photomicrograph).

These results indicate that MoSi2-base composites are still competitive with

state-of-the-art ceramic composites as replacements for superalloys in jet

engines. Further work on processing, fiber coatings, and mechanical and

environmental durability is in progress.

Bibliography

Hebsur, M.G; and Nathal, M.V.: Strong, Tough, and Pest Resistant MoSi2-Base Hybrid

Composite for Structural Applications. Structural Intermetallics 1997, M.V. Nathal,

et al., eds., 1997, The Minerals, Metals and Materials Society, Warrandale, PA,

pp. 949–958.

Lewis contacts:

Dr. Mohan G. Hebsur, (216) 433–3266, Mohan.G.Hebsur@grc.nasa.gov; and

Dr. Michael V. Nathal, (216) 433–9516, Michael.V.Nathal@grc.nasa.gov

C-98-977

Hybrid composite coupon of MoSi2/

Si3N4 particulate/Hi-Nicalon SiC fiber

after engine testing at Pratt & Whitney.

Office of Naval Research (ONR)

contact:

Dr. A.K. Vasudevan, (703) 696–8181

Authors: Dr. Michael V. Nathal and

Dr. Mohan G. Hebsur

Headquarters program office: OAT

Programs/Projects: P&PM

Nickel- (Ni-) base superalloy single crystals represent the state-of-the-art

for turbine engine airfoil applications because they offer the best balance

of properties under the high operating temperatures required for efficient engine operation. Current trends in alloy design take advantage of

improved creep rupture strength with the addition of higher levels of

refractory elements. In particular, the addition of significantly higher levels

of rhenium in third-generation superalloys is key for both microstructural

stability and creep rupture strength. Although refractories provide strength

benefits, alloys tend to be unstable when their refractory content is high

because of topologically close-packed (TCP) phases. The formation of these

phases in sufficient amount is detrimental to the performance of these

alloys because of their brittle nature and because they deplete the Ni-rich

matrix of potent solid-solution strengthening elements.

TCP phase formation in modern single-crystal superalloys is being investigated in-house at the NASA Lewis Research Center, where particular

attention is being paid to the potential synergistic effects of alloying elements. The approach chosen in this investigation was to develop a

model based on a design-of-experiments methodology. A design of

Deleterious Phase Formation in Next-Generation

Nickel-Base Superalloys Predicted

experiments consisting of 44 alloys

was set up to quantify both the

linear and pairwise interactive

effects of aluminum (Al), cobalt

(Co), chromium (Cr), molybdenum

(Mo), rhenium (Re), tantalum

(Ta), and tungsten (W) in a thirdgeneration Ni-base superalloy

according to the resultant amount

of TCP phase in the microstructure.

All alloys were produced by vacuum induction melting and were

cast using high-purity melting

stock. This was followed by a

homogenization heat treatment

and a simulated long–duration

engine exposure.
16

10 mm 1 mm

After alloys were aged at 1093 °C for 400 hours, an examination of all

microstructures showed that the alloys fell into three groups on the basis

of the TCP distribution. Group 1 alloys contained no TCP phase. Group 2

contained less than 3 vol % TCP, which formed exclusively at the grain

boundaries. Group 3 contained between 3 and 17 vol % TCP, which

formed within grains as well as at grain boundaries.

A regression model was developed to describe the presence of TCP phase

in the microstructures of a polycrystalline third-generation superalloy on

the basis of chemical content. The result of this analysis, in terms of atomic

percent, follows:

(vol % TCP)1/2 = 16.344782 – 1.019587(Al) – 2.624322(Cr)

–3.821997(Mo) + 1.109575(Re) – 3.207295(Ta)

+ 6.462984(W) – 2.271803(Co) + 0.052884(Al)(Co)

+ 0.214059(Al)(Cr) + 0.300698(Al)(Mo)

+ 0.80011(Co)(Re) + 0.257108(Cr)(Mo)

– 5.081598(Re)(W) + 1.824441(Ta)(W)

This relationship has been shown to explain 95 percent of the TCP variations seen in the experiments. It is also useful for identifying principal

Third-generation, Ni-base, superalloy aged microstructures. Left: Group 2. Right: Group 3.

elemental effects as well as interactive effects toward TCP phase

instability and for providing insight

for stable alloy development.

Additional efforts are underway

to further refine these results.

Bibliography

Ritzert, F., et al.: The Effect of Alloying

on Topologically Close Packed Phase

Instability in Advanced Nickel-Based

Superalloy Rene N6. NASA TM—1998-

206622, 1998.

Lewis contact:

Frank J. Ritzert, (216) 433–8199,

Frank.J.Ritzert@grc.nasa.gov

Author: Frank J. Ritzert

Headquarters program office: OAT

Programs/Projects: HITEMP

Efforts to improve the performance of modern gas turbine engines have

imposed increasing service temperature demands on structural materials.

Through active cooling, the useful temperature range of nickel-base

superalloys in current gas turbine engines has been extended, but the

margin for further improvement appears modest. Because of their low

density, high-temperature strength, and high thermal conductivity, in situ

toughened silicon nitride ceramics have received a great deal of attention

for cooled structures. However, high processing costs have proven to be a

Affordable Manufacturing Technologies Being Developed for

Actively Cooled Ceramic Components

major obstacle to their widespread

application. Advanced rapid prototyping technology, which is developing rapidly, offers the possibility

of an affordable manufacturing

approach.
17

Research and Technology

Silicon nitride (Si3N4) plate with cooling holes.

~0.1 cm

(~0.040 in.) ~0.02 cm

(~0.008 in.) (~0.008 in.)

~0.05 cm

(~0.02 in.) (~0.02 in.)

Researchers at the NASA Lewis Research Center, in cooperation with a

local university and industry, are developing actively cooled and functionally graded ceramic structures. The objective of this program is to develop

a cost-effective manufacturing technology and the analytical modeling

capability to predict thermomechanical stresses in burner rig tests of

thermal-barrier-coated, actively cooled, in situ toughened silicon nitride

turbine nozzle vanes under simulated engine conditions. In the initial

studies, green and sintered bodies of silicon nitride plates with cooling

holes were successfully fabricated through the use of a Sanders prototype

molding machine for mold detail fabrication, ceramic gel casting, and conventional sintering techniques. A two-dimensional analytical model was

developed to optimize the cooling

hole geometry and spacing, and to

reduce thermal stresses.

Lewis contact: Dr. Ramakrishna T.

Bhatt, (216) 433–5513,

Ramakrishna.T.Bhatt@grc.nasa.gov

Author: Dr. Ramakrishna T. Bhatt

Headquarters program office: OAT

Programs/Projects: HITEMP

There is strong interest in the development of silicon carbide-based monolithic ceramic and composite materials and components for demanding,

high-temperature applications. Thorough characterization of material properties, including high-temperature testing under simulated or actual operating conditions, is a high priority for programs involved in developing

these silicon carbide- (SiC) based materials and components.

Members of the Sensors and Electronics Technology Branch at the NASA

Lewis Research Center are developing minimally intrusive methods of

measuring the properties (such as the surface temperature, strain, and

heat flux characteristics) of components and subelements that are being

tested or operated in hostile, high-temperature environments. Their primary goal is to instrument the test article or operating component with

durable sensors that have a minimal effect on test conditions such as the

gas flow across the surface of the item and the material response (including the through-thickness conduction of heat). Therefore, the main thrust

of their work has been the development of thin-film sensors (e.g., thermocouples or strain gauges) for use on various advanced material test articles,

including SiC/SiC composite components.

Sensor Lead Wires Positioned on SiC-Based Monolithic

Ceramic and Fiber-Reinforced Ceramic Matrix Composite

Subcomponents With Flat and Curved Surfaces

There was a need for a better

method of securing sensor lead

wires on SiC-based components

and subelements that would be

tested at temperatures to 1000 °C

(or higher), to enhance the durability of the overall minimally intrusive

sensor system. To address this

need, Lewis researchers devised an

alternative approach for positioning

the sensor lead wires (which are

connected to the thin-film sensors)

on SiC or SiC/SiC components. A

reaction-forming method of joining

was used to strongly bond hoopshaped monolithic SiC and SiC/SiC

composite attachments of various sizes to both flat and curved

surfaces of SiC/SiC composite
18

subelements (see the photos). This approach is based on an affordable,

robust ceramic joining technology, named ARCJoinT, which was developed at Lewis for the joining of SiC-based ceramics and fiber-reinforced

composites.

The ability to join these attachments to a curved surface is important

because many SiC/SiC components that need to be tested have some

curved surfaces. The use of thin attachment hoops having a radius slightly

larger than the diameter of the sensor lead wire cable should (1) prevent

the wires from moving, which reduces the risk of sensor failure, (2) minimize the disruption of high-velocity combustion or cooling gas flow, and

(3) minimize the influence of the attachments on component performance

during testing under simulated engine conditions because the hoops have

minimal contact with the test article. On the basis of results obtained in

previous joining studies, the joined attachments should maintain their

integrity at temperatures up to 1350 °C in air. Preliminary testing under

static high-temperature conditions has yielded positive results.

The projected excellent high-temperature strength of the joined attachments should make it possible to obtain better sensor durability than has

previously been achieved by using refractory, adhesive cement to secure

lead wire assemblies. This novel approach is being evaluated further

through burner rig testing of instrumented panels to determine the potential for obtaining enhanced durability, minimally intrusive, high-temperature

sensor systems. The technology apparently could also be used to reinstall

attachments on ceramic components that have been damaged in service.

SiC hoops successfully attached (via reaction joining) to a

SiC/SiC composite subelement having a curved surface.

SiC hoops joined to SiC/SiC composite panels. These

attachments are used to position and secure sensor

lead wires.

CD-97-2383 CD-97-2305

For more information, visit Lewis on

the World Wide Web:

Ceramics Branch: http://

www.grc.nasa.gov/WWW/Ceramics/

Sensors and Electronics Technology

Branch: http://

www.grc.nasa.gov/WWW/sensors/

Lewis contacts:

James D. Kiser, (216) 433–3247,

James.D.Kiser@grc.nasa.gov;

Dr. Jih-Fen Lei, (216) 433–3922,

Jih-Fen.Lei@grc.nasa.gov; and

Lisa C. Martin, (216) 433–6468,

Lisa.C.Martin@grc.nasa.gov

Dynacs Engineering Co., Inc. contact:

Dr. Mrityunjay Singh, (216) 433–8883,

Mrityunjay.Singh@grc.nasa.gov

Authors:

James D. Kiser, Dr. Mrityunjay Singh,

Dr. Jih-Fen Lei, and Lisa C. Martin

Headquarters program office: OAT

Programs/Projects: HITEMP, EPM
19

Research and Technology

Ceramic matrix composites are being developed for elevated-temperature

engine applications. A leading material system in this class of materials is

silicon carbide (SiC) fiber-reinforced SiC matrix composites. Unfortunately,

the nonoxide fibers, matrix, and interphase (boron nitride in this system)

can react with oxygen or water vapor in the atmosphere, leading to

strength degradation of the composite at elevated temperatures. For this

study, constant-load stress-rupture tests were performed in air at temperatures ranging from 815 to 960 °C until failure. From these data, predictions

can be made for the useful life of such composites under similar stressedoxidation conditions.

During these experiments, the sounds of failure events (matrix cracking

and fiber breaking) were monitored with a modal acoustic emission (AE)

analyzer through transducers that were attached at the ends of the tensile

bars. Such failure events, which are caused by applied stress and oxidation reactions, cause these composites to fail prematurely. Because of

the nature of acoustic waveform propagation in thin tensile bars, the location of individual source events and the eventual failure event could be

detected accurately.

The graph shows the load versus time and the AE event location versus

time for an experiment performed at 960 °C. The AE events are plotted

according to their respective energies; that is, the highest energy events are

the loudest events and correspond to large matrix cracks. Also noted on

this figure is the location of the hot zone, which is a 15-mm region of the

furnace that is at 960 °C. Outside of this region the temperature decreases.

It is evident that after a few thousand seconds very little activity occurs in

the hot zone. However, just outside the hot zone, a significant amount

of AE activity occurs, with eventual composite failure occurring at about

880 °C. For experiments performed at 815 °C, the failure location is always

Modal Acoustic Emission Used at Elevated Temperatures

to Detect Damage and Failure Location in Ceramic

Matrix Composites

Acoustic emission (AE) event location and load versus time for stress-rupture experiment performed in air at 960 °C.

in the hot zone. Because at ~800

to 900 °C chemical reactions

between the interphase and environment reduce the load-carrying

ability of the reinforcing fibers, this

constitutes the “pest” temperature

regime for this material.

Modal AE accurately located the

failure location of the composite.

In addition, it detected a number

of loud AE events in the same

relative failure location prior to

failure. Since eventual failure of

these composites occurs at a specific location after a progression

of smaller matrix cracks and fiber

failures which weaken that area of

the composite, this type of monitoring can be used to actually

predict the location of composite

failure.

Lewis contact:

Gregory N. Morscher, (216) 433–5512

Gregory.N.Morscher@grc.nasa.gov

Author: Gregory N. Morscher

Headquarters program office: OAT

Programs/Projects: HITEMP

0

1

2

3

4

5

0 2 4 6 8 10 12x103

AE event location, mm

Hot zone

Time, sec

Load, kN

–20

0

20

40

Failure 

location

AE energy

Load

Lowest

2nd highest

Highest
20

High-Temperature Solid Lubricants Developed by

NASA Lewis Offer Virtually “Unlimited Life” for

Oil-Free Turbomachinery

PS304-coated bearing after 100 000 cycles at 1000 °F. This oil-free turbocharger features two journal foil bearings, two thrust

foil bearings, a NASA PS304 coating, and a rigid rotor. All dimensions are in inches.

Journal

foil bearings

Thrust

foil 

bearings

Compressor inlet 

(100 °F)

Compressor 

exhaust (500 °F)

Turbine 

exhaust

(700 to 900 °F)

Turbine inlet

(1100 to 1400 °F)

O 5.500 /

r j c

 5.250

+_0.003

O 5.750 /

2.890 6.671 3.441

3.600
21

Research and Technology

The NASA Lewis Research Center is capitalizing on breakthroughs in foil

air bearing performance, tribological coatings, and computer analyses to

formulate the Oil-free Turbomachinery Program. The program’s long-term

goal is to develop an innovative, yet practical, oil-free aeropropulsion gas

turbine engine that floats on advanced air bearings. This type of engine

would operate at higher speeds and temperatures with lower weight and

friction than conventional oil-lubricated engines. During startup and

shutdown, solid lubricant coatings are required to prevent wear in such

engines before the self-generating air-lubrication film develops.

NASA’s Tribology Branch has created PS304, a chrome-oxide-based plasma

spray coating specifically tailored for shafts run against foil bearings. PS304

contains silver and barium fluoride/calcium fluoride eutectic (BaF2/CaF2)

lubricant additives that, together, provide lubrication from cold start temperatures to over 650 °C, the maximum use temperature for foil bearings.

Recent lab tests show that bearings lubricated with PS304 survive over

100 000 start-stop cycles without experiencing any degradation in performance due to wear. The accompanying photograph shows a test bearing

after it was run at 650 °C. The rubbing process created a “polished” surface

that enhances bearing load capacity.

This type of bearing performance suggests virtually unlimited life for

foil-bearing-supported, oil-free turbomachinery. By adapting these

technologies, future aeropropulsion

engines can run more efficiently

and be maintenance free as well

as oil free.

Lewis contacts: Dr. Christopher

DellaCorte, (216) 433–6056,

Christopher.DellaCorte@grc.nasa.gov;

and Dr. Mark J. Valco, (216) 433–3717,

Mark.J.Valco@grc.nasa.gov

Authors: Dr. Christopher DellaCorte

and Dr. Mark J. Valco

Headquarters program office: OAT

Programs/Projects: P&PM, SGE

Special recognition:

A patent will be issued on this coating

technology in January 1999.

To meet the increased thermal stresses of future advanced aircraft engines,

new lubricants will have to be developed to replace the currently used

ester-based liquid lubricants. If a suitable conventional replacement cannot

be found, a different lubrication method will have to be used. The conventional method circulates bulk lubricant (stored in a sump) through a lubricating system containing cooling and filtering elements. Solid lubricants

have been studied as a replacement for bulk liquid lubricants, and have

been found to provide reasonable lubrication for lightly loaded systems.

Solid lubricants, however, have proved inadequate for highly loaded, highspeed applications. Vapor/mist phase lubrication (VMPL), on the other

hand, may be a viable alternative.

VMPL has been used successfully to lubricate high-temperature bearings or

gears. It can be used as an emergency backup system or as the primary

source of lubrication. With VMPL, minimal weight is added to the system

and minimal debris is formed. It works over a wide temperature range. The

concept is illustrated in the following figure.

The basis for the VMPL method stems from the idea that a “suitable”

organic vapor or fine mist can react inside high-temperature sliding surfaces to form a lubricious film. The primary compounds used for VMPL

studies have been organophosphates, which work well with only certain

ferrous material. Their success is due to an initial, rapid formation of a

predominant iron phosphate film, which is followed by the formation and

growth (by cationic diffusion) of a lubricious pyrophosphate-type coating

over the iron phosphate. As long as iron is present at a wearing surface,

New Vapor/Mist Phase Lubricant Formulated

Vapor/mist phase lubrication method.

Airstream

Vapor 

stream

Organic 

liquid

Bearings

and gears

(1) A small quantity of an organic 

 liquid is vaporized into the air-

 stream (liquid in air, <1 wt%).

(2) The vapor steam is delivered 

 to the bearings and gears.

(3) A lubricating polymeric film 

 is deposited at the wearing 

 surfaces.

1

2

3

1

3

2
22

the VMPL method using organophosphates works well. Evidence, however, has been reported that continued organophosphate interaction with

ferrous bearing material can lead to depletion of surface iron and to

eventual lubrication failure. If VMPL is to work for prolonged periods of

time, the depletion of surface iron must be circumvented.

Work at the NASA Lewis Research Center has resulted in a new organophosphate formulation that prevents surface iron depletion. This new

formulation incorporates a soluble iron additive that reacts with the organophosphate to generate an iron phosphate film. This formulation was used

to successfully vapor/mist phase lubricate a spur gearbox in a preliminary

study. In a related study, this formulation was used to lubricate a silicon

nitride pin-on-plate couple that used a reciprocating friction and wear

apparatus at 500 °C.

Lewis contact:

Dr. Wilfredo Morales, (216) 433–6052,

Wilfredo.Morales@grc.nasa.gov

Authors:

Dr. Wilfredo Morales and

Dr. Robert F. Handschuh

Headquarters program office: OAT

Programs/Projects: P&PM

Special recognition: A special

achievement award was given for

this work, and a patent was issued.

Large-scale simulations of dynamic processes at the atomic level have

developed into one of the main areas of work in computational materials

science. Until recently, severe computational restrictions, as well as the

lack of accurate methods for calculating the energetics, resulted in slower

growth in the area than that required by current alloy design programs.

The Computational Materials Group at the NASA Lewis Research Center is

devoted to the development of powerful, accurate, economical tools to

aid in alloy design. These include the BFS (Bozzolo, Ferrante, and Smith)

method for alloys (ref. 1) and the development of dedicated software for

large-scale simulations based on Monte Carlo-Metropolis numerical techniques, as well as state-of-the-art visualization methods. Our previous

effort linking theoretical and computational modeling resulted in the

successful prediction of the microstructure of a five-element intermetallic

alloy, in excellent agreement with experimental results (refs. 2 and 3). This

effort also produced a complete description of the role of alloying additions in intermetallic binary, ternary, and higher order alloys (ref. 4).

A natural continuation of this work includes dynamic processes. Preliminary results from Monte Carlo-Metropolis minimization techniques used

in conjunction with the BFS method for calculating the energetics, show

a promising and viable way to gain insight into the microscopic evolution

of intermetallic fracture, as shown in the illustration on the facing page.

In this example, a nickel-aluminide- (NiAl) ordered alloy in the B2 phase

is uniformly stretched until the computational cell develops cracks, which

ultimately lead to the sample’s fracture. Beyond this point, the surfaces

created during this process reconstruct, showing a strong preference for

Monte Carlo Simulation of Alloy Design Techniques:

Fracture and Welding Studied Using the BFS Method

for Alloys

mixed-composition termination, as

is to be expected. This Monte Carlo

approach allows for individual

atomic relaxations (which were

exaggerated in the illustration for

visualization purposes).

The final illustration shows a

temperature-dependent process

where the computational cell is

divided in two halves with an

intermediate copper (Cu) layer, in

what amounts to the first simulation of the brazing process. In this

example, the temperature is raised

steadily in small increments that

allow the sample to stabilize at

each step. The Monte CarloMetropolis approach used in this

example allows only for atomic

exchanges between nearest

neighbors, thus showing the

increasing interdiffusion of Cu in

the NiAl matrix as the temperature

increases beyond the melting

temperature of Cu.
23

Research and Technology

Onset of fracture for NiAl for various states of the computational cell. Ni and Al atoms are denoted with lighter and darker

spheres, respectively. (This figure is shown in color in the online version of this article: http://www.grc.nasa.gov/WWW/

RT1998/5000/5140noebe.htm.) (a) Beginning of the simulation. (b) After uniaxial strain, just before the onset of fracture.

(c) Just after the onset of fracture. (d) During relaxation of the newly formed fracture surfaces. (e) After fracture relaxation

is essentially complete.

Ni

Al

Although Monte Carlo numerical techniques allow for a detailed visualization of the simulated process, the main advantage of this type of analysis

is that it helps researchers to understand the underlying features that drive

these processes. Because of the simplicity of the BFS method for the energetics used in these calculations, a detailed atom-by-atom analysis can be

performed at any point in the simulation, providing necessary insight into

the details of the process.

A main objective of this research program is to make the calculations as

simple as possible. By reducing computational effort without losing physical accuracy, we expect that powerful simulation tools will be developed in

the immediate future which will

allow material scientists to easily

visualize and analyze processes at a

level not achievable experimentally.

References

1. Bozzolo, G.; Ferrante, J.; and Kobistek,

R.: Modelling of Surfaces. J. Comput.-

Aided Mater. Des., vol. 1, no. 3, 1994,

pp. 305–324.

(a) (b) (c) (d) (e)
24

2. Bozzolo, G., et al.: Progress in the

Modelling of NiAl-Based Alloys

Using the BFS Method. Structural

Intermetallics 1997, M.V. Nathal,

et al., eds., The Minerals, Metals

and Materials Society, Warrendale,

PA, 1997, pp. 655–664.

3. Noebe, R.D.: Research & Technology

1996. BFS Method for Alloys Optimized and Verified for the Study of

Ordered Intermetallic Material.

NASA TM–107350, 1997, pp. 49–50.

(Available online: http:/www.grc.

nasa.gov/WWW/RT1996/5000/

5120b.htm)

4. Bozzolo, G.; Noebe, R.D.; and

Honecy, F.: Modeling of Substitutional Site Preference in Ordered

Intermetallic Alloys. The Minerals,

Metals and Materials Society,

Warrendale, PA, 1998,

pp. 341–368.

Lewis contact:

Ronald D. Noebe, (216) 433–2093,

Ronald.D.Noebe@grc.nasa.gov

Ohio Aerospace Institute contact:

Guillermo Bozzolo, Ohio Aerospace

Institute, (440) 962–3103,

Guillermo.H.Bozzolo@grc.nasa.gov

Authors:

Guillermo Bozzolo, Brian Good,

Ronald D. Noebe, Frank Honecy, and

Phillip Abel

Headquarters program office: OAT

Programs/Projects: HITEMP, P&PM

NiAl + Cu computational cell. Ni, Al, and Cu atoms are denoted with medium, dark,

and light spheres, respectively. (This figure is shown in color in the online version

of this article: http://www.grc.nasa.gov/WWW/RT1998/5000/5140noebe.htm.)

(a) Original cell, with four atomic layers of Cu between two identical NiAl cells.

(b) After stabilization at 1400 °F.

Ni

Al

Cu

(a) (b)
25

Research and Technology

PMR–15 polyimide, developed in the mid-1970’s at the NASA Lewis

Research Center, is recognized as a state-of-the-art high-temperature resin

for composite applications in the temperature range of 500 to 550 °F (260

to 288 °C). PMR–15 offers easy processing and good property retention at

a reasonable cost. For these reasons, it is widely used in both military and

commercial aircraft engine components. Traditionally, polyimide composites have been designed for long-term use at 500 to 600 °F over thousands of hours. However, new applications in reusable launch vehicles

(RLV’s) require lightweight materials that can perform for short times (tens

of hours) at temperatures between 800 and 1000 °F (425 and 538 °C).

Current efforts at Lewis are focused on raising the use temperature of

polyimide composites by increasing the glass-transition temperature of

the matrix resins. Achieving this dramatic increase in the upper use temperature without sacrificing polymer and composite processability is a major

technical challenge.

A recent development from these efforts is a low-cost, high glass-transition

temperature, thermosetting polyimide (DMBZ–15), prepared from 3,3’,4,4’-

benzophenonetetracarboxylic acid dimethyl ester (BTDE) and 2,2’-

dimethylbenzidine (DMBZ) with nadic ester as the endcap. The glasstransition temperature of a DMBZ–15 polyimide/T650–35 carbon fiber

composite (414 °C) is much higher than that of the corresponding

PMR–15 composite (348 °C). DMBZ–15 polyimide-based composites

exhibit good compressive strengths at room and elevated temperatures

(233 and 288 °C), as well as good hot/wet compressive strength. (See the

Low-Cost, High Glass-Transition Temperature, Thermosetting

Polyimide Developed

Compressive strength of polyimide/T650–35 composities. Flexure strength of DMBZ–15 and PMR–15 composites. Compressive strength, ksi

120

100

60

40

80

20

0

23

(74)

232

(450)

288

(500)

Test temperature, °C (°F)

5 hot/wet

cycles

PMR–15

DMBZ–15

Relative flexural strength

1

0

Room temperature 550 1000

Test temperature, °F

PMR–15

DMBZ–15

left figure, where one hot/wet

cycle consists of a 93 °C (200 °F)

water soak to a gain of >1 wt %,

followed by a dry out at 288 °C

(500 °F) to <0.1 wt % moisture.)

In addition, flexural strengths of

PMR–15 and DMBZ–15 composites

measured at room temperature,

288 °C, and 538 °C are comparable (see the right figure). These

materials, along with other highperformance fiber-reinforced

polymer matrix composites are

currently being evaluated in Lewis/

industry/university/Government

cooperative efforts for potential use

in reusable launch vehicles (RLV’s)

and other space vehicles.

Lewis contact:

Dr. Kathy C. Chuang, (216) 433–3227,

Kathy.Chuang@grc.nasa.gov

Author: Dr. Kathy C. Chuang

Headquarters program office: OAT

Programs/Projects: HITEMP, RLV, AST
26

CH3 O

OH

Ph

N

O

Ph

O

N

HO Ph O

O

Ph

1 2

3

4

Polyimides have found broad application in fiber-reinforced composites for

aerospace components and as thin films for electronics packaging. Typical

routes to processing these materials require temperatures above 200 °C.

As a result, tooling costs for fabricating components from these composites

can be quite high. Recent efforts within the Polymers Branch at the NASA

Lewis Research Center have been aimed at developing radiation-curable

(with light or electron beams) polyimides. Such materials may enable the

processing of polymers and composites at or near room temperature,

leading to reduced tooling requirements and costs.

Novel Ultraviolet-Light-Curable Polyimides

Generation and trapping of photoenol, 2, from o-methylbenzophenone, 1.

A new Diels-Alder route to polyimides has been developed that employs

ultraviolet light (UV), rather than heat, to effect polymerization. This

approach, which can be carried out at room temperature, is based on a

well-known photochemical reaction—the photoenolization of o-methylphenyl ketones. Irradiation of o-methylphenyl ketones, such as 1 in

the preceding figure, with UV wavelengths above 300 nm produces a

photoenol, 2. This photoenol is unstable, but it persists long enough to

undergo Diels-Alder reactions with good dienophiles, such as maleimide,

3. By utilizing a diketone, such as 2,5-dibenzoyl-p-xylene, 5, and a

bismalemide, 6, this chemistry has been used to make a number of

polyimides, 7 (see the figure on the following page).

These polyimides have glass-transition temperatures as high as 300 °C and

have modest stabilities in both air and nitrogen. Onsets of decomposition,

measured by thermal gravimetric analysis, were as high as 400 °C in air

and 450 °C in nitrogen. Higher glass-transition temperatures and onsets of

decomposition can be obtained by conversion of polyimide 7 into polyimide 8 through acid-catalyzed dehydration followed by dehydrogenation

(as shown in the second figure). These polyimides have glass-transition

temperatures as high as 330 °C and onsets of decomposition as high as

550 °C in air and 525 °C in nitrogen.

This chemistry has been demonstrated in solution (benzene or cyclohexanone). However, it should be easily adapted to achieve solid-state

(solvent-free) curing, making it well-suited for thin-film applications

(coatings, electronics packaging, and photonic and optical materials).

This new UV curing process could

offer several advantages over other

methods of preparing polyimides.

UV-cured films should undergo less

shrinkage during cure than thermally cured films. UV curing would

also be useful for processing

polyimides containing thermally

sensitive groups or additives, such

as photonic materials. Current

efforts are aimed at modifying

this chemistry to enable electronbeam curing of polyimides and

polyimide-based composites for

aeropropulsion and reusable

launch vehicle applications.

Lewis contact: Dr. Michael A. Meador,

(216) 433–9518,

Michael.A.Meador@grc.nasa.gov

Author: Dr. Michael A. Meador

Headquarters program office: OAT

Programs/Projects: P&PM, RLV
27

Research and Technology

CH3

Ar'

O

CH3

Ar'

O

Ar N

O

O

N

O

O

N

O

O

Ph hν

(Pyrex)

C6H6, N2

Ph N N

O

O

O

Ar'

Ar' OH O

Ar' OH

N N

O

O Ar' OH

O

O

Ar' OH

Ph

+ +

n

5 6 3

7

Ar = 

, 

O

, ,

O O

CN

,

5

OH

Ar'

HO

Ar'

Ph OH

Ph OH

N N

O

O

O

O

Ar

hν

CH3

CH3

Ar'

O

Ar'

O

6

Ar N

O

O

N

O

O

n

a b c

d e

Ar' = Ph, p-(C 12H25O)Ph

Ar'

Ph N N

O

O

O

Ar

Ar' O

Ar'

N N

O

O

O

O

Ph

Ar'

n

8

HCl/Ac2O

Reflux

N2, 3h

Sulfur/Ph2O

Reflux

N2, 6h

Polyimides from Diels-Alder trapping of photochemically generated bisdienes.
28

Power systems with operating temperatures in the range of 815 to 982 °C

(1500 to 1800 °F) frequently require alloys that can operate for long times

at these temperatures. A critical requirement is that these alloys have adequate oxidation resistance. The alloys used in these power systems require

thousands of hours of operating life with intermittent shutdown to room

temperature. Intermittent power plant shutdowns, however, offer the

possibility that the protective scale will tend to spall (i.e., crack and flake

off) upon cooling, increasing the rate of oxidative attack in subsequent

heating cycles. Thus, it is critical that candidate alloys be evaluated for

cyclic oxidation behavior. It was determined that exposing test alloys to

ten 1000-hr cycles in static air at 982 °C (1800 °F) could give a reasonable

simulation of long-time power plant operation.

Iron- (Fe-), nickel- (Ni-), and cobalt- (Co-) based high-temperature alloys

with sufficient chromium (Cr) and/or aluminum (Al) content can exhibit

excellent oxidation resistance. The protective oxides formed by these

classes of alloys are typically Cr2O3 and/or Al2O3, and are usually influenced by their Cr, or Cr and Al, content. Sixty-eight Co-, Fe-, and Ni-base

high-temperature alloys, typical of those used at this temperature or higher,

were used in this study. (Detailed chemical compositions of these alloys

are listed in ref. 1.)

At the NASA Lewis Research Center, the alloys were tested and compared

on the basis of their weight change as a function of time, x-ray diffraction

of the protective scale composition, and the physical appearance of the

exposed samples. Although final appearance and x-ray diffraction of the

final scale products were two factors used to evaluate the oxidation

resistance of each alloy, the main criterion was the oxidation kinetics

inferred from the specific weight change versus time data. These data

indicated a range of oxidation behavior including parabolic (typical of

isothermal oxidation), paralinear, linear, and mixed-linear kinetics. Paralinear kinetics was the most typical behavior, with scale growth at the

operating temperature and scale spalling as the sample cooled between

exposure cycles. Of the 132 cyclic oxidation tests (including replicates), 94

indicated paralinear behavior, only 4 showed parabolic behavior (scale

growth only), and 34 were linear or mixed linear, where spalling tended to

be massive.

The gravimetric data were fit to the basic paralinear equation:

∆W/A = k1

1/2t

1/2 + k2

t ± SEE

where ∆W is the change of sample weight (in milligrams) with time, A is

the surface area in square centimeters, t is time in hours, SEE is the stan10 000-hr Cyclic Oxidation Behavior of 68 HighTemperature Co-, Fe-, and Ni-Base Alloys

Evaluated at 982 °C (1800 °F)

dard error of estimate by the

multiple linear regression method,

and depending upon the degree of

fit, the significance and sign of the

constants k1

1/2 and k2 define the

kinetic model. (The use of this

equation is discussed in detail in

ref. 1.) The regression coefficients

were combined into a single

parameter, which along with

the alloys’ physical appearance,

provided a relative ranking of the

alloys’ oxidation resistance.

Of the 68 alloys tested, 16 alloys

were ranked as excellent, 18 good,

5 fair, 10 poor, and 19 catastrophic.

The top bar graph shows 23 Crcontaining Ni-base alloys that

formed a protective chromia/

chromite scale. The bottom bar

graph shows the best of the

excellent-ranked alloys. They

include four Ni-base and five Febase alloys. All of the most highly

cyclic-oxidation-resistant alloys

were protected by the formation of

alumina and aluminate scales.

Reference

1. Barrett, C.A.: 10,000-Hour Cyclic

Oxidation Behavior at 982 °C

(1800 °F) of 68 High-Temperature

Co-, Fe-, and Ni-Base Alloys. NASA

TM–107394, 1997. (Available online:

http://gltrs.grc.nasa.gov/cgi-bin/

GLTRS/browse.pl?1997/TM107394.html)

Lewis contact:

Charles A. Barrett, (216) 433–3163,

Charles.A.Barrett@grc.nasa.gov

Author: Charles A. Barrett

Headquarters program office: OAT

Programs/Projects: HITEMP
29

Research and Technology

Relative oxidation attack resistance ratings. Dashed lines indicate cyclic oxidation resistance ratings, KB4: <0.2, excellent; 0.2

to 0.5, good; 0.5 to 1.0, fair; 1.0 to 5.0, poor; >5.0, catastrophic. Top: Chromia/chromite-forming Ni-base alloys (ten 1000-hr

cycles at 982 °C). Bottom: The best of the excellent-ranked oxidation-resistant alloys (ten 1000-hr cycles at 982 °C). Oxidation resistance rating, KB4

102

101

100

10–1

10–2

Oxidation resistance rating, KB4

101

100

10–1

10–2

10–3

IN–671

Ni-40Cr

Top.-30

IN–804 Chromel A

Chromel AA

IN–601

HAS–G

HAS–X

RA–333 IN–617

DH–241

DH–242

IN–706

IN–738 HC–276

HAS-S

Chromel C Rene-80 HAS–N

IN–600 IN–X750 Chromel P

Cr levels decreasing

U–700 HOS–875 TRW-valve NASA–18T NASA-VIA Thermenol IN–702 B–1900 18SR

Alloys

Ni-base

Fe-base
30

Time (1-hr cycles)

0 250 300 50 100 150 200

125

100

75

50

25

0

Weight change, mg

NiAl-1Ti

NiAl-5Ti

NiAl-4Ti-Cr

4 to 5 at.%

Ti added

1 to 2 at.%

Cr added

Single-crystal nickel aluminide (NiAl) has been investigated extensively throughout the last several years as a potential structural

material in aero-gas turbine engines. The attractive features of NiAl

in comparison to Ni-base superalloys include a higher melting point,

lower density, higher thermal conductivity, and excellent oxidation

resistance. However, NiAl suffers from a lack of ductility and fracture

toughness at low temperatures and a low creep strength at high

temperatures. Alloying additions of hafnium (Hf), gallium (Ga),

titanium (Ti), and chromium (Cr) have each shown some benefit

to the mechanical properties over that of the binary alloy. However,

the collective effect of these alloying additions on the environmental resistance of NiAl-X was unclear. Hence, the present study was

undertaken to examine the hot corrosion behavior of these alloys.

A companion study examined the cyclic oxidation resistance of

these alloys.

Several single-crystal NiAl-X alloys (where X is Hf, Ti, Cr, or Ga) underwent

hot corrosion testing in a Mach 0.3 burner rig at the NASA Lewis Research

Center. Samples were tested for up to 300 1-hr cycles at a temperature

of 900 °C. It was found that increasing the Ti content from 1 to 5 at.%

degraded the hot corrosion behavior. This decline in the behavior was

reflected in high weight gains and large corrosion mound formation during

testing (see the figures). However, the addition of 1 to 2 at.% Cr to alloys

containing 4 to 5 at.% Ti appeared to greatly reduce the susceptibility of

these alloys to hot corrosion attack and negated the deleterious effect of

the increased Ti addition.

After testing, the surface morphology consisted of either mounds or an

inward, uniform attack that preserved surface features. The mounds

initiated early and grew and coalesced through the 300 hours of testing.

However, few new corrosion mounds were observed after 50 to 100 hours.

Microstructurally, below the surface, the hot corrosion attack initiated as

pits but evolved to a rampant attack consisting of the rapid inward growth

of Al2O3 and TiO2. This attack stage progressed rapidly inward, totally

Single-Crystal NiAl-X Alloys Tested for Hot Corrosion

Weight change as a function of time for three

single-crystal NiAl-X alloys.

Hot corrosion of three single-crystal alloys.

depleting the Al and Ti in the

degraded regions.

It was also observed that the

corrosion morphology was strongly

affected by the surface preparation

treatments. The sample surfaces

were prepared by either electropolishing, chemically milling, or

mechanical polishing. It was found

that mechanical polishing changed

the corrosion morphology and

improved the hot corrosion

response over that of chemically

milled surfaces. Further details of

this work are given in reference 1.

NiAl-1Ti NiAl-5Ti NiAl-5Ti-1Cr
31

Research and Technology

Reference

1. Nesbitt, J.A.: Hot Corrosion of Single-Crystal NiAl-X Alloys. NASA TM–113128,

1998.

Lewis contact:

Dr. James A. Nesbitt, (216) 433–3275, James.A.Nesbitt@grc.nasa.gov

Author: Dr. James A. Nesbitt

Headquarters program office: OAT

Programs/Projects: HITEMP

Recently, refractive secondary solar concentrator systems were developed

for solar thermal power and propulsion (ref. 1). Single-crystal oxides—

such as yttria-stabilized zirconia (Y2O3-ZrO2), yttrium aluminum garnet

(Y3Al5O12, or YAG), magnesium oxide (MgO), and sapphire (Al2O3)—

are candidate refractive secondary concentrator materials. However, the

refractive concentrator system will experience high-temperature thermal

cycling in the solar thermal engine during the sun/shade transition of a

space mission. The thermal mechanical reliability of these components in

severe thermal environments is of great concern. Simulated mission tests

are important for evaluating these candidate oxide materials under a variety of transient and steady-state heat flux conditions. In this research at the

NASA Lewis Research Center, a controlled heat flux test approach was

developed for investigating the thermal mechanical stability of the candidate oxide. This approach used a 3.0-kW continuous-wave (wavelength,

10.6 μm) carbon dioxide (CO2) laser (ref. 2). The CO2 laser is especially

Thermal Mechanical Stability of Single-Crystal-Oxide

Refractive Concentrators Evaluated for HighTemperature Solar-Thermal Propulsion

High-temperature thermal gradient testing of single-crystal-oxide refractive solar

concentrator using a high-power laser.

well-suited for single-crystal thermal shock tests because it can

directly deliver well-characterized

heat energy to the oxide surfaces.

Since the oxides are opaque at the

10.6-μm wavelength of the laser

beam, the light energy is absorbed

at the surfaces rather than transmitting into the crystals, and thus

generates the required temperature

gradients within the specimens.

The following figure is a schematic

diagram of the test rig.

Pyrometers

Cooling air

Specimen

Expanded 

laser beam

Thermography

system

Aperture

Charge-coupled 

discharge (CCD)

camera system

(a) Direct laser heating (b) Indirect laser heating

Superalloy Specimen plate

Cooling air

Laser beam

Thermocouple

 3.0-kW CO2 high-power laser

Laser Laser
32

The thermal stress fracture behavior and failure mechanisms of these oxide

materials were investigated, and the critical temperature gradients and

corresponding absorbed power densities imposed on the crystals to cause

failure were determined under various temperature and heating conditions. Steady-state thermal gradient tests were conducted on thin-disk

specimens of Y2O3-ZrO2, YAG, MgO, and Al2O3 (diameter, 25.4 mm;

thickness, 1 mm) to determine the critical temperature differences across

the specimen thickness that will initiate cracking. The laser power was

continuously increased at a slow rate to heat the specimen under near

steady-state conditions and to increase the temperature gradient across

the specimen thickness with the heating time. Because the disk specimen

broke off as soon as the material began to crack, the onset of cracking and

the fracture of the specimen were recorded as a sudden temperature drop

in the pyrometer temperature reading. The average critical temperature

differences at failure were 81±20 °C, 88±21 °C, and 200±35 °C for

Y2O3-ZrO2, YAG, and Al2O3, respectively. Because the thermal gradients

established in the secondary solar concentrator systems are due to the

heat receiver cavity heating/cooling cycles and to a small portion of solar

light energy being absorbed in the crystal components, the information

obtained from this experiment can be of great importance in component

design. The crack origins were typically observed near the back edges of

the specimen because of a tensile stress state near the backside, as shown

in the photomicrograph on the left (Y2O3-ZrO2).

Laser transient thermal shock tests were conducted on single-crystal

cylindrical specimens (diameter, 12.7 mm; length, 12.7 mm) with a total

heating time of 60 sec and a subsequently increased laser power. During

the test, heating/cooling profiles and temperature distributions were

experimentally measured with thermography and pyrometry. Results were

compared with analytical solutions and one-dimensional finite difference

models. The critical power densities for crystal failure were approximately

15.3 W/cm2, 56.9 W/cm2, 39.1 W/cm2, and 109.4 W/cm2 for Y2O3-ZrO2,

YAG, MgO, and Al2O3 (sapphire), respectively. In the materials with low

thermal conductivity and high thermal expansion, Y2O3-ZrO2 and MgO,

100 μm 1 mm

Crack origin

Surface

Surface

Crack and fracture morphologies of single-crystal oxides after laser testing. Left: SEM fractograph of the

Y2O3-doped ZrO2 disk specimen after laser steady-state thermal gradient testing. Right: Surface shear-stressinitiated cleavage along the (110) plane in MgO after laser thermal transient testing (optical microscopy).

high compressive stresses and/or

compressive stress-induced

(resolved) shear stresses initiated

the surface cracks. On the other

hand, in the materials with high

thermal conductivity and low thermal expansion, Al2O3 and YAG, the

cracks were more likely initiated

deep in the crystals.

The photomicrograph on the right

illustrates the surface shear-stressinitiated cleavage along the (110)

plane in MgO. The modeled transient thermal stress distributions

indicate that the surface compressive stresses and subsurface tensile

stresses are responsible for the

crystal failures observed under the

laser transient heating conditions.

Single-crystal cylindrical specimens

were also tested under a thermal

gradient to simulate the heating/

cooling cycles of a space mission. A

typical thermography temperature

distribution in a YAG specimen

during the testing is illustrated in

the final figure. Surface cracking

was observed for the Y2O3-ZrO2

and MgO specimens at a surface

heating rate of approximately

15 °C/min at 1200 °C. The YAG

and Al2O3 specimens did not crack
33

Research and Technology

13.0

11.0

12.0

10.0

9.0

7.0

8.0

6.0

5.0

3.0

4.0

2.0

1.0

0.0

–6.00 –4.00 –2.00 0.00 2.00 4.00 6.00

Distance from the surface, mm

0

1000

900

800

700

600

500

400

300

200

100

1100

1200

1300

Temperature, 

°C

Tsurface = 1232 °C

Tback Tback= 600 = 600 °°CC

Laser heating

Distance from the center, mm –6.30 6.30

under the test conditions. Al2O3 (sapphire) exhibited the best thermal

shock resistance of all oxide materials tested. The good thermal shock

resistance of sapphire is attributed to its high strength, high thermal

conductivity, and low thermal expansion coefficient.

Thermography temperature distribution in a YAG specimen (diameter, 12.7 mm;

length, 12.7 mm) during near-steady-state laser heating.

References

1. Soules, J.A., et al.: Design and

Fabrication of a Dielectric Total

Internal Reflecting Solar Concentrator

and Associated Flux Extractor for

Extreme High Temperature (2500 K)

Applications. NASA CR–204145,

1997.

2. Zhu, D.; Jacobson, N.S.; and Miller,

R.A.: Thermal-Mechanical Stability of

Single Crystal Oxide Refractive

Concentrators for High-Temperature

Solar Thermal Propulsion. NASA

TM—1999-208899, 1999.

Ohio Aerospace Institute contact:

Dr. Dongming Zhu, (216) 433–5422,

Dongming.Zhu@grc.nasa.gov

Lewis contact:

Dr. Nathan S. Jacobson, (216) 433–

5498, Nathan.S.Jacobson@grc.nasa.gov

Authors:

Dr. Dongming Zhu, Dr. Nathan S.

Jacobson, and Dr. Robert A. Miller

Headquarters program office: OAT

Programs/Projects:

Secondary Concentrator Project, Space

Transfer Technology Program
34

Modeling solar cell performance for a specific radiation environment to

obtain the end-of-life photovoltaic array performance has become both

increasingly important and, with the rapid advent of new types of cell

technology, more difficult. For large constellations of satellites, a few

percent difference in the lifetime prediction can have an enormous economic impact. The tool described here automates the assessment of solar

array on-orbit end-of-life performance and assists in the development

and design of ground test protocols for different solar cell designs. Once

established, these protocols can be used to calculate on-orbit end-of-life

performance from ground test results.

The Solar Array Verification Analysis Tool (SAVANT) utilizes the radiation

environment from the Environment Work Bench (EWB) model developed

by the NASA Lewis Research Center’s Photovoltaic and Space Environmental Effects Branch in conjunction with Maxwell Technologies. It then modifies and combines this information with the displacement damage model

proposed by Summers et al. (ref. 1) of the Naval Research Laboratory to

determine solar cell performance during the course of a given mission.

The resulting predictions can then be compared with flight data.

The Environment WorkBench (ref. 2) uses the NASA AE8 (electron) and

AP8 (proton) models of the radiation belts to calculate the trapped radiation flux. These fluxes are integrated over the defined spacecraft orbit for

the duration of the mission to obtain the total omnidirectional fluence

spectra.

Power and On-Board Propulsion Technology

Components such as the solar cell

coverglass, adhesive, and antireflective coatings can slow and attenuate the particle fluence reaching

the solar cell. In SAVANT, a continuous slowing down approximation

is used to model this effect.

The displacement damage model

is based on a damage correlation

using the “displacement damage

dose” derived from the product of

the calculated nonionizing energy

loss and the fluence. It can very

accurately predict the expected

degradation of solar cell performance in a complex space radiation environment based on Earthmade measurements with only one

proton energy and two electron

energies, or proton and electron

energy, and Co60 gammas.

The displacement damage dose is

the nonionizing equivalent of the

ionizing dose, and the nonionizing

energy loss coefficients are the nonionizing counterpart to the collision

stopping power. The displacement

damage dose is calculated by integrating the product of the differential fluence and the nonionizing

energy loss values over energy.

This displacement damage dose is

calculated for both electrons and

protons, and the total dose is

found by summing the two after

the electron dose has been converted to an effective dose by dividing it by a dose ratio experimentally

determined for the particular solar

cell type. The displacement damage dose can then be related to

solar cell power loss.

where Pmax is the maximum power

output of the cell, P0 is the original

0 50 100 150 200 250 300 350 400

Mission elapsed time, days

Savant

PASP plus

0.90

0.85

0.80

0.95

1.00

Power ratio

0 50 100 150 200 250 300 350 400

Mission elapsed time, days

Savant

PASP plus

0.90

0.95

1.00

Power ratio

Solar Array Verification Analysis Tool (SAVANT) Developed

 Normalized power ratio for the silicon module.

 Normalized power ratio for the gallium arsenide module.

P

P

C D

D

T

x

max 1 – ln 1+

0

= 



 




35

Research and Technology

SAVANT screen capture showing solar cell

radiation damage.

Authors:

Dr. Sheila G. Bailey, Kenwyn J. Long,

Henry B. Curtis, Barbara Gardner,

Victoria Davis, Scott Messenger, and

Robert Walters

Headquarters program office:

OSS (ATMS)

Programs/Projects: Spacecraft for all

solar-powered flight missions

In 1997, scientists and engineers of the Photovoltaic and Space Environments Branch of the NASA Lewis Research Center, Maxwell Technologies, and Space Systems/Loral discovered a new failure mechanism for

solar arrays on communications satellites in orbit. Sustained electrical

arcs, initiated by the space plasma and powered by the solar arrays

themselves, were found to have destroyed solar array substrates on

some Space Systems/Loral satellites, leading to array failure. The mechanism was tested at Lewis, and mitigation strategies were developed

to prevent such disastrous occurrences on-orbit in the future.

Deep Space 1 is a solar-electric-powered space mission to a comet,

launched on October 24, 1998. Early in 1998, scientists at Lewis and

Ballistic Missile Defense Organization (BMDO) realized that some

aspects of the Deep Space 1 solar arrays were nearly identical to those

that had led to the failure of solar arrays on Space Systems/Loral satellites. They decided to modify the Deep Space 1 arrays to prevent catastrophic failure in space. The arrays were suitably modified and are now

performing optimally in outer space.

Finally, the Earth Observing System (EOS) AM1, scheduled for launch in

mid 1999, is a NASA mission managed by the Goddard Space Flight Center.

Realizing the importance of Lewis’ testing on the Loral arrays, EOS–AM1

management asked Lewis scientists to test their solar arrays to show that

they would not fail in the same way. The first phase of plasma testing

Space Plasma Shown to Make Satellite Solar Arrays Fail

Sustained arc on the EOS–AM1 solar array

test article in a simulated space plasma in

the Lewis Plasma Interactions Facility.

showed that sustained arcing

would occur on the unmodified

EOS–AM1 arrays, so the arrays

were removed from the spacecraft

and fixed. Now, Lewis scientists

have finished plasma testing of the

modified array configuration to

ensure that EOS–AM1 will have no

sustained arcing problems on-orbit.

power output, C and Dx depend on the specific solar cell, and DT is

the total displacement damage dose.

A comparison of the SAVANT prediction with the albedo-free data

from the Photovoltaic Array Space Power Plus Diagnostics (PASP

Plus) flight experiment for the silicon and gallium arsenide modules

is shown in the preceding page.

The SAVANT tool fully integrates the required models end to end,

including the trapped particle environment, cell shielding, geometry

and type, slowing of the incident fluence by shielding, resultant

displacement damage dose, and effect on cell power output. Parameter trade studies are extremely easy to perform with SAVANT. The

effect of shielding material or thickness, orbit parameters, mission

length, or cell type can be examined. The SAVANT tool provides a

time-effective way to compare a variety of array parameters.

References

1. Summers, G.P.; Burke, E.A.; and Zapsos, M.A.: Displacement Damage Analogs to

Ionizing Radiation Effects. Radiation Measurements, vol. 24, no. 1, 1995, pp. 1–8.

2. EWB: The Environment WorkBench. Online information: http://

satori2.grc.nasa.gov/DOC/EWB/ewbhome.html

Lewis contact: Dr. Sheila G. Bailey, (216) 433–2228, fax (216) 433–6106,

Sheila.G.Bailey@grc.nasa.gov
36

Lewis contact: Dr. Dale C. Ferguson, (216) 433–2298, Dale.C.Ferguson@grc.nasa.gov

Author: Dr. Dale C. Ferguson

Headquarters program office: OSS (ATMS)

Programs/Projects

EOS, New Millenium, GEO communications satellites, LEO satellite constellations

A recently completed 3 1/2-year project funded by the Defense Advanced

Research Projects Agency (DARPA) under the Technology Reinvestment

Program has resulted in the development and scaleup of new lithium-ion

polymer battery technology for military and aerospace applications. The

contractors for this cost-shared project were Lockheed Martin Missiles &

Space and Ultralife Batteries, Inc. The NASA Lewis Research Center provided contract management and technical oversight. The final products of

the project were a portable 15-volt (V), 10-ampere-hour (A-hr) military

radio battery and a 30-V, 50-A-hr marine/aerospace battery. Lewis will test

the 50-A-hr battery.

The new lithium-ion polymer battery technology offers a threefold or fourfold reduction in mass and volume, relative to today’s commonly used

nickel-cadmium, nickel-hydrogen, and nickel-metal hydride batteries. This

is of special importance for orbiting satellites. It has been determined for a

particular commercial communications satellite that the replacement of

1 kg of battery mass with 1 kg of transponder mass could increase the

annual revenue flow by $100 000! Since this lithium-ion polymer technology offers battery mass reductions on the order of hundreds of kilograms

for some satellites, the potential revenue increases are impressive.

By the end of the project, the best high-energy formulations for the

electrodes and electrolyte were providing 170 watt-hours per kilogram

(W-hr/kg) specific energy and nearly 450 W-hr/liter energy density for the

lithium-ion polymer cell. Battery cells have been tested to 1000 deepdischarge cycles before 20 percent of capacity has been lost. We anticipate

that continued improvements in cell chemistry coupled with cycling to less

than full discharge will result in the cycle lives required for low-Earth-orbit

applications. Electronic circuitry has been developed to provide for chargedischarge control, cell protection, and cell rebalancing at the individual cell

level. In the lithium-ion polymer cell, there is no metallic lithium, so the

safety concerns associated with that metal are obviated. In a typical cell

formulation, the negative electrode is a carbon material that can accept

(intercalate) lithium ions at a low potential, and the positive electrode is a

transition metal oxide that intercalates the ions at a potential about 4 volts

higher. The two electrodes and the electrolyte structure all have as one

component a thermoplastic polymeric material. Thus, the three cell elements can be independently processed using common high-speed plastic

film processing technology. They can then be brought together, thermally

bonded to one another, cut to the desired size and shape, activated,

connected to form batteries, packaged, sealed, and tested. This

entire process, which can be fully

automated, uses inexpensive materials and offers the promise of an

extremely attractive advanced

technology at a low cost.

Lewis contact:

Norman H. Hagedorn, (216) 433–6157,

Norman.H.Hagedorn@grc.nasa.gov

Author: Norman H. Hagedorn

Headquarters program office: OSS

Programs/Projects:

NASA/AF Lithium-ion Battery Development, NASA Aerospace Flight Battery

Systems, RLV, Bantam, Future-X, LEO

and GEO missions, EAPU, EMA, and

in-house lithium-ion polymer battery

development

50 A-hr, 30-V lithium-ion polymer

battery.

Lithium-Ion Polymer Rechargeable Battery Developed for

Aerospace and Military Applications
37

Research and Technology

The International Space Station’s (ISS) power system is designed

with high-voltage solar arrays that typically operate at output voltages of 140 to 160 volts (V). The ISS grounding scheme electrically

ties the habitat modules, structure, and radiators to the negative tap

of the solar arrays. Without some active charge control method, this

electrical configuration and the plasma current balance would cause

the habitat modules, structure, and radiators to float to voltages as

large as –120 V with respect to the ambient space plasma.

With such large negative floating potentials, the ISS could have deleterious interactions with the space plasma. These interactions could

include arcing through insulating surfaces and sputtering of conductive surfaces as ions are accelerated by the spacecraft plasma sheath.

A plasma contactor system was baselined on the ISS to prevent arcing and

sputtering. The sole requirement for the system is contained within a single

directive (SSP 30000, paragraph 3.1.3.2.1.8): “The Space Station structure

floating potential at all points on the Space Station shall be controlled

to within ±40 V of the ionospheric plasma potential using a plasma

contactor.” NASA is developing this plasma contactor as part of the ISS

electrical power system.

For ISS, efficient and rapid emission of high electron currents is required

from the plasma contactor system under conditions of variable and uncertain current demand. A hollow cathode plasma source is well suited for

this application and was, therefore, selected as the design approach for the

station plasma contactor system. In addition to the plasma source, which is

referred to as a hollow cathode assembly, or HCA, the plasma contactor

system includes two other subsystems. These are the power electronics

unit and the xenon gas feed system.

The Rocketdyne Division of Boeing North American is responsible for the

design, fabrication, assembly, test, and integration of the plasma contactor

system. Because of technical and schedule considerations, the NASA Lewis

Research Center was asked to manufacture and deliver the engineering

model, the qualification model, and the flight HCA units for the plasma

Cathodes Delivered for Space Station Plasma

Contactor System

Flight-ready hollow cathode assembly units for

the space station plasma contactor flight

hardware (includes one qualification unit).

contactor system as government

furnished equipment. To date,

multiple units have been built.

One cathode has demonstrated

approximately 28 000-hr lifetime,

two development HCA units have

demonstrated over 15 000-hr

lifetime, and one HCA unit has

demonstrated more than 38 000

ignitions. All eight flight HCA’s have

been manufactured, acceptance

tested, and are ready for delivery to

the flight contractor.

Lewis contact:

Michael J. Patterson, (216) 977–7481,

Michael.J.Patterson@grc.nasa.gov

Author: Michael J. Patterson

Headquarters program office: OSF

Programs/Projects: ISS

TRW, under contract to the NASA Lewis Research Center, has successfully

completed over 10 000 sec of testing of a rhenium thrust chamber manufactured via a new-generation powder metallurgy. High performance was

achieved for two different propellants, N2O4–N2H4 and N2O4–MMH. TRW

conducted 44 tests with N2O4–N2H4, accumulating 5230 sec of operating

time with maximum burn times of 600 sec and a specific implulse, I

sp, of

333 sec. Seventeen tests were conducted with N2O4–MMH for an additional 4789 sec and a maximum I

sp of 324 sec, with a maximum firing

duration of 700 sec. Together, the 61 tests totalled 10 019 sec of operating

High-Performance Bipropellant Engine

time, with the chamber remaining

in excellent condition. Of these

tests, 11 lasted 600 to 700 sec.

The performance of radiationcooled rocket engines is limited by

their operating temperature. For

the past two to three decades, the

majority of radiation-cooled rockets
38

or silicon carbide, would have a

thrust-to-weight ratio of 1500:1

compared to 15:1 for a conventional engine in this thrust range.

The regeneratively cooled silicon

structure is predicted to have an

operating temperature less than

1400 K.

Because of the fundamental nature

of the work, near-term efforts will

focus on the basic feasibility issues.

These include (1) MEMS fabrication technology, (2) component/

propellant compatibility, (3) bonding and leakage, (4) heat transfer

capabilities, (5) component life,

and (6) performance.

were composed of a high-temperature niobium alloy (C103) with a

disilicide oxide coating (R512) for oxidation resistance. The R512 coating

practically limits the operating temperature to 1370 °C. For the Earthstorable bipropellants commonly used in satellite and spacecraft propulsion systems, a significant amount of fuel film cooling is needed. The large

film-cooling requirement extracts a large penalty in performance from

incomplete mixing and combustion.

A material system with a higher temperature capability has been matured

to the point where engines are being readied for flight, particularly the

100-lb-thrust class engine. This system has powder rhenium (Re) as a

substrate material with an iridium (Ir) oxidation-resistant coating. Again,

the operating temperature is limited by the coating; however, Ir is capable

of long-life operation at 2200 °C. For Earth-storable bipropellants, this

allows for the virtual elimination of fuel film cooling (some film cooling is

used for thermal control of the

head end). This has resulted in

significant increases in specific

impulse performance (15 to

20 sec).

To determine the merits of a powder rhenium thrust chamber, Lewis’

On-Board Propulsion Branch

directed TRW (under the Space

Storable Rocket Technology Program and the High Pressure Earth

Storable Rocket Technology Program) to design, fabricate, and test

an engineering model to serve as a

technology demonstrator.

Lewis contact:

Dr. Steven J. Schneider, (216) 977–7484,

Steven.J.Schneider@grc.nasa.gov

Authors: James A. Biaglow and

Dr. Steven J. Schneider

Headquarters program office: OSS

Programs/Projects:

Space Storable Rocket Technology

Program, High-Pressure Earth Storable

Rocket Technology

Technology demonstrator engine fabricated from rhenium.

Fundamental research into the feasibility of microrockets for primary

propulsion and attitude control for far-term micro/integrated spacecraft is

being performed. These rockets would be fabricated using microelectrical

and mechanical systems (MEMS) technology. The enabling technology

is being developed at the Massachusetts Institute of Technology (MIT).

The NASA/MIT program leverages a very large Army Research Office and

Defense Advanced Research Projects Agency (DARPA) program for the

development of microturbine technology. The microrocket motor is complete with regenerative cooling, turbopumps, and control valves etched

onto the same chip. They would be fabricated in large numbers in parallel

using semiconductor manufacturing techniques. The technology may lead

to the development of microsatellites as fully integrated MEMS devices that

could be mass produced at a fraction of the cost of current satellites.

This microrocket would have a specific impulse of 300 sec at a chamber

pressure of 125 atm in a 15-N-thrust class engine and would use propellants such as liquid oxygen and ethanol. The motor, fabricated from silicon

Liquid Bipropellant Microrocket Concept Under

Development
39

Research and Technology

Lewis contact:

Dr. Steven J. Schneider, (216) 977–7484,

Steven.J.Schneider@grc.nasa.gov

Author: Dr. Steven J. Schneider

Headquarters program office:

OSS (ATMS)

Programs/Projects:

New Millenium Program, MTPE

Silicon half-rocket chamber, complete with cooling channels.

For scaling, this component is shown sitting atop a nickel.

The NASA Solar Electric Propulsion Technology Applications Readiness

(NSTAR) Program has provided a single-string primary propulsion system to

NASA’s Deep Space 1 spacecraft. This spacecraft will carry about 81 kg of

xenon propellant for the ion thruster, which can be throttled down from

2.3 to 0.5 kW as the spacecraft moves away from the Sun. The propellant

load will provide about 20 months of propulsion at the one-half power

throttle setpoint of 1.2 kW. This mission will validate the 2.5-kW ion propulsion system and will fly by the asteroid 1992 KD in 1999. If funding

Ion Thruster and Power Processor Developed for the

Deep Space 1 Mission

permits, Deep Space 1 also

will encounter comets WilsonHarrington and Borrelly in 2001.

NASA Lewis Research Center’s

On-Board Propulsion Branch was

responsible for the development

of the 30-cm-diameter ion thruster,

the 2.5-kW power processor unit

(PPU), and the Digital Control and

Interface Unit (DCIU) that controls

the PPU/thruster/feed system and

provides data and recovery from

fault conditions. Lewis transferred

the thruster and PPU technologies

to the Hughes Electron Dynamics

Division, which was selected to

build two sets of flight thrusters,

as well as the PPU’s and DCIU’s.

Hughes subcontracted the DCIU

development to Spectrum Astro

Incorporated. The Jet Propulsion

Laboratory (JPL) was primarily

responsible for the NSTAR project

management, thruster lifetests, the

feed system, diagnostics, and the

propulsion subsystem integration.

A total of four engineering model

thrusters and three breadboard

PPU’s were built, integrated, and

tested. More than 50 development Ion thruster mounted to its gimbal ring on the Deep Space 1 spacecraft during

solar thermal vacuum testing.
40

tests were conducted along with thruster design verification tests of 2000

and 1000 hours. In addition, an 8000-hr life demonstration test was successfully completed and demonstrated wear-rates consistent with fullpower life-times in excess of 12 000 hours.

Because of funding and schedule constraints, the PPU’s and DCIU’s fabricated by Hughes Electron Dynamics Division were flight qualified at Lewis.

Thrusters were acceptance tested at both Lewis and the Jet Propulsion

Laboratory. All the hardware passed the qualification test program and was

subsequently integrated onto the Deep Space 1 spacecraft, where the

propulsion system completed a compatibility test program that included a

thruster firing at full-power in the Jet Propulsion Laboratory’s 25-ft spacesimulation facility. Deep Space 1 was launched from Cape Canaveral Air

Station on October 24, 1998, aboard a Boeing Delta 7326.

Bibliography

Sovey, J.S., et al.: Development of an

Ion Thruster and Power Processor

for New Millennium’s Deep Space 1

Mission. AIAA Paper 97–2778, July 1997.

Rawlin, V.K., et al.: NSTAR Flight

Thruster Qualification Testing. AIAA

Paper 98–3936, July 1998.

Lewis contacts:

James S. Sovey, (216) 977–7454,

James.S.Sovey@grc.nasa.gov; and

Vincent K. Rawlin, (216) 977–7462,

Vincent.K.Rawlin@grc.nasa.gov

Author: James S. Sovey

Headquarters program offices:

OSS, OAT

Programs/Projects: New Millennium

program, NSTAR, Deep Space 1, Deep

Space 4, AST

The Advanced Refrigerator/Freezer (R/F) Technology Development Project

was initiated in 1994, on the basis of recommendations of a team of

NASA Scientists and engineers, who assessed the need for advanced

technology to support future life and biomedical sciences space flight

missions. The project, which was cofunded by NASA’s Office of AeroSpace Technology and Life and Biomedical Sciences & Applications

Division, has two phases. In the Phase I Advanced R/F Technology Assessment, candidate technologies were identified and ranked, on the basis

of a combination of their effect on system performance and their risk of

developmental success. In Phase II Technology Development, the advanced technologies with the highest combined ranking, which could be

accomplished within the budgetary constraints, were pursued. The effort

has been mainly by contract, with a modest in-house effort at the NASA

Lewis Research Center. Oceaneering Space Systems (OSS) of Houston,

Texas, was selected as the prime contractor for both contract phases.

Advanced Refrigerator/Freezer Technology

Development Project

Six-sided vacuum panel insulated box. Stirling Orbiter Refrigerator Freezer.

The Phase I report (ref. 1) identified nine candidate advanced

technologies for development.

Resource availability forced a

limiting of the Phase II technology

development contract effort to only

the highest priority technology,

rigid polymer vacuum panel multilayer insulation (MLI). Typical

3 1/2-in.-thick fiberglass insulation

has an insulation value of R–11/in.

or R–3.14/in. The best commercially available vacuum panel

insulation has an insulation value

of about R–35/in. and has an
41

Research and Technology

expected 10-yr life with insulation value of at least R–20/in. Vacuum panel

insulation tests by OSS indicated center of panel insulation values of more

than R–100/in. Unfortunately, severe design compromises, made to expedite construction of the panels, yielded high conduction heat flow at the

panel edges. As a result, the measured overall insulation value for a sixsided box (see the left photograph on the preceding page), constructed

of these panels, was about R–11.5/in. Although this was well below the

intended goal, it agreed reasonably well with the R–13.89/in. computer

model prediction. OSS is confident that overall insulation values of

R–60/in. to R–80/in. are achievable with optimal panel edge design. The

long-term performance of these panels remains to be demonstrated.

The OSS contract has been completed, and the final report has been delivered to Lewis for NASA review and editing. It should be published in

early 1999.

The Stirling Orbiter Refrigerator Freezer (SOR/F), which flew as a space

flight experiment on the STS–60 shuttle mission, was acquired from the

NASA Johnson Space Center to serve as a test bed for in-house, advanced

refrigerator/freezer technology testing at Lewis. The current effort is aimed

at improving the SOR/F performance and making it more versatile. Key

objectives are to (1) reduce the electrical power consumption and heat

rejection, (2) reduce the rate of temperature rise with the power turned off,

(3) reduce the levels of noise and vibration, and (4) extend the operating

temperature range down to –80 °C. The original SOR/F operated at temperatures from 4 to –26 °C.

So that these objectives can be realized, the original double Stirling cycle

cooler, built by Sunpower, Inc. of Athens, Ohio, is being replaced with two

improved, higher efficiency coolers (see the photograph on the left) built

by Global Cooling Manufacturing Company, under license from Sunpower.

Lewis, with the aid of technology furnished by the NASA Goddard Space

Flight Center, is developing controls to drive the Stirling coolers to maintain operating temperature and to minimize vibration. The new coolers

use liquid cooling, which allows lower heat rejection temperature and

improved cooler efficiency in comparison to the original air-cooled Sunpower coolers. Liquid cooling also eliminates the noise of a cooling fan.

The original SOR/F used heat pipes to transport heat from the storage

volume to the Stirling cooler. The acetone fluid in these heat pipes limits

the lower operating temperature to

about –40 °C. During 1998, new

heat pipes (see the photograph on

the right) with propylene working

fluid were designed, fabricated,

and tested by Thermacore, Inc.,

of Lancaster, Pennsylvania, and

delivered to Lewis. The new heat

pipes can operate at temperatures

to –100 °C. Also, they are designed

to operate as thermal diodes,

allowing heat to flow easily from

the storage space to the cooler,

but they restrict heat flow in the

reverse direction, when the cooler

is not operating.

For further information, check out

our web page:

http://powerweb.grc.nasa.gov/stirling/

DOC/adfreeze.html

Reference

1. Gaseor, T.; Hunter, R.; and Hamill, D.:

Advanced Refrigerator/Freezer Technology Development Technology

Assessment. NASA CR–198484, 1996.

Lewis contacts:

James E. Cairelli, (216) 433–6142,

James.E.Cairelli@grc.nasa.gov; and

Steven M. Geng, (216) 433–6145,

Steven.M.Geng@grc.nasa.gov

Authors:

James E. Cairelli and Steven M. Geng

Headquarters program office:

OAT, OLMSA (LS)

Programs/Projects: Life and biomedical sciences spaceflight missions

Improved Stirling cooler. Low-temperature diode heat pipes.
42

A major modification of and addition to existing Closed Brayton Cycle

(CBC) space power system optimization codes was completed. These

modifications relate to the global minimum mass search driver programs

containing three nested iteration loops comprising iterations on cycle

temperature ratio, and three separate pressure ratio iteration loops—one

loop for maximizing thermodynamic efficiency, one for minimizing radiator

area, and a final loop for minimizing overall power system mass. Using the

method of steepest ascent, the code sweeps through the pressure ratio

space repeatedly, each time with smaller iteration step sizes, so that the

three optimum pressure ratios can be obtained to any desired accuracy for

each of the objective functions referred to above (i.e., maximum thermodynamic efficiency, minimum radiator area, and minimum system mass).

Two separate options for the power system heat source are available:

(1) A nuclear fission reactor can be used. It is provided with a radiation

shield (composed of a lithium hydride (LiH) neutron shield and tungsten

(W) gamma shield). Suboptions can be used to select the type of reactor

(i.e., fast spectrum liquid metal cooled or epithermal high-temperature gas

reactor (HTGR)).

(2) A solar heat source can be used. This option includes a parabolic concentrator and heat receiver for raising the temperature of the recirculating

working fluid.

A useful feature of the code modifications is that key cycle parameters are

displayed, including the overall system specific mass in kilograms per

Power System Optimization Codes Modified

kilowatt and the system specific

power in watts per kilogram, as the

results for each temperature ratio

are computed. As the minimum

mass temperature ratio is encountered, a message is printed out.

Several levels of detailed information on cycle state points, subsystem mass results, and radiator

temperature profiles are stored for

this temperature ratio condition

and can be displayed or printed by

users.

Lewis contact:

Albert J. Juhasz, (216) 433–6134,

Albert.J.Juhasz@grc.nasa.gov

Author: Albert J. Juhasz

Headquarters program office:

OSS (ATMS)

Programs/Projects:

SSP, Lunar Base Power System, Mars

Base Power, multimegawatt power

system for interplanetary electric

propulsion systems

A battery charge regulator based on the series-connected boost regulator (SCBR) technology has been developed for high-voltage spacecraft

applications. The SCBR regulates the solar array power during insolation

to prevent battery overcharge or undercharge conditions. It can also be

used to provide regulated battery output voltage to spacecraft loads if

necessary. This technology uses industry-standard dc-dc converters and

a unique interconnection to provide size, weight, efficiency, fault tolerance, and modularity benefits over existing systems.

The high-voltage SCBR shown in the photograph has demonstrated

power densities of over 1000 watts per kilogram (W/kg). Using four

150-W dc-dc converter modules, it can process 2500 W of power at

120 Vdc with a minimum input voltage of 90 Vdc. Efficiency of the SCBR

was 94 to 98 percent over the entire operational range.

Internally, the unit is made of two separate SCBR’s, each with its own

analog control circuitry, to demonstrate the modularity of the technology.

The analog controllers regulate the output current and incorporate the

output voltage limit with active current sharing between the two units.

Lightweight Battery Charge Regulator Used to Track

Solar Array Peak Power

2.5-kW Series-connected boost regulator.

They also include voltage and

current telemetry, on/off control,

and baseplate temperature sensors.

For peak power tracking, the SCBR

was connected to a LabViewbased data acquisition system for
43

Research and Technology

telemetry and control. A digital control algorithm for tracking the peak

power point of a solar array was developed using the principle of matching the source impedance with the load impedance for maximum energy

transfer. The algorithm was successfully demonstrated in a simulated

spacecraft electrical system at the Boeing PhantomWorks High Voltage Test

Facility in Seattle, Washington. The system consists of a 42-string, highvoltage solar array simulator, a 77-cell, 80-ampere-hour (A-hr) nickelhydrogen battery, and a constant power-load module. The SCBR and the

LabView control algorithm successfully tracked the solar array peak power

point through various load transients, including sunlight discharge transients

when the total load exceeded the maximum solar array output power.

Lewis contacts:

James F. Soeder, (216) 433–5328,

James.F.Soeder@grc.nasa.gov; and

Robert M. Button, (216) 433–8010,

Robert.M.Button@grc.nasa.gov

Authors: James F. Soeder and

Robert M. Button

Headquarters program office:

OSS (ATMS)

Programs/Projects: High-voltage

Spacecraft; satellites powered by solar

arrays; other solar array and battery

applications

Spacecraft in low Earth orbit (LEO) are subjected to many components of

the environment, which can cause them to degrade much more rapidly

than intended and greatly shorten their functional life. The atomic oxygen,

ultraviolet radiation, and cross contamination present in LEO can affect

sensitive surfaces such as thermal control paints, multilayer insulation,

solar array surfaces, and optical surfaces. The LEO Spacecraft Materials Test

(LEO–SMT) program is being conducted to assess the effects of simulated

LEO exposure on current spacecraft materials to increase understanding of

LEO degradation processes as well as to enable the prediction of in-space

performance and durability. Using ground-based simulation facilities to test

the durability of materials currently flying in LEO will allow researchers to

International Low-Earth-Orbit Spacecraft Materials

Test Program Initiated for Better Prediction of

Durability and Performance

compare the degradation evidenced in the ground-based facilities with that evidenced on orbit.

This will allow refinement of

ground laboratory test systems

and the development of algorithms

to predict the durability and performance of new materials in LEO

from ground test results. Accurate

predictions based on ground tests

could reduce development costs

and increase reliability.

The wide variety of national and

international materials being tested

represent materials being functionally used on spacecraft in LEO. The

more varied the types of materials

tested, the greater the probability

that researchers will develop and

validate predictive models for

spacecraft long-term performance

and durability. Organizations that

are currently participating in the

program are ITT Research Institute

(USA), Lockheed Martin (USA),

MAP (France), SOREQ Nuclear

Research Center (Israel), TNO

Institute of Applied Physics (The

Netherlands), and UBE Industries,

Ltd. (Japan). These represent some

of the major suppliers of thermal

Directed atomic oxygen and synergistic vacuum ultraviolet radiation control and sensor materials

exposure facility at NASA Lewis can measure the reflectance of samples

from 250 to 2500 nm while they are in the vacuum chamber.
44

currently flying in LEO. The participants provide materials that are exposed

to selected levels of atomic oxygen, vacuum ultraviolet radiation, contamination, or synergistic combined environments at the NASA Lewis Research

Center. Changes in characteristics that could affect mission performance or

lifetime are then measured. These characteristics include changes in mass,

solar absorptance, and thermal emittance. The durability of spacecraft

materials from U.S. suppliers is then compared with those of materials

from other participating countries. Lewis will develop and validate performance and durability prediction models using this ground data and available space data. NASA welcomes the opportunity to consider additional

international participants in this program, which should greatly aid future

spacecraft designers as they select materials for LEO missions.

Lewis contacts:

Bruce A. Banks, (216) 433–2308,

Bruce.A.Banks@grc.nasa.gov; and

Sharon K. Rutledge, (216) 433–2219,

Sharon.K.Rutledge@grc.nasa.gov

Author: Sharon K. Rutledge

Headquarters program office:

OSS (ATMS)

Programs/Projects: ISS, EOS

A Russian solar array panel removed in November 1997 from the nonarticulating photovoltaic array on the Mir core module was returned to Earth on

STS–89 in January 1998. The panel had been exposed to low Earth orbit

(LEO) for 10 years prior to retrieval. The retrieval provided a unique opportunity to study the effects of the LEO environment on a functional solar

array. To take advantage of this opportunity, a team composed of members

from RSC–Energia (Russia), the Boeing Company, and the following NASA

Centers—Johnson Space Center, Kennedy Space Center, Langley Research

Center, Marshall Space Flight Center, and Lewis Research Center—was

put together to analyze the array. After postretrieval inspections at the

Spacehab Facility at Kennedy in Florida, the array was shipped to Lewis in

Cleveland for electrical performance tests, closeup photodocumentation,

and removal of selected solar cells and blanket material. With approval

from RSC–Energia, five cell pairs and their accompanying blanket and

mesh material, and samples of painted handrail materials were selected

for removal on the basis of their ability to provide degradation information.

Sites were selected that provided different sizes and shapes of micrometeoroid impacts and different levels of surface contamination. These materials were then distributed among the team for round robin testing.

A preliminary examination of the

solar array panel indicated significant darkening and diffuse light

scattering on the surface of the

panel and structural members. By

their location and appearance, the

darkened surfaces appear to be

silicone contamination from

adhesives used to assemble the

solar array panel. NASA Lewis is

performing tests to look at the

contamination on the array and to

determine the effect of atomic

oxygen and vacuum ultraviolet

radiation on contaminant fixation.

Two samples have been evaluated

to date—a handrail section (whitepainted aluminum) and a woven

polymer fabric used as an overwrap

tape on a handrail. Both samples

had visible dark-brown contamination. Optical microscopy revealed

mud-tile-like crazing patterns in

the contaminated areas. Field emission microscopy showed that the

contaminant layer was very thick

(approximately 1.6 μm in some

areas) and had spalled off in some

regions. Energy-dispersive spectroscopy confirmed that the contaminant layer on both the handrail

and the overwrap tape was silicon

oxides (SiOx) resulting from the

oxidation of silicone.

Results are being shared with the

other team members, who are

each performing a different analysis, such as micrometeoroid and

Retrieval of Mir Solar Array

Solar reflector on the back of one of the Mir solar cells, showing a

micrometeoroid impact and contamination darkening in areas not

shielded by the woven mesh on the back of the array (net pattern).
45

Research and Technology

debris impact measurements and electrical measurements. Degradation

mechanisms described through this analysis will provide valuable information that can be used in predicting the performance of solar arrays on the

International Space Station and that could prevent contamination of the

level observed on the Mir array.

Lewis contacts:

Bruce A. Banks, (216) 433–2308, Bruce.A.Banks@grc.nasa.gov;

Kim K. de Groh, (216) 433–2297, Kim.K.deGroh@grc.nasa.gov; and

Sharon K. Rutledge, (216) 433–2219,

Sharon.K.Rutledge@grc.nasa.gov

Authors:

Sharon K. Rutledge and Kim K. de Groh

Headquarters program office:

OSS (ATMS)

Programs/Projects: ISS

During the Hubble Space Telescope (HST) second servicing mission,

astronauts noticed that the outer layer of the multilayer insulation (MLI)

was cracked in many locations around the telescope. The insulation’s

outer layer is composed of 5-mil (0.127 mm) Teflon FEP (DuPont; fluorinated ethylene propylene) with vapor-deposited aluminum (VDA) on the

backside. The MLI blankets, which are used on over 80 percent of the

external surface of the telescope, provide passive thermal control for

equipment. Two large cracks were observed on the light shield directly

above the high gain antenna. The upper light shield crack propagated

from a cut placed in the blanket during installation around a handrail

standoff. Two cracks propagated almost normal to each other, with the

outer layer curling tightly as the cracks propagated. A second much larger

vertical crack had started to curl and lift away from the telescope, as seen

in the figure. MLI on the equipment bays was cracked extensively also,

and in some areas it had pulled away from the bay. Continued degradation of the equipment bay insulation would potentially cause limited

observations because of housekeeping boxes; for example, the

data interface unit could overheat

during some portions of the year,

in certain sun angles.

The tightly curled, cracked FEP/VDA

on the upper light shield was

retrieved by astronauts and found

to be severely embrittled, as witnessed by ground testing. The

NASA Goddard Space Flight Center

organized a Hubble Space Telescope MLI Failure Review Board,

which included three members

from Lewis, to (1) determine the

damage mechanism of the (FEP/

VDA) in the telescope’s orbit environment, (2) predict the condition

of the MLI at the time of the next

servicing mission, and (3) recommend a replacement thermal

control outer layer material to be

installed on the telescope during

subsequent servicing missions. The

Failure Review Board chose candidate thermal control replacement

materials by using a multiplicative

process recommended by the

NASA Lewis Research Center. Ten

candidate materials were tested

for environmental durability under

various exposures and durations

as overseen by Goddard and Lewis.

Lewis conducted x-ray radiation

exposure, thermal cycling under

load, and large-sample thermal

cycling of candidate materials.

Lewis’ durability testing of candiThermal Control Replacement Materials Evaluated for

Durability and Selected for the Hubble Space Telescope

Photograph of Hubble Space Telescope taken during the second servicing mission,

showing the very large, vertical light shield cracked area and the tightly curled

upper light shield cracked area.
46

date replacement materials and involvement on the Failure Review Board

had a direct impact on the choice of the replacement thermal control

material to be placed on the Hubble Space Telescope during subsequent

servicing missions in 2000 and 2003. The recommended replacement

thermal control material chosen by the review board and adopted by the

Hubble Space Telescope project office is 5-mil FEP/VDA bonded with a

low-outgassing ultraviolet-durable adhesive to a Nomex (DuPont) scrim.

Lewis contact:

Kim K. de Groh, (216) 433–2297,

Kim.K.deGroh@grc.nasa.gov

Authors:

Kim K. de Groh, Bruce A. Banks, and

Patricia A. Hansen

Headquarters program office:

OSS (ATMS)

Programs/Projects: HST

Solar-dynamic space power systems require durable, high-emittance

surfaces on a number of critical components, such as heat receiver interior

surfaces and parasitic load radiator (PLR) elements. An alumina-titania

coating, which has been evaluated for solar-dynamic heat receiver canister

applications, has been chosen for a PLR application (an electrical sink for

excess power from the turboalternator/compressor) because of its demonstrated high emittance and high-temperature durability in vacuum. Under

high vacuum conditions (±10–6 torr), the alumina-titania coating was

found to be durable at temperatures of 1520 °F (827 °C) for ≈2700 hours

with no degradation in optical properties. This coating has been successfully applied to the 2-kW solar-dynamic ground test demonstrator at the

NASA Lewis Research Center, to the 500 thermal-energy-storage containment canisters inside the heat receiver and to the PLR radiator. The solardynamic demonstrator has successfully operated for over 800 hours in

Lewis’ large thermal/vacuum space environment facility, demonstrating

the feasibility of solar-dynamic power generation for space applications.

Effects Investigated of Ambient High-Temperature Exposure

on Alumina-Titania High-Emittance Surfaces

Thermal emittance (at 1200 °F) versus time of alumina-titania-coated samples

that were heated in air at various temperatures.

Although the alumina-titania

coating has operated successfully

in the solar-dynamic ground test

demonstrator system, solardynamic flight PLR hardware would

require ambient pressure operation

to verify proper system electrical

function and to check out alternator control software. If operation of

the PLR in air decreases the emittance, it will operate at a higher

temperature than desired while

on orbit. This could shorten the

PLR’s life, and possibly damage

the elements. Therefore, a program

was conducted at Lewis to test the

high-temperature durability of the

alumina-titania coating in air.

Sixteen alumina-titania-coated

Incoloy samples were prepared,

and fifteen were exposed to high

temperatures (316 °C (600 °F) to

816 °C (1500 °F)) for various times

(2 to 32 hours). Samples were

characterized prior to and after

heat treatment for reflectance,

solar absorptance, room temperature emittance, and emittance at

1200 °F (649 °C). Samples also

were examined for physical defects

and surface chemistry using optical

microscopy, scanning electron

microscopy operated with an

energy dispersive spectroscopy

(EDS) system, and x-ray photoelectron spectroscopy (XPS).

Thermal emittance at 1200 °F 1.0

0.8

0.6

0 5 10

Time, hr

15 20 35 25 30

600

1000

1250

1500

Heating

temperature,

°F
47

Research and Technology

Visual examination of the heat-treated samples showed a whitening of

samples exposed to temperatures of 1000 °F (538 °C) and above. Correspondingly, the optical properties of these samples were degraded.

Decreases in thermal emittance at 1200 °F are shown in the graph. A sample exposed to 1500 °F (816 °C) for 24 hours appeared white, and the

thermal emittance at 1200 °F had decreased from the non-heat-treated

value of 0.94 to 0.62. The coating on this sample was embrittled and

spalling off the substrate in several locations. On the basis of this research,

we do not recommend that components with alumina-titania coatings be

operated at temperatures above 600 °F (316 °C) in air because optical

degradation will occur, and structure degradation of the coating may occur

also. These results also indicate that components with the alumina-titania

coating are likely to experience optical property degradation with atomic

oxygen exposure in space. Therefore, the atomic oxygen durability of this

coating needs to be assessed prior to use on surfaces directly exposed to

atomic oxygen in the low-Earth-orbit space environment.

Lewis contact:

Kim K. de Groh, (216) 433–2297,

Kim.K.deGroh@grc.nasa.gov

Authors:

Kim K. de Groh and Daniela C. Smith

Headquarters program office:

OSS (ATMS)

Programs/Projects:

SDFD, SD GTD, future solar dynamic

power systems and systems utilizing

heat received technology

The Polymers Erosion and Contamination Experiment (PEACE) is currently

being developed at the NASA Lewis Research Center by the Electro-Physics

Branch in conjunction with students and faculty from Hathaway Brown

School in Cleveland. The experiment is a Get Away Special Canister shuttle

flight experiment sponsored by the American Chemical Society. The two

goals of this experiment are (1) to measure ram atomic oxygen erosion

rates of approximately 40 polymers that have potential use in space applications and (2) to validate a method for identifying sources of silicone

contamination that occur in the shuttle bay. Equipment to be used in this

flight experiment is shown in the schematic diagram.

Spacecraft materials subjected to

attack by atomic oxygen in the

space environment experience

significant degradation over the

span of a typical mission. Therefore, learning the rates of atomic

oxygen erosion of a wide variety

of polymers would be of great

benefit to future missions. PEACE

will use two independent techniques to determine the atomic

oxygen erosion rates of polymers.

Large (1-in.-diameter) samples

will be used for obtaining mass

loss. Preflight and postflight dehydrated masses will be obtained,

and the mass lost during flight

will be determined. Small (0.5-in.-

diameter) samples will be protected with isolated particles

(such as NaCl crystals) and then

exposed to the space environment. After flight, the protective

particles will be removed (washed

off) and atomic force microscopy

(AFM) will be used to measure

the erosion depth from protected

mesas. Erosion depth measurements are more sensitive than

traditional mass measurements

Polymers Erosion and Contamination Experiment

Being Developed

Schematic diagram of PEACE showing polymer erosion samples, silicone

contamination samples, atomic oxygen pinhole cameras, and contamination

pinhole cameras.

Atomic oxygen

camera

Contamination

camera

Contamination

samples

Erosion

samples
48

and are very useful for materials with low erosion yields or with very low

fluence missions.

PEACE will validate silicone contamination sources by using special contamination pinhole cameras with views of three different types of silicones.

Six silicone samples will be viewed—three with atomic oxygen exposure

and three without atomic oxygen exposure. The atomic oxygen interaction

with the silicone should result in greater contamination; this can be validated with the cameras. The contamination cameras will also have a view

of atomic oxygen, for atomic oxygen fixing of the contamination. Scanning

ellipsometry will be used to produce an image of the source of contaminants. Using pinhole cameras to detect sources of silicone contamination

will provide a model for possible use in future shuttle operations.

Three students from Hathaway Brown School spent the 1998 summer at

Lewis, and five students worked during the 1998–1999 school year, designing the sample configuration, ordering samples, and learning some of

the technical aspects of putting together an experiment. Members of the

Electro-Physics Branch and Dynacs Engineering Co., Inc., are overseeing

the development of the experiment, the faculty from Hathaway

Brown School are coordinating

the students’ efforts, and the

Space Flight Projects Branch is

serving as the project manager.

Lewis contact:

Kim K. de Groh, (216) 433–2297,

Kim.K.deGroh@grc.nasa.gov

Authors:

Kim K. de Groh, Bruce A. Banks, Elyse

A. Baney-Barton, Edward Sechkar, and

Patricia Hunt

Headquarters program office:

OSS (ATMS)

Programs/Projects: ISS, MPTE, future

low-Earth-orbit satellites

Since the Hubble Space Telescope (HST) was deployed in low Earth orbit

in April 1990, two servicing missions have been conducted to upgrade its

scientific capabilities. Minor cracking of second-surface metalized Teflon

FEP (DuPont; fluorinated ethylene propylene) surfaces from multilayer

insulation (MLI) was first observed upon close examination of samples

with high solar exposure retrieved during the first servicing mission, which

was conducted 3.6 years after deployment. During the second HST servicing mission, 6.8 years after deployment, astronaut observations and

photographic documentation revealed significant cracks in the Teflon FEP

layer of the MLI on both the solar- and anti-solar-facing surfaces of the

telescope. NASA Goddard Space Flight Center directed the efforts of the

Hubble Space Telescope MLI Failure Review Board, whose goals included

identifying the low-Earth-orbit environmental constituent(s) responsible for

the cracking and embrittling of Teflon FEP which was observed during the

second servicing mission. The NASA Lewis Research Center provided

significant support to this effort.

Because soft x-ray radiation from solar flares had been considered as a

possible cause for the degradation of the mechanical properties of Teflon

FEP (ref. 1), the effects of soft x-ray radiation and vacuum ultraviolet light

on Teflon FEP were investigated. In this Lewis-led effort, samples of Teflon

FEP with a 100-nm layer of vapor-deposited aluminum (VDA) on the

backside were exposed to synchrotron radiation of various vacuum ultraviolet and soft x-ray wavelengths between 18 nm (69 eV) and 0.65 nm

(1900 eV). Synchrotron radiation exposures were conducted using the

National Synchrotron Light Source at Brookhaven National Laboratory.

Samples of FEP/VDA were exposed with the FEP surface facing the synchrotron beam. Doses and fluences were compared with those estimated

for the 20-yr Hubble Space Telescope mission.

The table summarizes the results

of tensile testing of synchrotron

radiation-exposed FEP/VDA and

gives comparisons to samples

retrieved from the HST. As indicated in the table, specimens with

the worst damage—those with a

loss of elongation greater than

50 percent—received synchrotron

exposure doses significantly in

excess of HST’s 20-yr lifetime

estimated dose. However, these

samples did not experience the

complete loss of elongation

shown by the samples retrieved

from the telescope, which were

exposed to the space environment

for 6.8 years. This indicates that, at

HST 6.8-yr equivalent exposure

doses, using synchrotron radiation

between 69 and 1900 eV, it is not

possible to produce the extent of

degradation observed for HST

space-exposed materials. Some

of the least-damaged specimens,

those with less than 10-percent

loss of elongation, were also

exposed to radiation doses significantly in excess of the HST

Synchrotron Vacuum Ultraviolet Light and Soft X-Ray

Radiation Effects on Aluminized Teflon FEP Investigated
49

Research and Technology

20-yr estimated dose. Although exposure to synchrotron radiation in

the vacuum ultraviolet and soft x-ray range of energies can degrade the

mechanical properties of Teflon FEP, this radiation alone could not be the

cause for the severe degradation of FEP surfaces that was observed after

the 6.8-yr exposure on the Hubble Space Telescope. Other environmental

factors, such as continuum x-ray radiation of energies up to 10 keV, electron and proton radiation, thermal cycling, and synergistic effects of radiation and thermal cycling, have also been investigated by NASA Lewis and

NASA Goddard. Exposure to electron and proton radiation followed by

thermal cycling similar to HST mission exposures caused greater degradation than exposure to radiation alone. However, further testing is required

to positively identify the environmental exposure conditions responsible

for the severe degradation of the FEP/VDA material retrieved after 6.8-yr

exposure on the HST.

Reference

1. Milintchouk, A., et al.: Influence of

X-Ray Solar Flare Radiation on

Degradation of Teflon“ in Space.

J. Spacecr. Rockets, vol. 34, no. 4,

Jul.-Aug. 1997, pp. 542–548.

Lewis contact:

Joyce A. Dever, (216) 433–6294,

Joyce.A.Dever@grc.nasa.gov

Goddard contact: Jacqueline A.

Townsend, (301) 286–6685,

jtownsend@pop300.gsfc.nasa.gov

Authors:

Joyce A. Dever, Jacqueline A. Townsend,

Dr. James R. Gaier, and Alice I. Jalics

Headquarters program office: OSS

Programs/Projects: HST

SUMMARY OF TENSILE TESTING OF SYNCHROTRON-EXPOSED FEP/VDA SAMPLES AND

COMPARISON TO SAMPLES RETRIEVED FROM THE HST

Energy,

eV

Attenuation

length,b

μm

Fluence,

J/m2

Sample dose/

HST 20-yr

x-ray dose

Loss of

ultimate

tensile strength

UTS,

percent

Loss of

elongation,

percent

Most damaged: >50 percent loss of elongation

1489 2.74 95 900 1482 c 28 83

1489 2.74 16 100 249c 30 67

290

700a

.262

.403

1 600

1 500

 207c

 136c

26 66

290

700a

.262

.403

1 600

1 500

 216c

 137c

27 59

Least damaged: <10 percent loss of elongation

69 0.033 5 300 5638 c –0.2 1.5

1489 2.74 200 2.5c –5.5 –7.2

1489 2.74 8 200 127c –6.7 –7.7

1900 5.36 3 000 27.5c –.2 –2.0

Samples retrieved from HST

1500 to 25 000 ------ 225 0.32d 42

37

46

68

1500 to 25 000 ------ 255 .36 d 50

92

100

100 a

Samples were exposed sequentially to 290- and 700-eV radiation.

bAttenuation length for a specific energy of radiation is defined as the depth into material where

radiation intensity has been reduced to 1/e times the intensity at the surface.

cSample dose is that absorbed at an attenuation length of synchrotron exposure energy. The HST 20-yr

x-ray dose is for x-rays in the energy range 1.5 to 12 keV absorbed within the attenuation length for the

synchrotron exposure energy.

d

Dose absorbed within the full 127-μm specimen thickness for the range of x-ray radiation energies

between 1.5 and 25 keV.
50

During the Hubble Space Telescope (HST) Second Servicing Mission,

6.8 years after the telescope was deployed in low Earth orbit, degradation

of unsupported Teflon FEP (DuPont; fluorinated ethylene propylene), used

as the outer layer of the multilayer insulation (MLI) blankets, was evident

as large cracks on the telescope light shield. A sample of the degraded

outer layer (see the photograph) was retrieved during the second servicing mission and returned to Earth for ground testing and evaluation. Also

retrieved was a Teflon FEP radiator surface from a cryogen vent cover that

was exposed to the space environment on the aft bulkhead of the HST.

NASA Goddard Space Flight Center directed the efforts of the Hubble Space

Telescope MLI Failure Review Board, whose goals included determining the

FEP degradation mechanisms. As part of the investigations into the degradation mechanisms, specimens retrieved from the first and second HST

servicing missions, 3.6 and 6.8 years after launch, respectively, were characterized through exhaustive mechanical, optical, and chemical testing.

Testing led by Goddard included scanning electron microscopy, optical

microscopy, tensile testing, solar absorptance measurements, time-of-flight

secondary ion mass spectroscopy (TOF–SIMS), Fourier transform infrared

microscopy (μ-FTIR), attenuated total reflectance infrared microscopy

(ATR/FTIR), and x-ray diffraction (XRD). The NASA Lewis Research Center

contributed significantly to the analysis of the retrieved HST materials by

leading efforts and providing results of bend testing, surface microhardness measurements, x-ray photoelectron spectroscopy, solid-state nuclear

magnetic resonance spectroscopy, and density measurements. Other

testing was conducted by Nano Instruments, Inc., and the University of

Akron.

Results of mechanical properties

testing of FEP specimens retrieved

after 3.6 and 6.8 years of HST

exposure indicated significant

decrease in tensile test elongation

at break and reduced tolerance

to strain upon bending. Spaceexposed samples also showed

increased surface hardness that

decreased with depth. Solar

absorptance of the space-exposed

FEP was slightly increased. In general, analyses showed that all of

the retrieved specimens underwent chain scission. Evidence of

increased crystallinity was found

only in the light shield specimen

from the second servicing mission,

which had curled on orbit and thus

experienced a significantly higher

temperature than is typical for the

HST light shield. Heating specimens

from the first servicing mission also

generated increased crystallinity.

The conclusions of the Hubble

Space Telescope MLI Failure

Review Board were based on the

combined evidence of HST damage

and data uncovered in groundbased experiments. The review

board concluded the following:

“The observations of HST MLI and

ground testing of pristine samples

indicate that thermal cycling with

deep-layer damage from electron

and proton radiation are necessary

to cause the observed Teflon FEP

embrittlement and the propagation

of cracks along stress concentrations. Ground testing and analysis

of retrieved MLI indicate that

damage increases with the combined total dose of electrons, protons, UV, and x-rays along with

thermal cycling.”

Teflon FEP Analyzed After Retrieval From the

Hubble Space Telescope

Hubble Space Telescope light shield sample retrieved during the second

servicing mission; shown prior to removal.
51

Research and Technology

Tests continue to determine the effects of the higher temperature limit that

the second servicing mission MLI specimen experienced.

Lewis contacts:

Joyce A. Dever, (216) 433–6294, Joyce.A.Dever@grc.nasa.gov; and Kim K. de Groh,

(216) 433–2297, Kim.K.deGroh@grc.nasa.gov

Goddard contacts:

Patricia A. Hansen, (301) 286–0564, phansen@mscmail.gsfc.nasa.gov; and

Jacqueline A. Townsend, (301) 286–6685, jtownsen@pop300.gsfc.nasa.gov

Authors:

Joyce A. Dever, Kim K. de Groh,

Jacqueline A. Townsend, Patricia A.

Hansen, Bruce A. Banks, Len Wang, and

Charles He

Headquarters program office: OSS

Programs/Projects: HST

Weight savings as high as 80 percent could be achieved if graphite

polymer composites could replace aluminum in structures such as

electromagnetic interference shielding covers and grounding planes.

This could result in significant cost savings, especially for the mobile

electronics found in spacecraft, aircraft, automobiles, and hand-held

consumer electronics. However, such composites had not yet been

fabricated with conductivity sufficient to enable these applications.

To address this lack, a partnership of the NASA Lewis Research Center,

Manchester College, and Applied Sciences, Inc., fabricated nonmetallic

composites with unprecedented electrical conductivity. For these composites, heat-treated, vapor-grown graphite fibers were selected which

have a resistivity of about 80 μΩ-cm, more than 20 times more conductive than typical carbon fibers. These fibers were then intercalated with

iodine bromide (IBr). Intercalation is the insertion of guest atoms or

molecules between the carbon planes of the graphite fibers. Since the

carbon planes are not highly distorted in the process, intercalation has

little effect on mechanical and thermal properties. Intercalation does,

however, lower the carbon fiber resistivity to less than 10 μΩ-cm, which

is comparable to that of metal fibers.

Scaleup of the reaction was required since the initial intercalation experiments would be carried out on 20-mg quantities of fibers, and tens of

grams of intercalated fibers would be needed to fabricate even small

demonstration composites. The reaction was first optimized through a time

and temperature study that yielded fibers with a resistivity of 8.7±2 μΩ-cm

when exposed to IBr vapor at 114 °C for 24 hours. Stability studies indicated that the intercalated fibers rapidly lost their conductivity when

exposed to temperatures as low as 40 °C in air. They were not, however,

susceptible to degradation by water vapor in the manner of most graphite

intercalation compounds. The 1000-fold scaleup experiments concluded

that 114 °C was near the optimum temperature, but that the intercalation

time needed to be lengthened by a factor of 3.

Laminar composites were hand laid up with the graphite fibers and two

different resins, a room temperature epoxy and a cyanate ester resin with a

cure temperature of 175 °C. The resistivity of the resulting composites was

200±25 μΩ-cm, nearly a factor of two lower than for similar composites

made with intercalated pitch-based fibers. The fibers in the cyanate ester

composites did not appear to lose conductivity even though the cure

temperature was well above the

temperature at which the fibers

lost conductivity in air. Although

this is the lowest resistivity composite yet fabricated, calculations

indicate that it should be possible

to use these fibers to make composites that are 6 to 7 times more

conductive.

Lewis contact: Dr. James R. Gaier,

(219) 982–5075 (Feb—May, Sept—

Dec); (216) 433–6686 (Jan, Jun—Aug),

James.R.Gaier@grc.nasa.gov

Author: Dr. James R. Gaier

Headquarters Program Office:

OSS (ATMS)

Programs/Projects: High-conductivity,

low-weight applications, such as

notebook computers and cell phones

Highly Conducting Graphite Epoxy Composite Demonstrated

Composite resistivity of different graphite

fibers. Log resistivity, μΩ-cm

4

2

3

1

0

PAN

P–55

P–100+Br

VGCF+IBr

Theory

Aluminum

VGCF
52

Materials located

on the exterior of

spacecraft in low

Earth orbit are subjected to a number

of environmental

threats, including

atomic oxygen,

ultraviolet radiation,

thermal cycling, and

micrometeroid and

debris impact.

Atomic oxygen

attacks materials

vulnerable to

oxidation. Ultraviolet radiation can

break chemical

bonds and cause

undesirable

changes in optical

properties. Thermal

cycling can cause

cracking, and micrometeroid and debris

impacts can damage protective coatings. Another threat is contamination. The outgassing of volatile chemicals

can contaminate nearby surfaces, changing their thermal control properties. Contaminated surfaces may undergo further change as a result of

atomic oxygen and ultraviolet radiation exposure.

The Passive Optical Sample Assembly (POSA) experiment was designed

as a risk mitigation experiment for the International Space Station. Samples

were characterized before launch, exposed for 18 months on the exterior

of Mir, and characterized upon their return. Lessons learned from POSA

about the durability of material properties can be applied to the space

station and other long-duration missions.

Thirty-two Lewis samples were on POSA–2. The suitcase-sized mounting

hardware, which was attached to the shuttle docking module, allowed half

of the samples to face Mir and the other half of the samples to face space.

Those samples facing space also faced the shuttle when it was docked to Mir.

Postflight evaluation of the samples revealed atomic oxygen erosion,

changes in the optical properties of solar absorptance and infrared emittance, and contamination from at

least two different sources. Dehydrated Kapton mass loss measurements identified an atomic oxygen

fluence of 8.2x1019 atoms/cm2 for

the Mir-facing side and 2.1x1020

atoms/cm2 for the space-facing

side. The mass of contamination

was found to be negligible.

Contamination changed some

optical properties. For example,

the solar absorptance of aluminum

mirror samples increased in all

cases, with the increase being

greater for the Mir-facing samples.

Such increases are undesirable

because they cause unwanted

heating of spacecraft surfaces. It

was also observed that the infrared

emittance of thermal control paint

samples decreased, with the

amount of decrease essentially

the same for both the Mir-facing

and space-facing samples. Such

decreases are undesirable because

they diminish the heat-rejection

capability of the thermal control

paint.

Chemical evaluation of several

samples identified a silicon signature. The origin of the silicon is

thought to be the outgassing of a

silicone paint during one or more

zero-eclipse orbits. The source for

the other contaminant(s) is still

under investigation.

Lewis contact:

Dr. Donald A. Jaworske, (216) 433–2312,

Donald.A.Jaworske@grc.nasa.gov

Author: Dr. Donald A. Jaworske

Headquarters program office:

OSS (ATMS)

Programs/Projects: ISS, POSA

Changes in the Optical Properties of Materials Are Observed

After 18 Months in Low Earth Orbit

Opening the Passive Optical Sample Assembly (POSA)

experiment after 18 months in low Earth orbit.
53

Research and Technology

In many future NASA missions, such as deep space planetary exploration and the Next Generation Space Telescope, electrical components and systems must operate reliably and efficiently in extremely

low temperature environments. Most modern electronic components cannot operate below moderately low operating temperatures

(–40 to –55 °C).

The low-temperature electronics program at the NASA Lewis

Research Center is focusing on the development and characterization of low-temperature components and the integration of

the developed devices into demonstrable very low-temperature

(–200 °C) power systems such as dc-dc converters. Such lowtemperature electronics will not only tolerate hostile environments

but also will reduce system size and weight by eliminating radioisotope heating units, thereby reducing launch cost, improving

reliability and lifetime, and increasing energy densities.

Low-temperature electronic components will also have a great influence

on terrestrial applications such as medical instrumentation, magnetic levitation transportation systems, and arctic and antarctic exploration. Lewis

researchers are now performing extensive evaluations of commercially

available as well as custom-made devices. These include various types of

energy storage and signal capacitors, power switching devices, magnetic

and superconducting materials, and primary lithium batteries, to name

a few. The following photograph shows solid tantalum and ceramic

capacitors that were recently evaluated for one of our customers. The

components were subjected to screening and then to comprehensive characterization of parameters such as frequency, applied bias, temperature,

Low-Temperature Electronic Components Being Developed

Solid tantalum (darker) and ceramic capacitors. For a color version of this

photograph, see the online version of this document (http://www.grc.nasa.gov/

WWW/RT1998/5000/5480patterson.htm).

Capacitance of film and ceramic capacitors at

20-kHz versus temperature.

and multistress conditions. The

graph shows the capacitance of

film and ceramic capacitors as a

function of temperature.

These research and development

efforts are being carried out

through collaboration with other

government agencies, industrial

and aerospace companies, and

academia. The Low Temperature

Electronics Program supports

missions and development programs at the NASA Goddard Space

Flight Center and the Jet Propulsion

Laboratory.

To find out more, visit us on the

World Wide Web:

http://www.grc.nasa.gov/WWW/

epbranch/ephome.htm (select “Low

Temperature Electronics” at the bottom

of the list on the left)

Lewis contact:

Richard L. Patterson, (216) 433–8166,

Richard.L.Patterson@grc.nasa.gov

Authors:

Richard L. Patterson and

Ahmad Hammoud

Headquarters program office:

OSS (ATMS)

Programs/Projects: X2000 (JPL),

Next Generation Space Telescope

(Goddard), Low Temperature

Electronics

Change in capacitance, percent

20

0

10

–10

–20

–200 –150 –100

Temperature, °C

–50 0 50

Film

Ceramic
54

NASA Lewis Research Center’s Atomic Oxygen Lamp Cleaning Facility was

designed to produce an atomic oxygen plasma within a metal halide lamp

to remove carbon-based contamination. It is believed that these contaminants contribute to the high failure rate realized during the production of

these lamps.

The facility is designed to evacuate a metal halide lamp and produce a

radiofrequency-generated atomic oxygen plasma within it. Oxygen gas,

with a purity of 0.9999 percent and in the pressure range of 150 to

250 mtorr, is used in the lamp for plasma generation while the lamp is

being cleaned. After cleaning is complete, the lamp can be backfilled

with 0.9999-percent pure nitrogen and torch sealed.

The facility comprises various vacuum components connected to a

radiation-shielded box that encloses the bulb during operation. Radiofrequency power is applied to the two parallel plates of a capacitor, which

are on either side of the lamp. The vacuum pump used, a Leybold Trivac

Type D4B, has a pumping speed of 4-m3/hr, has an ultimate pressure of

<8310–4, and is specially adapted for pure oxygen service. The electronic

Atomic Oxygen Lamp Cleaning Facility Fabricated

and Tested

Front view of Atomic Oxygen Lamp Cleaning Facility.

power supply, matching network,

and controller (500-W, 13.56-MHz)

used to supply the radiofrequency

power were purchased from RF

Power Products Inc. Initial test

results revealed that this facility

could remove the carbon-based

contamination from within bulbs.

For more information, visit the

Electro-Physics Branch on the

World Wide Web:

http://www.grc.nasa.gov/WWW/

epbranch/ephome.htm

Lewis contacts:

Edward A. Sechkar, (216) 433–2299,

Edward.A.Sechkar@grc.nasa.gov; and

Thomas J. Stueber, (216) 433–2218,

Thomas.J.Stueber@grc.nasa.gov

Authors: Edward A. Sechkar and

Thomas J. Stueber

Headquarters program office:

OSS (ATMS)

Programs/Projects:

Broad applicability to lamp industry—

wide variety of glass bulb lamps for

lighting and specific wavelength

spectral source lamps; technique

would assure that hydrocarbon contaminants within the lamp enclosure are

not present during sealing and that the

lamp function and durability would

thereby be improved
55

Research and Technology

Stirling Technology Co., as part of a NASA Lewis Research Center Phase II

Small Business Innovation Research contract, has successfully demonstrated paralleling two thermodynamically independent Stirling converters.

A system of four Stirling converters is being developed by NASA and the

Department of Energy as an alternative high-efficiency radioisotope power

source for spacecraft onboard electric power for NASA deep space missions. The high Stirling efficiency, exceeding 20 percent for this application,

will greatly reduce the necessary isotope inventory in comparison to the

current radioisotope thermoelectric generators (RTG’s), significantly reducing mission cost and risk. Stirling is the most developed converter option

of the advanced power technologies under consideration.

Previous Stirling development had focused on single converters and had

not addressed how to connect multiple converters in a system. However,

in most potential space applications, multiple converters are important

for reliability and modularity. Thermodynamically independent converters

allow one converter to fail without affecting the performance of the other.

Finally, the use of multiple converters is important to controlling vibrations,

a critical issue for a dynamic space power system. Synchronization of converter pairs operating in an opposed configuration provides balanced

operation with minimal vibration.

Synchronization was achieved with two 350-We converters (see the

photograph) operating over a wide range of conditions, including simulated degraded operation for one of the two converters. This produced a

40- to 50-fold reduction in vibrations in comparison to an unbalanced converter. Equal power generation between the two converters was shown for

normal operation and simulated

degradation with constant power

input. Successful system operation

was demonstrated with the two

synchronized converters feeding a

battery charger load, as would

most likely be used in a radioisotope power system. Transient data

taken during various connections

and disconnections of the two converters showed that the converters

could be synchronized reliably and

rapidly. The two converters were

connected electrically in parallel

and mechanically through external

attachments on the cold-end

pressure vessels. A mechanical

coupler was developed that precisely aligns the two converters and

can compensate for any inherent

misalignments.

This connection method for multiple converters will now be utilized

to connect the prototype converters that are being developed by the

Department of Energy and Stirling

Technology Co. for potential use in

a 150-We Stirling isotope power

system.

For further background information,

visit the World Wide Web:

Research at Stirling Technology Co.:

http://www.stirlingtech.com

Research at Lewis:

http://powerweb.grc.nasa.gov/stirling/

home.html

Lewis contact:

Lanny G. Thieme, (216) 433–6119,

Lanny.G.Thieme@grc.nasa.gov

Author: Lanny G. Thieme

Headquarters program office: OSS

Programs/Projects:

SBIR, advanced radioisotope power

Parallel Stirling Converters Being Developed for

Spacecraft Onboard Power

Parallel Stirling converters synchronized for system operation and low vibrations.
56

Space-based solar thermal power systems are very appealing as a space

power source because they generate power efficiently. However, solar

thermal (dynamic) systems currently incorporate rigid concentrators that

are relatively heavy and require significant packaging volume and robust

deployment schemes. In many ways, these requirements make these

systems less appealing than photovoltaic systems. As an alternative to

solar thermal power systems with rigid concentrators, solar thermal power

systems with thin-film inflation-deployed concentrators have low cost, are

lightweight, and are efficiently packaged and deployed. Not only are inflatable concentrators suitable for low Earth orbit and geosynchronous orbit

applications, but they can be utilized in deep space missions to concentrate solar energy to high-efficiency solar cells.

The objective of this Small Business Innovation effort with SRS Technologies, Inc., was to design, build, demonstrate, and characterize a thin-film,

inflatable concentrator that has interfaces for solar thermal power systems.

The actual concentrator, which was delivered to the NASA Lewis Research

Center in late October 1998, is a 2.0- by 3.0-m off-axis parabolic concentrator with a 90° tilt angle and 30° half angle. The ultimate goal of this effort

is to demonstrate, through testing, that inflatable concentrators for power

generation are both a viable and desirable alternative to the rigid concentrator designs that are currently being used. Lewis has been working

closely with the U.S. Air Force, Edwards Air Force Base, to finalize contractual agreements and test requirements to support joint testing.

The objectives of the test are to (1) quantify the thermal performance and

optical performance of the test article in the simulated space environment

that is provided by

Lewis’ Tank 6 Solar

Simulator and (2)

use the test data to

qualify analytical

computer models.

This will be the first

in a series of several

planned space environmental tests that

Lewis and the Air

Force are planning as

part of a technology

road map for an

inflation-rigidized

concentrator flight

experiment that is

currently scheduled

for 2001. During

the test, thermal

performance will be

measured via thermocouples on the

Inflatable Solar Thermal Concentrator Delivered

surface of the inflated concentrator

and the support torus. Optical

performance will be characterized

via flux intensity measurements

at the focal plane via a chargecoupled discharge (CCD) camera.

Additional performance data will

be collected on the catenary strain,

torus and canopy pressure, and

make-up gas requirements.

Major activities for the Tank 6 test

include

(1) Setup of the test article and test

support equipment

(2) Integration of the instrumented

test article with Tank 6 data collection equipment and the Thiokol

inflation controller

(3) Calibration and verification of

data collection, evacuation of the

chamber, and thermal cycling and

data collection

(4) Repressurization of the chamber, removal and storage of test

articles, and restoration of Tank 6

to pretest status

(5) Data reduction

Lewis contact:

Carol M. Tolbert, (216) 433–6167,

Carol.M.Tolbert@grc.nasa.gov

Author: Carol M. Tolbert

Headquarters program office: OAT

Programs/Projects:

Launch Services and Space Transportation Program Office at Lewis

1.7- by 2.7-m inflatable concentrator designed and built

by SRS Technologies, Inc.
57

Research and Technology

∑

Microphone

Flame s(t)

m(t)

∑

Microphone

r(t)

n(t)

d

x(t) = s(t) + m(t)

y(t) = r(t) + n(t)

Internal gas temperature is one of the most fundamental parameters

related to engine efficiency and emissions production. The most common

methods for measuring gas temperature are physical probes, such as

thermocouples and thermistors, and optical methods, such as Coherent

Anti Stokes Raman Spectroscopy (CARS) or Rayleigh scattering. Probes are

relatively easy to use, but they are intrusive, their output must be corrected

for errors due to radiation and conduction, and their upper use temperature is limited. Optical methods are nonintrusive, and they measure some

intrinsic property of the gas that is directly related to its temperature (e.g.,

lifetime or the ratio of line strengths). However, optical methods are

usually difficult to use, and optical access is not always available. Lately,

acoustic techniques have been receiving some interest as a way to overcome these limitations (ref. 1).

One of these techniques, acoustic pyrometry, is also nonintrusive, is simpler to use than optical pyrometry, and doesn’t require optical access. To

measure the mean gas temperature between two points, one places

a sound (acoustic) source at one point and a microphone at a known distance from the source, then measures the time required for the signal to

travel from the source to the microphone. The speed of the sound is the

distance divided by the time of flight, and it is directly proportional to the

square root of the absolute temperature of the gas.

Measuring gas temperature in the noisy environment of a jet engine, however, might require a very powerful sound source. Instead, an experiment

was performed at the NASA Lewis Research Center in collaboration with

the University of Nevada at Reno (UNR), to determine the feasibility of

using the combustion noise itself as the sound source (ref. 2). The experiment consisted of installing two microphones downstream of the combustion zone of a burner rig, measuring the time of flight of the signal, and

calculating the gas temperature. The setup is shown schematically in the

accompanying figure. Since no signal was externally introduced at the

upstream location, the time of flight had to be determined from the cross

correlation of the signals from the two microphones. The upstream and

downstream signals were recorded on a two-channel digital audio recorder

and processed at the University of Nevada at Reno to calculate the cross

correlation between signals s(t) and r(t) and to remove the uncorrelated

noise n(t) and m(t), as shown in

the figure (where t indicates time).

The shape of the cross-correlation

function indicated that the signals

consisted of wideband random

noise. The measured speed of

sound for this particular experiment was 787 m/sec, which corresponds to 1270 °C.

Future work will include experiments to determine whether the

various operating conditions of

an engine have distinct acoustic

signatures, raising the possibility of

acoustic combustion monitoring as

well as temperature measurement.

References

1. Kleppe, J.A.: High Temperature Gas

Measurements Using Acoustic

Pyrometry. Sens., vol. 13, no. 1, Jan.

1996, pp. 17–22.

2. Kleppe, J.; Sanches, J.; Fralick, G.: The

Application of Acoustic Pyrometry to

Gas Turbines and Jet Engines. AIAA

Paper 98–3611, 1998.

Lewis contact:

Gustave C. Fralick, (216) 433–3645,

Gustave.C.Fralick@grc.nasa.gov

Author: Gustave C. Fralick

Headquarters program office: OAT

Programs/Projects:

Propulsion Systems R&T

Acoustic Pyrometry Applied to Gas Turbines and Jet Engines

Measurement of sound speed using two microphones.

Instrumentation and Controls
58

A flow-visualizing system with a scanning optical beam offers

greater advantages than the conventional approaches. In addition

to a higher signal-to-noise ratio and lower source power, the

scanning permits visualization of weak phenomena such as the

scattering and diffraction of light on shocks. Scanning beam

flow-visualization techniques were evaluated at the NASA Lewis

Research Center for shock position sensing. In an effort to eliminate moving parts, a spectral scanning technique was invented

and demonstrated.

This spectral scanner consists of a tunable laser, a diffraction

grating, and a lens. The tunable laser generates a narrow beam

of light whose wavelength changes in a prescribed manner. When

a light beam emitted by the laser interacts with the grating, the

grating causes the beam to change its direction, producing a

phenomenon called diffraction. The new direction depends on

the wavelength. This space-wavelength scanning generates a

“rainbow” in the time domain, where each “color” appears in its

place at a given time. Thus, if the wavelength λ changes monotonically in time, the diffracted beam draws a cone with the apex

being the point of impact of the beam on the grating. The lens

is positioned so that its focal point coincides with the point of

impact. It transforms the cone of light into a number of beams of

different wavelengths separated from each other in time and parallel to

each other in space. In other words, the lens converts an angular scanning

into a linear one. When a beam encounters a shock or another type of

inhomogeneity, it diffracts, splits, and forms secondary fringes and tails.

The drawing shows a blunt body inserted into a stream of air. The blunt

body generates the shock, and a charge-coupled discharge (CCD) camera

array behind the shock observes the phenomenon. The following graphs

demonstrate splitting a beam of light. At a certain wavelength (left graph)

the beam strikes the shock. At a different wavelength (right graph), the

light beam misses the shock and the beam shape, in this case, is not

affected. This information can be used to determine the location of a

shock wave.

Scanning Mode Shock Position Sensor Invented and

Demonstrated

Intensity profiles observed by a CCD array at two wavelengths during the spectral

scanning. Left: At the wavelength corresponding to the beam striking the shock.

Right: At the wavelength corresponding to the beam missing the shock.

Bibliography

Adamovsky, G.; and Johnson, D.K.:

Optical Techniques for Shock Visualization and Detection. Optical Techniques

in Fluid, Thermal, and Combustion

Flow. S.S. Cha and J.D. Trolinger, eds.,

SPIE Proceedings, Issue 2546, 1995,

pp. 348–357.

Adamovsky, G.: Scanning Mode Sensor

for Detection of Flow Inhomogeneities.

U.S. Patent 5,715,047, Feb. 1998.

Lewis contact:

Dr. Grigory Adamovsky, (216) 433-3736,

Grigory.Adamovsky@grc.nasa.gov

Author: Dr. Grigory Adamovsky

Headquarters program office: OAT

Programs/Projects:

Information Technology Base

Principle of operation of a spectral scanner.

Blunt body

l1, . . . , ln

Nozzle

Lens

Shock

Air

Tunable

laser

Diffraction 

grating

l1 ln

CCD array
59

Research and Technology

Aircraft icing occurs when a plane flies through a cloud of supercooled

water droplets. When the droplets impinge on aircraft components, ice

starts to form and accumulate. This accumulation of ice severely increases

the drag and lift of the aircraft, and can ultimately lead to catastrophic

failures and even loss of life. Knowledge of the air pressures on the surfaces of ice and models in wind tunnels allows researchers to better

predict the effects that different icing conditions will have on the performance of real aircraft. The use of pressure-sensitive paint (PSP) has provided valuable information on similar problems in conventional wind

tunnel testing.

In NASA Lewis Research Center’s Icing Research Tunnel, Lewis researchers

recently demonstrated the world’s first application of PSP on actual ice

formed on a wind tunnel model. This proof-of-concept test showed that

a new paint formulation developed under a grant by the University of

Washington adheres to both the ice shapes and cold aluminum models,

provides a uniform coating that preserves the detailed ice shape structure,

and responds to simulated pressure changes.

Three different samples of ice were acquired on a length of aerodynamic

tubing attached to the ceiling of the IRT. The samples consisted of glaze

ice (–0.7 °C, 30.7 °F), mixed ice (–8.1 °C, 17.4 °F) and rime ice (–15.4 °C,

4.2 °F). Prior to painting, the samples were removed from the wind tunnel

and stored in the facility cold room at a temperature of –6.7 °C (20 °F).

The PSP was cooled to the cold-room temperature prior to application.

Then PSP was applied to the samples using a automotive type spray gun,

with cooled nitrogen providing the atomization pressure. After the curing

time, which was less than 5 min with this paint, finish details were noted.

Pressure-Sensitive Paint Applied to Ice Accretions

Painted rime ice and the detailed ice structures

that remain after the application of the

pressure-sensitive paint coating.

False-color reduced data image of a

nitrogen jet impinging on the ice surface.

These images are shown in color in the

online version of this article (http://

www.grc.nasa.gov/WWW/RT1998/5000/

5520bencic.htm).

A low-pressure nitrogen jet was

used to determine the pressure

responsiveness of the paint. The

nitrogen locally displaces air and

simulates a region of low pressure. PSP works on the oxygen

quenching principle: in a vacuum,

no quenching occurs and the

intensity of the paint is at its maximum. At higher air pressures,

oxygen quenches the luminescence and the normalized intensity ratio behaves in a linear

fashion. The success of this test

lends optimism to the final application of PSP to icing research

wind tunnel testing.

Lewis contact:

Timothy J. Bencic, (216) 433–5690,

Timothy.J.Bencic@grc.nasa.gov

Author: Timothy J. Bencic

Headquarters program office: OAT

Programs/Projects:

Aeronautics Base R&T
60

Electronic time-average holograms are convenient for comparing the

measured vibration modes of fan blades with those calculated by finiteelement models. At the NASA Lewis Research Center, neural networks

recently were trained to perform what had been a simple visual comparison of the predictions of the design models with the measurements.

Finite-element models were used to train neural networks to recognize

damage and strain information encoded in subtle changes in the timeaverage patterns of cantilevers. But, the design-grade finite element

models were unable to train the neural networks to detect damage in

complex blade shapes. The design-model-generated patterns simply did

not agree well enough with the measured patterns. Instead, hybrid-training

records, with measured time-average patterns as the input and modelgenerated strain information as the output, were used to effect successful

training. One inspection process is outlined in the figure.

A twisted blade appears at the top left. The full time-average or characteristic pattern of the first vibration mode is shown next. The third and fourth

pictures at the top show measured-region time-average patterns for

Neural Networks Used to Compare Designed and Measured

Time-Average Patterns

Performance of a measured-input, model-output neural network. This figure is shown in color in the online version of this

article (http://www.grc.nasa.gov/WWW/RT1998/5000/5520decker.htm).

Measured region

Strain

patterns

Uncracked Cracked

Neural network processor

undamaged and cracked blades,

where a crack was induced by

high-cycle fatigue. These patterns

were sampled on a nonuniform

finite-element-model grid (not

shown). The neural networks

processed the samples as often as

30 times/sec. The outputs of the

neural networks, in this case, were

chordwise strains.

Three kinds of neural-net training

have been implemented with

software. These are listed in

increasing order of effectiveness.

(1) Neural networks can be model

trained with model-generated

time-average patterns and model-
61

Research and Technology

generated strain patterns. The effectiveness of this technique depends

strongly on the accuracy of the models.

(2) Neural networks can be trained with measured time-average patterns

and model-generated strain patterns.

(3) Neural networks can be trained very effectively to categorize measured

time-average patterns. Categories can consist of damaged and undamaged

fan blades, for example.

We plan to expand this work in the future from nonrotating to rotating fan

blades.

Bibliography

Decker, A.J., et al.: Comparison of Computational-Model and Experimental-Example

Trained Neural Networks for Processing Speckled Fringe Patterns. NASA TM—1998-

208814 (also International Conference on Optical Technology and Image Processing in Fluid, Thermal, and Combustion Flow, Proceedings of VSJ–SPIE98, Yokohama,

Japan Dec. 6–10, 1998, CD ROM format only), 1998.

Lewis contact:

Dr. Arthur J. Decker, (216) 433–3639,

Arthur.J.Decker@grc.nasa.gov

Authors:

Dr. Arthur J. Decker, E. Brian Fite, Oral

Mehmed, and Scott A. Thorp

Headquarters program office:

OAT, DDF

Program/Project: AST

Nondestructive Strain Measurement System Used to

Determine Surface Strain on Fibers

Sensor head optics of laser speckle-shift strain measurement

system illuminating test specimen.

Small-diameter structural fibers are being considered as

reinforcements for high-temperature ceramic matrix

composite materials, and thus they require characterization. At the NASA Lewis Research Center, a nondestructive optical technique was used to determine surface

strain on a structural fiber, in real time, as it was pulled

in a tensile test machine.

With this technique, interference or speckle patterns

from the laser illuminated fiber test specimen are

recorded. As the fiber is pulled, its speckle pattern shifts

in proportion to the strain, translation, and rotation components of the sample deformation. Shifting speckle

patterns are detected in real time by two linear chargecoupled discharge (CCD) camera arrays, and the images

are processed by a hardware correlator. Surface strain

is selectively detected on fibers with diameters on the

order of 100 μm and can be resolved to 19 microstrain.

This system was designed to be robust and compact

and generally does not require surface preparation of

the structural fibers. For strain detection, two laser

beams are positioned incident on the structural fiber

being tested, as shown in the photograph, where the

test specimen is mounted in a tensile test machine via

two coupons. As the fiber is pulled, the speckle pattern

produced from each laser beam is detected by one of

two CCD arrays located inside the tube on the right side

of the photograph.
62

Data obtained from three standard specimens (silica, silicon carbide, and

tungsten) using the real-time two-color laser speckle-shift system compared well with published values. A linear fit of the raw data was performed to determine the experimental modulus of each specimen (the

slope of the stress versus the strain curve). Errors between the known and

experimental values were 5.2 percent or less. Numerical results for silica,

silicon carbide, and tungsten fibers are given in the following table.

The key feature of this system, which separates it from current strain

detection systems, is the real-time data analysis. This allows for closed-loop

system control based on the detected strain value. Some other advantages

of this system follow:

(1) It is easy to use.

(2) Generally, it does not require preparation of specimen surface.

(3) It does not require post-processing.

(4) It will operate at elevated temperatures without modifications.

Because the strain of three standard structural fibers was measured successfully, this system can

be used with confidence to characterize unknown samples.

Find out more about this research on

the World Wide Web:

http://www.grc.nasa.gov/WWW/

OptInstr/stra.html

Lewis contacts:

Dr. Margaret L. Tuma, (216) 433–8665,

Margaret.L.Tuma@grc.nasa.gov; and

Lawrence G. Oberle, (216) 433–3657,

Lawrence.G.Oberle@grc.nasa.gov

Author: Dr. Margaret L. Tuma

Headquarters program office: OAT

Programs/Projects: HITEMP

High Reliability Engine Control Demonstrated for

Aircraft Engines

For a dual redundant-control system, which is typical for short-haul aircraft,

if a failure is detected in a control sensor, the engine control is transferred

to a safety mode and an advisory is issued for immediate maintenance

action to replace the failed sensor. The safety mode typically results in

severely degraded engine performance. The goal of the High Reliability

Engine Control (HREC) program was to demonstrate that the neuralnetwork-based sensor validation technology can safely operate an engine

by using the nominal closed-loop control during and after sensor failures.

With this technology, engine performance could be maintained, and the

sensor could be replaced as a conveniently scheduled maintenance

action.

The neural network architecture

used here for the sensor validation is

the Auto-Associative Neural Network

(AANN). This feed-forward network

architecture has output data that

reproduce the network input data

(see the figure to the right). AANN

consists of two layers, the mapping

layer and the demapping layer,

which are interconnected through

the bottleneck layer. The mapping

layer compresses the data into a

reduced-order representation,

eliminating redundancies and

extracting the key features (principal components) in the data.

Reducing the number of dimensions is the key characteristic of this

Mapping

layer layer

 Demapping

Bottleneck

layer

XNLc

XNHc

P25c

P3c

T3c

WFc

DPBLD

XNLc

XNHc

P25c

P3c

T3c

WFc

DPBLD

Input Estimates

Auto-associative neural network schematic (7–10–4–10–7 network) for turbofan

engine estimates.

COMPARISON OF EXPERIMENTAL STRAIN MEASUREMENTS TO ACCEPTED VALUES

Sample Fiber diameter,

μm

Accepted modulus,

GPa

Experimental modulus,

Gpa

Error,

percent

Silica 142 73.1 76.9 5.2

Silicon carbide 143 410.0 418.0 2.0

Tungsten 177 380.0 389.0 2.4
63

Research and Technology

Closed-loop control with sensor validation.

Outer Servoloops controller

Actuator

servos Engine

PLA

ALT

XM

P2, T2

WFMA

CVGMA

MMVFB,

XNL, XNH, ITT, T2P5

CVGFB

WF,

CVG

Full Authority Digital Engine

Controller (FADEC) Plant

Sensor

validation

architecture. The demapping layer recovers the encoded information from

the principal components. In addition, a fault detection logic identifies the

unique sensor failure based on the pattern of the error vector of the input

and output of the AANN.

In March 1998 the High Reliability Engine Control program successfully

demonstrated the neural-network-based sensor failure detection and

accommodation algorithm on a real-time simulation of the Allison’s

AE3007 engine and its Full Authority Digital Engine Controller (FADEC,

ref. 1). Successful accommodation of faults for low- and high-speed rotor

speed sensors was demonstrated in the closed-loop engine simulation

(see the following flow diagram).

Simulation results show that the neural-network-based sensor validation

algorithm can identify both hard and soft sensor failures and can estimate

how long the engine control will operate without degradation (ref. 2).

The High Reliability Engine Control

program is a collaborative effort

between NASA, the Allison Engine

Company, and Scientific Monitoring

Inc., a small business. In this program, NASA provided the sensor

failure detection technology which

included the neural network

modeling, training, and detection

algorithm. Allison Engine Company

provided the expertise in AE3007

engine model and control, and

Scientific Monitoring Inc. did the

system integration and implementation. This program gave these

research collaborators an opportunity to further develop and transfer

NASA-developed advanced technology for sensor failure detection,

isolation, and accommodation to

aircraft engine industry.

References

1. Mattern, D.L., et al.: Using Neural

Networks for Sensor Validation. AIAA

Paper 98–3547, 1998.

2. Allison Engine Co.: AE 3007 High

Reliability Engine Control Program.

Development/Demonstration of

Reliability Enhanced Control

Technologies to Improve Aircraft

Dispatch Reliability. NASA3–27394,

Task Order 008, EDR 18549, Final

Contractor Report, April 1998.

Lewis contact:

Dr. Ten-Huei Guo, (216) 433–3734,

Ten-Huei.Guo@grc.nasa.gov

Author: Dr. Ten-Huei Guo

Headquarters program office: OAT

Program/Project: HREC

Variable Engine parameter

DPBLD

P25

P3

T3

WF

XNL

XNH

Delta pressure between compressor discharge and bypass duct

Low-pressure compressor inlet pressure

Compressor discharge pressure, burner inlet

Temperature at burner inlet

Fuel flow

Fan rotor speed

Core rotor speed

Variable Engine parameter

ALT

CVGMA

CVGFB

ITT

MMVFB

PLA

P2

T2

T2P5

WFMA

XM

XNL

XNH

Aircraft altitude

Compressor variable geometry command

Compressor variable geometry feedback

Interstage turbine temperature

Main metering valve feedback signal

Power lever angle (thrust command)

Fan inlet pressure

Fan inlet temperature

Compressor inlet temperature

Main metering valve command

Aircraft Mach number

Fan rotor speed

Core rotor speed
64

Future aircraft turbine engines, both commercial and military, will have to

be able to successfully accommodate expected increased levels of steadystate and dynamic engine-face distortion. Advanced tactical aircraft are

likely to use thrust vectoring for enhanced aircraft maneuverability. As a

result, the engines will see more extreme distortion levels than currently

encountered with present-day aircraft. Also, the mixed-compression inlets

needed for the High-Speed Civil Transport (HSCT) will likely encounter

disturbances similar to those seen by tactical aircraft, in addition to planar

pulse, inlet buzz, and high distortion levels at low flight speed and offdesign operation.

The current approach of incorporating sufficient component design stall

margin to tolerate these expected levels of distortion would result in

significant performance penalties. The objectives of NASA’s High Stability

Engine Control (HISTEC) program, which has reached a highly successful

conclusion, were to design, develop, and flight demonstrate an advanced,

high-stability, integrated engine control system that uses measurementbased real-time estimates of distortion to enhance engine stability. The

resulting distortion tolerant control adjusts the stall margin requirement

online in real time. This reduces the design stall margin requirement, with

a corresponding increase in performance and decrease in fuel burn.

Distortion Tolerant Control Flight Demonstration

Shown to Be Successful

The HISTEC approach includes two

major systems: a Distortion Estimation System (DES) and Stability

Management Control (SMC). The

DES is an aircraft-mounted, highspeed processor that estimates

the amount and type of distortion

present and its effect on the

engine. It uses high-response

pressure measurements at the

engine face to calculate indicators

of the type and extent of distortion

in real time. From these indicators,

the DES determines the effects of

the distortion on the propulsion

system. In addition, the DES uses

maneuver information from the

flight control to anticipate high distortion conditions. The DES output

consists of fan and compressor

pressure ratio trim commands that

Distortion estimates

Engine sensitivities

Onboard stability audit

Stall margin control laws

Actuator

commands

Flight control

data

Trim

commands

Engine face

pressure 

sensors

Distortion tolerant control. (Engine cutaway copyright Pratt & Whitney; used with permission.)
65

Research and Technology

0 10 20 30 40 50

–0.12

–0.08

–0.04

0.0

0.04

0.08

0.12

Time, sec

Normalized engine pressure ratio (EPR)

Control EPR limit

Distortion EPR limit

In-flight fan distortion accommodation.

F–15 ACTIVE aircraft.

are passed to the SMC. The SMC performs a stability audit online by using

the trims from the DES and then accommodates the distortion through the

production engine actuators.

Last year, the HISTEC distortion tolerant control system was flight tested on

the NASA F–15 ACTIVE aircraft at the NASA Dryden Flight Research Center

in Edwards, California. The flight demonstration showed closed-loop control operation with the engine fan and compressor stall margins being

adjusted on the basis of estimated distortion. Project pilots flew the F–15

ACTIVE aircraft through a variety of maneuvers—such as high angle of

attack flight, windup turns, and takeoff—which create distorted airflow

conditions at the inlet. Detailed analysis of the flight test data performed

this year verified that both the DES and SMC performed as designed. The

system was able to successfully modify the engine pressure ratio (EPR)

limit in response to high levels of distortion, thus temporarily moving the

operating line to accommodate the lower stability condition.

The NASA Lewis Research Center

conducted the HISTEC program in

partnership with the NASA Dryden

Flight Research Center, Pratt &

Whitney, Boeing Phantom Works

(formerly McDonnell Douglas), and

the U.S. Air Force.

Find out more about the High

Stability Engine Control distortion

tolerant control technologies on the

World Wide Web:

http://www.grc.nasa.gov/WWW/cdtb/

projects/histec/

Bibliography

DeLaat, J.C.; Southwick, R.D.; and

Gallops, G.W.: High Stability Engine

Control (HISTEC). AIAA Paper 96–2586

(NASA TM–107272), 1996. (Available

online: http://gltrs.grc.nasa.gov/cgibin/GLTRS/browse.pl?1996/TM107272.html)

DeLaat, J.C., et al.: The High Stability

Engine Control (HISTEC) Program:

Flight Demonstration Phase. AIAA

Paper 98–3756 (NASA TM–1998–

208482), 1998. (Available online:

http://gltrs.grc.nasa.gov/cgi-bin/

GLTRS/browse.pl?1998/TM-1998-

208482.html)

Southwick, R.D., et al.: High Stability

Engine Control (HISTEC) Flight Test

Results. AIAA Paper 98–3757 (NASA

TM–1998–208481), 1998. (Available

online: http://gltrs.grc.nasa.gov/cgibin/GLTRS/browse.pl?1998/TM-1998-

208481.html)

Orme, J.S., et al.: Development and

Testing of a High Stability Engine

Control (HISTEC) System. AIAA Paper

98–3715, 1998.

DeLaat, J.C., et al.: High Stability Engine

Control (HISTEC): Flight Demonstration

Results. SAE Paper 985556 (NASA TM—

1998-208655), 1998.

Lewis contact:

Jonathan S. Litt, (216) 433–3748,

Jonathan.S.Litt@grc.nasa.gov

Author: Jonathan S. Litt

Headquarters program office: OAT

Program/Projects:

Propulsion Systems R&T, PHSV, HISTEC

Special recognition:

1998 NASA Group Achievement Award
66

NASA Lewis Research Center’s Advanced Controls and Dynamics Technology Branch is investigating active control strategies to mitigate or eliminate

the combustion instabilities prevalent in lean-burning, low-emission combustors. These instabilities result from coupling between the heat-release

mechanisms of the burning process and the acoustic flow field of the

combustor.

Control design and implementation require a simulation capability that is

both fast and accurate. It must capture the essential physics of the system,

Combustion Instabilities Modeled

Results from a sample simulation. Top: Combustor diagram. Middle:

Time-averaged pressure, temperature, velocity, and heat-release rate.

Bottom: Contours of pressure and velocity showing self-sustained,

coupled, thermoacoustic oscillations.

yet be as simple as possible. A

quasi-one-dimensional, computational fluid dynamics (CFD) based

simulation has been developed

which may meet these requirements. The Euler equations of

mass, momentum, and energy

have been used, along with a

single reactive species transport

equation to simulate coupled

thermoacoustic oscillations. A

very simple numerical integration

scheme was chosen to reduce

computing time. Robust boundary

condition procedures were incorporated to simulate various flow

conditions (e.g., valves, open ends,

and choked inflow) as well as to

accommodate flow reversals that

may arise during large flow-field

oscillations.

The accompanying figure shows a

sample simulation result. A combustor with an open inlet, a choked

outlet, and a large constriction

approximately two thirds of the

way down the length is shown.

The middle plot shows normalized,

time-averaged distributions of the

relevant flow quantities, and the

bottom plot illustrates the acoustic mode shape of the resulting

thermoacoustic oscillation. For this

simulation, the limit cycle peak-topeak pressure fluctuations were

13 percent of the mean.

The simulation used 100 numerical

cells. The total normalized simulation time was 50 units (approximately 15 oscillations), which took

26 sec on a Sun Ultra2.

Lewis contact:

Dr. Daniel E. Paxson, (216) 433–8334,

Daniel.E.Paxson@grc.nasa.gov

Author: Dr. Daniel E. Paxson

Headquarters program office: OAT

Program/Project: HSCT

0.0 0.2 0.4 0.6 0.8 1.0

0.0

1.0

2.0

3.0

4.0

Ratio of axial distance to length, x/L

Normalized q uantities

Constant-pressure inlet Choked outlet

Reactants Products

Heat release rate

Temperature

Pressure

Velocity

Pressure

45

40

35

Nondimensional time

0.0 0.5

x/L

1.0

Velocity

45

40

35

Nondimensional time

0.0 0.5

x/L

1.0
67

Research and Technology

Extensive research conducted by NASA Lewis Research Center’s Satellite

Networks and Architectures Branch and the Ohio University has demonstrated performance improvements in World Wide Web transfers over

satellite-based networks. The use of a new version of the Hypertext Transfer Protocol (HTTP) reduced the time required to load web pages over a

single Transmission Control Protocol (TCP) connection traversing a satellite

channel. However, an older technique of simultaneously making multiple

requests of a given server has been shown to provide even faster transfer

time (refs. 1 to 3).

Unfortunately, the use of multiple simultaneous requests has been shown

to be harmful to the network in general. Therefore, we are developing new

mechanisms for the HTTP protocol which may allow a single request at any

given time to perform as well as, or better than, multiple simultaneous

requests (ref. 4). In the course of study, we also demonstrated that the

time for web pages to load is at least as short via a satellite link as it is via

a standard 28.8-kbps dialup modem channel (ref. 5). This demonstrates

that satellites are a viable means of accessing the Internet.

References

1. Kruse, H., et al.: HTTP Page Transfer Rates Over Geo-Stationary Satellite Links.

Proceedings of the Sixth International Conference on Telecommunication

Systems, 1998, p. 518. (Available online: http://roland.grc.nasa.gov/~mallman/

papers/nash98.ps)

2. Kruse, H., et al.: Performance Analysis of the HTTP Protocol on Geostationary

Satellite Links. Satellite Networks: Architectures, Applications, and Technologies

Workshop. NASA CP—1998-208524, 1998, pp. 65–74. (Available online: http://

www.csm.ohiou.edu/kruse/Presentations/Cleve98/index.htm)

3. Kruse, H., et al.: A Comprehensive Evaluation of HTTP Over Satellite Channels.

Submitted to ACM Mobile Networks and Applications Journal (Dec. 1998).

Web Transfer Over Satellites Being Improved

Communications Technology

4. Kruse, H.; Allman, M.; and Mallasch,

P.: Network and User-Perceived

Performance of Web Page Retrievals.

Proceedings of the First International

Conference on Telecommunications

and Electronic Commerce (ICTEC),

Nov. 1998. (Available online: http://

roland.grc.nasa.gov/~mallman/

papers/ecom98.ps)

5. Griner, J., et al.: Internet Over a

Bi-Directional Satellite Link. Satellite

Networks: Architectures, Applications, and Technologies Workshop.

NASA CP—1998-208524, 1998,

pp. 49–50. (Available online: http://

roland.grc.nasa.gov/~jgriner/

papers/workshop-demo.pdf )

Find out more about this HTTP study

on the World Wide Web:

http://ctd.grc.nasa.gov/5610/

httpextensions.html

Lewis contact:

Mark Allman, (216) 433–6586,

Mark.A.Allman@grc.nasa.gov

Author: Mark Allman

Headquarters program office: OSS

Programs/Projects:

Aeronautics Base R&T

Because of the unique networking characteristics of communications

satellites, global satellite networks are moving to the forefront in enhancing

national and global information infrastructures. Simultaneously, broadband

data services, which are emerging as the major market driver for future

satellite and terrestrial networks, are being widely acknowledged as the

foundation for an efficient global information infrastructure. In the past

2 years, various task forces and working groups around the globe have

identified pivotal topics and key issues to address if we are to realize such

Advances Made in the Next Generation of Satellite

Networks

networks in a timely fashion. In

response, industry, government,

and academia undertook efforts to

address these topics and issues.

A workshop was organized to provide a forum to assess the current

state-of-the-art, identify key issues,
68

Space-based networks scenario.

and highlight the emerging trends in the next-generation architectures,

data protocol development, communication interoperability, and applications. The Satellite Networks: Architectures, Applications, and Technologies

Workshop was hosted by the Space Communication Program at the NASA

Lewis Research Center in Cleveland, Ohio. Nearly 300 executives and

technical experts from academia, industry, and government, representing

the United States and eight other countries, attended the event (June 2 to

4, 1998). The program included seven panels and invited sessions and

nine breakout sessions in which 42 speakers presented on technical topics.

The proceedings (ref. 1) covers a

wide range of topics: access technology and protocols, architectures

and network simulations, asynchronous transfer mode (ATM) over

satellite networks, Internet over

satellite networks, interoperability

experiments and applications,

multicasting, NASA interoperability

experiment programs, NASA mission applications, and Transmission

Control Protocol/Internet Protocol

(TCP/IP) over satellite: issues, relevance, and experience.

Reference

1. Bhasin, K.: Satellite Networks:

Architectures, Applications, and

Technologies, NASA/CP—1998-

208524, 1998.

Lewis contact:

Dr. Kul B. Bhasin, (216) 433–3676,

Kul.B.Bhasin@grc.nasa.gov

Author: Dr. Kul B. Bhasin

Headquarters program office: OSS

Programs/Projects: NII/GII

Extensive research conducted by the Satellite Networks and Architectures

Branch of the NASA Lewis Research Center led to an experimental change

to the Internet’s Transmission Control Protocol (TCP) that will increase

performance over satellite channels. The change raises the size of the initial

burst of data TCP can send from 1 packet to 4 packets or roughly 4 kilobytes (kB), whichever is less.

TCP is used daily by everyone on the Internet for e-mail and World Wide

Web access, as well as other services. TCP is one of the feature protocols

used in computer communications for reliable data delivery and file

transfer. Increasing TCP’s initial data burst from the previously specified

single segment to approximately 4 kB may improve data transfer rates by

up to 27 percent for very small files. This is significant because most file

transfers in wide-area networks today are small files, 4 kilobytes or less. In

addition, because data transfers over geostationary satellites can take 5 to

20 times longer than over typical terrestrial connections, increasing the

Internet Protocol Enhanced Over Satellite Networks

initial burst of data that can be sent

is extremely important.

This research along with research

from other institutions has led to

the release of two new Request for

Comments (RFC, ref. 1) from the

Internet Engineering Task Force

(IETF, the international body that

sets Internet standards). In addition, two studies of the implications of this mechanism were also

funded by NASA Lewis (refs. 2 and

3). Additional information is available in reference 4.
69

Research and Technology

Internet Protocol Testbed.

Find out more about this research on the World Wide Web:

http://info.internet.isi.edu:80/in-notes/rfc/files/rfc2414.txt

http://info.internet.isi.edu:80/in-notes/rfc/files/rfc2416.txt

http://sulu.grc.nasa.gov/5610/inetprotocols.html

References

1. Allman, M.; Floyd, S.; and Partridge, C.: Increasing TCP’s Initial Window. RFC 2414,

Experimental, Sept. 1998. (Available online: http://info.internet.isi.edu:80/

in-notes/rfc/files/rfc2414.txt)

2. Shepard, T.; and Partridge, C.: When TCP Starts Up With Four Packets Into Only

Three Buffers. RFC 2416, Informational, BBN Technologies, Sept. 1998. (Online

information: http://info.internet.isi.edu:80/in-notes/rfc/files/rfc2416.txt)

3. Allman, M.; Hayes, C.; and Ostermann, S.: An Evaluation of TCP With Larger Initial

Windows. ACM SIGCOMM, vol. 28, no. 3, 1998, pp. 41–52.

Lewis contact:

William D. Ivancic, (216) 433–3494,

William.D.Ivancic@grc.nasa.gov

Author: William D. Ivancic

Headquarters program office:

OSS, OAT

Programs/Projects:

Level 1 HEDS 22 improved groundbased data networks, 64 advanced

data handling; ESE 119 advanced endto-end mission information system

technologies; Level 2 3694 desktop

instrument operations end-to-end

protocol standards, 3700 and 2030

high-data-rate communications, 3753

communications protocols; SOMO

Interoperability and Standardization 8,

9, 10; Innovative Mission Information

Systems 56; and Hybrid Network

Ubiquity 61, 62, 63

An important element of satellite-supported asynchronous transfer mode

(ATM) networking will involve support for the routing and rerouting of

active connections. Work published under the auspices of the Telecommunications Industry Association (http://www.tiaonline.org), describes basic

architectures and routing protocol issues for satellite ATM (SATATM) networks (ref. 1). The architectures and issues identified will serve as a basis

for further development of technical specifications for these SATATM

networks.

Three ATM network architectures for bent pipe satellites and three ATM

network architectures for satellites with onboard ATM switches were

developed. The architectures differ from one another in terms of required

level of mobility, supported data rates, supported terrestrial interfaces,

and onboard processing and switching requirements. The documentation

(ref. 1) addresses low-, middle-, and geosynchronous-Earth-orbit satellite

configurations.

Satellite ATM Networks: Architectures and Guidelines

Developed

The satellite environment may

require real-time routing to support

the mobility of end devices and

nodes of the ATM network itself.

This requires the network to be

able to reroute active circuits in real

time. In addition to supporting

mobility, rerouting can also be used

to (1) optimize network routing,

(2) respond to changing quality-ofservice requirements, and (3) provide a fault tolerance mechanism.

Reference 1 examines the various

reroute events and the time-scale

in which they would occur.
70

Traffic management and control functions are necessary in ATM to ensure

that the quality-of-service requirements associated with each connection

are not violated and also to provide flow and congestion control functions.

Functions related to traffic management were identified and described.

Most of these traffic management functions will be supported by onground ATM switches, but in a hybrid terrestrial-satellite ATM network,

some of the traffic management functions may have to be supported by

the onboard satellite ATM switch. It is expected that the ATM traffic loading

on the satellite will be high; hence, it has been proposed to transfer traffic

management functions (to the degree possible) to Earth station components. This will also reduce the complexity of the satellite hardware. Future

work is planned to examine the tradeoffs of placing traffic management

functions onboard a satellite as opposed to implementing those functions

at the Earth station components. This work was done at the NASA Lewis

Research Center in collaboration with ComSat Labs.

Find out more about this research on the World Wide Web:

http://ctd.grc.nasa.gov/5610/5610.html

References

1. Satellite ATM Networks: Architectures

and Guidelines, TSB91, 1998, Telecommunications Industry Association.

Lewis contact:

Thomas C. vonDeak, (216) 433–3277,

fax (216) 433–8705,

Thomas.C.Vondeak@grc.nasa.gov

Authors:

Thomas C. vonDeak and Ferit Yegendu

Headquarters program office: OSS

Programs/Projects: ATM research and

development, SOMO, satellite communications industry

The current usage of broadband (data rate greater than 64 kilobits per

second (kbs)) for multimedia network computer applications is increasing,

and the need for network communications technologies and systems to

support this use is also growing. Satellite technology will likely be an

important part of the National Information Infrastructure (NII) and the

Global Information Infrastructure (GII) in the next decade. Several candidate communications technologies that may be used to carry a portion of

the increased data traffic have been reviewed, and estimates of the future

demand for satellite capacity have been made. A study was conducted by

the NASA Lewis Research Center to assess the satellite addressable markets for broadband applications. This study effort included four specific

milestones: (1) assess the changing nature of broadband applications and

their usage, (2) assess broadband satellite and terrestrial technologies, (3)

estimate the size of the global satellite addressable market from 2000 to

2010, and (4) identify how the impact of future technology developments

could increase the utility of satellite-based transport to serve this market.

Broadband Satellite Technologies and Markets Assessed

Lewis’ approach was to assess the

requirements for satellite broadband services, which are based on

an understanding of user perspectives and data needs. This research

effort included interviews with

leading industry players and technology experts to understand the

unfolding markets for broadband

services and the status of terrestrial network deployments. In addition, the suitability of available and

emerging terrestrial broadband

technologies to compete with the

newly licensed Ka-band satellite

systems in providing multimedia

Estimated penetration of broadband into top five regional satellite markets and less developed regions. (Adapted from

Booz-Allen & Hamilton analysis, ref. 1.)

40

30

20

10

0

2000 2002 2004

Year

2006 2008 2010

Subscribers, milions

6

4

5

3

2

1

0

2000 2002 2004

Year

2006 2008 2010

Subscribers, milions
71

Research and Technology

services to business and consumer markets was studied. The analysis

was supported by an end-to-end communications services supply-anddemand model used to make quantitative assessments. It considered

relevant technical, demographic, and competitive factors including application transport requirements, user preferences, changing usage levels

and application mixes, affordability by user segments and geographies,

investment costs and delays for upgrades, and expansion of the terrestrial

infrastructure.

We estimate that North America will have the highest number of satellite

service subscribers as shown in the graph. South America is projected to be

the second largest regional opportunity. In this model, the underdeveloped

terrestrial networks in the region and rapidly growing gross domestic product (GDP) drive this result. Western European subscriptions could peak by

2007 and slowly decline as terrestrial buildup catches up with demand in

areas not served by terrestrial technologies. In less developed regions,

overall satellite service penetration is likely to be significantly lower initially

than for the top five regions but to grow consistently over time. Of these,

Russia is projected to be the largest single market, with over 5 million subscribers, followed by Eastern Europe, with over 3 million. We believe this

result reflects the potentially underreported level of investment in terrestrial infrastructure for Russia as reported by the International Telecommunications Union (ITU). Nonetheless, Russia will most likely remain a large

market for satellite services considering its large geography and low population density. A set of recommendations to improve the utility of satellite

systems for delivering broadband services was developed. We expect that

the methodology developed in this

study will allow more quantitative

assessments of specific broadband

satellite technologies and markets.

References

1. Booz-Allen & Hamilton: Assessment

of Broadband Satellite Technologies

and Markets, Volume 2. NASA CR—

1998-206595, 1998. Available from

Lewis’ Satellite Networks and

Architectures Branch.

2. Booz-Allen & Hamilton: Assessment

of Broadband Satellite Technologies

and Markets, Volume 1—Executive

Summary. NASA CR—1998-206596,

1998. Available from Lewis’ Satellite

Networks and Architectures Branch.

Lewis contact:

Dr. Thomas M. Wallett, (216) 433–

3673, Thomas.M.Wallett@grc.nasa.gov

Author: Dr. Thomas M. Wallett

Headquarters program office: OSS

Programs/Projects: Space Communications, communications industry

The phenomenal growth of commercial communications has created a

great demand for traveling-wave tube (TWT) amplifiers. Although the helix

slow-wave circuit remains the mainstay of the TWT industry because of its

exceptionally wide bandwidth, until recently it has been impossible to

accurately analyze a helical TWT using its exact dimensions because of the

complexity of its geometrical structure. For the first time, an accurate threedimensional helical model was developed that allows accurate prediction

of TWT cold-test characteristics including operating frequency, interaction

impedance, and attenuation. This computational model, which was developed at the NASA Lewis Research Center, allows TWT designers to obtain

a more accurate value of interaction impedance than is possible using

experimental methods.

Obtaining helical slow-wave circuit interaction impedance is an important

part of the design process for a TWT because it is related to the gain and

efficiency of the tube. This impedance cannot be measured directly; thus,

conventional methods involve perturbing a helical circuit with a cylindrical

dielectric rod placed on the central axis of the circuit and obtaining the

A More Accurate and Efficient Technique to Obtain Helical

Traveling-Wave Tube Interaction Impedance Using

Computational Methods

difference in resonant frequency

between the perturbed and

unperturbed circuits. A mathematical relationship has been derived

between this frequency difference

and the interaction impedance

(ref. 1). However, because of the

complex configuration of the

helical circuit, deriving this relationship involves several approximations. In addition, this experimental

procedure is time-consuming and

expensive, but until recently it was

widely accepted as the most

accurate means of determining

interaction impedance.
72

100

80

60

40

20

0

0 10 20 30 40 50

Impedance, 

V

Frequency, GHz

Measurement by perturbation

MAFIA, perturbation

MAFIA, direct calculation

References

1. Lagerstrom, R.P.: InteractionImpedance Measurements by

Perturbation of Traveling Waves.

Stanford Electronics Laboratories,

Technical Report No. 7, Naval

Research Contract, Nonr 225 (24),

NR 373 360, Stanford University,

Stanford, CA, Feb. 11, 1957.

2. Kory, C.L.; and Dayton, J.A., Jr.:

Accurate Cold-Test Model of Helical

TWT Slow-Wave Circuits. IEEE Trans.

Electron Devices, vol. 45, no. 4,

1998, pp. 966–971.

3. Weiland, T.: On the Numerical

Solution of Maxwell’s Equations and

Applications in the Field of Accelerator Physics. Part. Accel., vol. 15,

1984, pp. 245–292.

4. Weiland, T.: On the Unique Numerical Solution of Maxwellian Eigenvalue Problems in Three Dimensions. Part. Accel., vol. 17, 1985,

pp. 227–242.

5. Kory, C.L.; and Dayton, J.A., Jr.:

Computational Investigation of

Experimental Interaction Impedance

Obtained by Perturbation for Helical

Traveling-Wave Tube Structures.

IEEE Trans. Electron Devices, vol. 45,

no. 9, 1998, pp. 2063–2071.

Lewis contact:

Carol L. Kory, (216) 433–3512,

Carol.L.Kory@grc.nasa.gov

Author: Carol L. Kory

Headquarters program office: OSS

Programs/Projects:

CETDP; high-efficiency, high-data-rate

communications systems for future

missions

The advent of an accurate three-dimensional helical circuit model (ref. 2)

made it possible for Lewis researchers to fully investigate standard approximations made in deriving the relationship between measured perturbation

data and interaction impedance. The most prominent approximations

made in the analysis were addressed and fully investigated for their

accuracy by using the three-dimensional electromagnetic simulation code

MAFIA (Solution of Maxwell’s Equations by the Finite Integration Algorithm) (refs. 3 and 4). We found that several approximations introduced

significant error (ref. 5).

To further validate the three-dimensional computational helical model,

researchers duplicated the experimental perturbation method by simulating the helical circuit with a cylindrical dielectric rod of size and material properties consistent with the experimental setup. The difference in

frequency between the perturbed and unperturbed circuits was obtained.

Then, the interaction impedance was calculated using the previously

mentioned approximate formulation relating frequency difference to

interaction impedance. The following graph, which compares the results

with measured values, emphasizes the accuracy of the code. The interaction impedance was calculated with the exact formula using MAFIA and is

also plotted in this graph. The results calculated directly with MAFIA are

consistently lower than measured results, with an average difference of

26.6 percent.

The demonstrated inaccuracy of approximations in the derived experimental impedance formula, along with the large discrepancy between these

measured impedance data and direct calculations using MAFIA, verifies that

a more accurate value of interaction impedance can be obtained by using

three-dimensional computational methods. In addition, this implies that

MAFIA will yield large savings in time and cost in comparison to expensive

and time-consuming experimental cold-test measurements.

On-axis interaction impedance obtained by measurement and by using

MAFIA with the perturbation method and direct calculation.
73

Research and Technology

Diamond is a promising semiconductor material for novel electronic

applications because of its chemical stability and inertness, heat conduction properties, and so-called negative electron affinity (NEA). When a

surface has NEA, electrons generated inside the bulk of the material are

able to come out into the vacuum without any potential barrier (work

function). Such a material would have an extremely high secondary electron emission coefficient σ, very high photoelectron (quantum) yield, and

would probably be an efficient field emitter.

Chemical-vapor-deposited (CVD) polycrystalline diamond films have even

more advantages than diamond single crystals. Their fabrication is relatively

easy and inexpensive, and they can be grown with high levels of doping—

consequently, they can have relatively high conductivity. Because of these

properties, diamond can be used for cold cathodes and photocathodes in

high-power electronics and in high-frequency and high-temperature

semiconductor devices.

Using mostly secondary electron spectroscopy as a tool, the NASA

Lewis Research Center conducts research on the electronic properties of

both natural diamond crystals and microwave-assisted chemical-vapordeposited thin diamond films. Whereas artificial diamond films can be

made conductive, the effects occurring in natural diamond can be

obscured by the undissipated charge. Because of this effect, measurements on natural diamond often have to be done in the pulse mode. The

preceding block diagram shows the installation for measuring the secondary electron emission coefficient σ in such a mode, and the left graph on

Diamond Analyzed by Secondary Electron Emission

Spectroscopy

Experimental setup (b, battery; C, capacitor; R, resistor). Top: Block diagram. Bottom:

Timing diagram of various voltages supplied into the top drawing.

Wehnelt

Cathode Target

b1 b2 C1

Electronic

polarity

switch

Current-tovoltage

converter

Amplifier

Sample and hold

Output 1

Output 2

C

R

the next page shows measurements of σ for the (100) surface of

IIb natural diamond single crystals, both saturated with hydrogen

and dehydrogenated. For the dehydrogenated surface, the secondary

electron yield is quite low, especially in comparison to the surface

saturated with hydrogen, which has

a very high secondary electron yield

because of its NEA status. Hydrogen can be desorbed if the surface

is heated at elevated temperatures

or is exposed to an electron beam.

To improve the stability of the

surface under an electron beam,

Lewis researchers recently studied

various coatings that might preserve the NEA property of diamond

surfaces. Among the materials

tested for the coatings were alkali

halides, such as cesium iodide

(CsI). Tests showed that as a result

of the exposure to the electron

beam, the halogen (I in this case)

desorbed into the vacuum, leaving
74

Secondary electron emission coefficient σ for (100)

surface of natural IIb diamond crystal.

Surface concentrations versus fluence for diamond

coated with 100 nm of CsI.

80

40

60

20

0

0.0 1.0 2.0 3.0

Primary electron energy, E, eV

4.0 5.0x103

s(E)

Hydrogenated

surface

Dehydrogenated

surface

80

40

60

20

0

0 50 100 150

Fluence, C/cm2

200 250 300

Surface concentration, percent

Iodine

Carbon

Cesium

the surface coated with a thin layer of active ingredient. The resulting a

surface was stable under the electron beam exposure and had a very high

secondary electron yield, indicating that it was exhibiting NEA. The graph

on the right shows an example of this effect for a thin CsI film on diamond.

It can be seen that the iodine concentration at the surface diminishes with

the exposure to the electron beam.

Lewis contact:

Dr. Isay L. Krainsky, (216) 433–3509,

Isay.L.Krainsky@grc.nasa.gov

Author: Dr. Isay L. Krainsky

Headquarters program office: OSS

Programs/Projects:

Electron beam devices

Researchers in NASA Lewis Research Center’s Electron Device Technology

Branch are developing transmission lines for radiofrequency and wireless circuits that are more efficient, smaller, and make lower cost circuits

possible. Traditionally, radiofrequency and wireless circuits have employed

a microstrip or coplanar waveguide to interconnect the various electrical

elements that comprise a circuit. Although a coplanar waveguide (CPW)

is widely viewed as better than a microstrip for most applications, it too

has problems. To solve these problems, NASA Lewis and the University of

Finite Ground Coplanar (FGC) Waveguide: Characteristics

and Advantages Evaluated for Radiofrequency and Wireless

Communication Circuits

Michigan developed a new version

of a coplanar waveguide with

electrically narrow ground planes.

This new transmission line, which

we call the finite ground coplanar

(FGC) waveguide, is illustrated in

the following figure.

Coplanar waveguides. Waveguide width, W; center conductor width, S; ground plane width, B. Left:

Conventional coplanar waveguide. Center: Waveguide with lower ground plane. Right: Finite ground

coplanar waveguide.
75

Research and Technology

1.5

1.4

1.3

1.2

1.1

1.0

0.0 1.0 2.0 3.0 4.0

Normalized attenuation,

aFGC/aCPW

B' = B/SFGC

Conformal mapping

Measured data

B

W

S

W

B

LC

LC

Attenuation, α, of FGC waveguide with S = W = 25 μm

normalized to the attenuation of a conventional CPW

with S = W = 25 μm as a function of the normalized

ground plane width, B, where S is the center conductor

width and W is the waveguide width.

MIM capacitor. Center conductor width, S; waveguide

width, W; ground plane width, B; capacitor length, LC .

Top: Integrated in ground planes. Bottom: Integrated in

center conductor of the FGC waveguide.

Through extensive numerical modeling and experimental measurements,

we have characterized the propagation constant of the FGC waveguide, the

lumped and distributed circuit elements integrated in the FGC waveguide,

and the coupling between parallel transmission lines. Although the attenuation per unit length is higher for the FGC waveguide because of higher

conductor loss, the attenuation is comparable when the ground plane

width is twice the center conductor width as shown in the preceding

graph. An upper limit to the line width is derived from observations that

when the total line width is greater than λd/2, spurious resonances due

to the parallel plate waveguide mode are established. Thus, the ground

plane width must be less than λd/4 where λd is the wavelength in the

dielectric. Since the center conductor width S is typically less than λ/10 to

maintain good transverse electromagnetic mode characteristics, it follows

that a ground plane width of B = 2S would also be electrically narrow.

Thus, we can now treat the ground strips of the FGC waveguide the same

way that the center conductor is treated.

This allows the novel integration of

circuit elements in the FGC waveguide. To explore this, researchers

integrated metal-insulator-metal

(MIM) capacitors, inductors, and

thin film resistors into the center

conductor and the ground planes

of the transmission line as illustrated in the final figure. Results

indicate that even though the value

of the parasitic reactance associated with each element is independent of the element placement,

the value of the primary element

value (capacitance, inductance, and

resistance, respectively) does vary

upon placement. Specifically, when

two elements are connected in the

ground planes as shown in the

drawing, they add in parallel. Thus,

capacitors in the ground planes can

be made shorter and still provide

the same capacitance as capacitors

placed in the center conductor.

Furthermore, since the parasitic

reactance is independent of element placement, the self-resonant

frequency of the capacitors in the

ground planes is higher.

Lastly, coupling is reduced as the

ground plane width is reduced

while the distance between the

center of each line is kept constant.

Furthermore, the coupling between

parallel transmission lines is lower

for the FGC waveguide than for the

conventional coplanar waveguide.

In fact, even a small slot between

the ground planes of the conventional coplanar waveguide can

lower coupling by 10 dB.

All these results indicate that the

FGC waveguide is a better transmission line than the coplanar

waveguide and the microstrip for

radiofrequency and wireless

circuits. Further work is underway

to more fully develop design rules

for this new transmission line.
76

Bibliography

Ponchak, G.E.; Tentzeris, E.M.; and Katehi, L.P.B.: Characterization of Finite Ground

Coplanar Waveguide With Narrow Ground Planes. Int. J. Microcircuits Electronic

Packaging, vol. 20, no. 2, 1997, pp. 167–173.

Ponchak, G.E., et al.: Finite Width Coplanar Waveguide for Microwave and Millimeter-Wave Integrated Circuits. Proceedings of the International Symposium on Microelectronics, ISHM ’96, The Microelectronics Society, Reston, VA, 1996, pp. 517–521.

Ponchak, G.E.; and Katehi, L.P.B.: Characteristics of Finite Ground Coplanar Waveguide Lumped Elements. Proceedings of the IEEE MTT–S International Microwave

Symposium Digest, Vol. 2, 1997, pp. 1003–1006.

Ponchak, G.E.; and Katehi, L.P.B.: Open- and Short-Circuit Terminated Series Stubs

in Finite-Width Coplanar Waveguide on Silicon. IEEE Trans. Microwave Theory Tech.,

vol. 45, no. 6, June 1997, pp. 970–976.

Ponchak, G.E.: Development of Passive Components for Millimeter-Wave Circuits.

Ph.D. Thesis, Univ. of Michigan, 1997.

Ponchak, G.E.; Tentzeris, E.M.; and Katehi, L.P.B.: Coupling Between Adjacent Finite

Ground Coplanar (FGC) Waveguides. International Symposium on Microelectronics,

1997, pp. 7–10.

Lewis contact:

Dr. George E. Ponchak, (216) 433–3504,

fax (216) 433–8705,

George.E.Ponchak@grc.nasa.gov

Author:

Dr. George E. Ponchak

Headquarters program office: OSS

Programs/Projects: Space communications for NASA missions

Special recognition:

IMAPS’ 1997 International Symposium

on Microelectronics Best Paper Award

Target

Manipulator

Pumping system

Electron

beam Channeltron

Analyzer

It has been demonstrated that the suppression of secondary electron

emission significantly improves the performance of electron beam collectors (ref. 1). However, a complete analysis of the effects of secondary

electron emission with respect to collector performance has not been

possible because of the lack of quantitative data on angular distributions

Angular Distribution of Elastically Scattered Electrons

Determined and Its Effect on Collector Performance

Computed

Apparatus for determining the angular

distribution of secondary electrons.

of secondary electrons. Secondary

electrons are emitted with energies

ranging from near zero to the

energy of the incident primary. For

our purposes, we define elastically

scattered electrons as secondary

electrons within 20 percent of the

incident energy. Elastically scattered electrons are of great concern

because their energy allows them

to follow trajectories that can carry

them almost anywhere within

the vacuum envelope. If these secondaries leave the collector and

reenter the slow wave circuit,

they can produce undesired signal

distortion and oscillation.

This apparatus, which was built by

Krainsky (ref. 2), was used at the

NASA Lewis Research Center to

obtain detailed measurements of

the angular distributions of elastically scattered secondaries. Data

were obtained for three surfaces
77

Research and Technology

180

120

60

0

6.35310–5 units per meter

6.35

310–5 units per meter

0 60 120 180 240 300 360 420

of significant interest to collector applications: highly polished copper,

copper roughened by ion sputtering, and isotropic graphite. Lewis researchers discovered that elastically scattered electrons have a complex angular

distribution that is strongly dependent on the atomic number and surface

morphology of the target material, as well as the energy and angle of

incidence of the primary beam. At low energies, secondary emission from

polished copper in the chosen energy range is primarily directed back

to the source of primary electrons (backscattering). Forward scattering

increases with primary energy until, at high energies, forward scattering

dominates the angular distribution. Although back-scattered secondaries

dominate the distributions of the textured copper surface, the yield is

substantially lower. From the standpoint of secondary emission, isotropic

graphite is the most attractive material because it exhibits low yield and

little back scattering.

The measured data were curve-fitted into linear combinations of Gaussian

and Lorentzian functions, which were used in an interpolation routine to

calculate the approximate distributions at any arbitrary primary energy and

angle of incidence. The complexity of the data did not provide easy incorporation into the computational model. Furthermore, the volume of the

data exceeded computational resources and software limits. However,

improvements in computational power and software enhancements now

allow the inclusion of the newly available data into previous models of

electron beam collectors. Simulations on a collector designed for a 32-GHz

traveling-wave tube (TWT) showed that forward-scattered electrons had

Computer model of back-scattered electrons from one impact site in the

collector design for the Hughes 32-GHz traveling-wave tube proof-of-concept

design for the Cassini Mission.

little effect on collector performance. The angular distributions

for ion-textured copper dominated

by back-scattered secondaries were

of more concern because those

electrons are likely to reenter the

slow wave circuit. However, our

model of these distributions

showed that approximately 1 percent of the primary beam current

returned to the slow wave circuit,

which is in agreement with experimental observations. The figure to

the left shows a sample of such

a model. These additions have

brought even more accuracy and

insight to previous modeling

attempts.

References

1. Ramins, P.; and Ebihara, B.T.:

Improvements in MDC and TWT

Overall Efficiency Through Application of Carbon Electrode Surfaces.

IEEE Trans. Electron Devices, vol.

ED–33, no. 11, Nov. 1986, pp.

1915–1924.

2. Krainsky, I., et al.: The Angular

Distribution of Elastically Scattered

Electrons and Computed Impact on

Collector Performance. IEEE IEDM

92–953, 1992, pp. 37.4.1–37.4.4.

Lewis contacts:

Karl R. Vaden, (216) 433–8131,

Karl.R.Vaden@grc.nasa.gov; and

Dr. Isay L. Krainsky, (216) 433–3509,

Isay.L.Krainsky@grc.nasa.gov

Authors:

Dr. Isay L. Krainsky and Karl R. Vaden

Headquarters program office: OSS

Programs/Projects:

Traveling-wave tubes
78

Electronically tunable microwave components have become the subject of

intense research efforts in recent years. Many new communications systems would greatly benefit from these components. For example, planned

low-Earth-orbiting satellite networks have a need for electronically scanned

antennas. Thin ferroelectric films are one of the major technologies competing to fill these applications. When a direct-current (dc) voltage is

applied to ferroelectric film, the dielectric constant of the film can be

decreased by nearly an order of magnitude, changing the high-frequency

wavelength in the microwave device. Recent advances in film growth have

demonstrated high-quality ferroelectric thin films. This technology may

allow microwave devices that have very low power and are compact,

lightweight, simple, robust, planar, voltage tunable, and affordable.

The NASA Lewis Research Center has been designing, fabricating, and

testing proof-of-concept tunable microwave devices. This work, which is

being done in-house with funding from the Lewis Director’s Discretionary

Fund, is focusing on introducing better microwave designs to utilize these

materials. We have demonstrated Ku- and K-band phase shifters, tunable

local oscillators, tunable filters, and tunable diplexers. Many of our devices

employ SrTiO3 as the ferroelectric. Although it is one of the more tunable

and easily grown ferroelectrics, SrTiO3 must be used at cryogenic temperatures, usually below 100 K. At these temperatures, we frequently use hightemperature superconducting thin films of YBa2Cu3O7-δ to carry the microwave signals. However, much of our recent work has concentrated on

inserting room-temperature ferroelectric thin films, such as BaxSr1-xTiO3

into these devices. The BaxSr1-xTiO3 films are used in conjuction with

normal metal conductors, such as gold.

The first example of ferroelectric-based components is the tunable filter,

which has a compact, planar design. The metallic layer consists of a thin

YBa2Cu3O7-δ superconducting film. A 300-nm-thick SrTiO3 ferroelectric

film lies below the metallic layer on a LaAlO3 dielectric substrate. The

Thin-Film Ferroelectric Tunable Microwave Devices

Being Developed

S-parameter transmission and

reflection measurements in the

graph at the bottom left indicate

that filter’s center frequency shifted

from 16.5 GHz at no bias, to 18.8

GHz at ±400 V bias, showing a

tuning range of 12 percent at 30 K.

The passband losses are relatively

flat, and the minimum embedded

insertion loss is 1.5 dB. These

results demonstrate the feasibility

of using ferroelectric thin-film

planar microstrip filters at Ku-band

frequencies. Recent activities

include fabricating and testing

these tunable filters with ambient

temperature ferroelectrics (e.g.,

BaxSr1-xTiO3).

Extensive work has gone into

developing these novel ferroelectric

coupled microstrip phase shifters,

which were invented at Lewis. Two

configurations of these phase

shifters are given in the schematic

at the top of the next page. The

ferroelectric thin film again lies

below the metallic circuit on a

crystal substrate. Direct-current bias

is applied between the coupled

microstrips, strongly affecting the

dielectric constant of the ferroelectric in the gap and causing

a microwave phase shift. Larger

phase shifts are obtained by

placing several of these coupled

microstrip sections in series. At

cryogenic temperatures, these

devices have shown nearly 500° of

phase shift with figures of merit of

80° of phase shift per 1 dB of loss.

These results are shown in the final

graphs for a YBa2Cu3O7-δ /SrTiO3/

LaAlO3 multilayer structure.

On-going efforts are extending and

modifying these designs to different ferroelectric films, which are

being doped to reduce losses, and

to different dielectric substrates. Bias dependence of transmission signal, S21 , and

reflectance signal, S11 , measurements for a tunable

bandpass filter at 30 K. Magnitude of 

S21 and 

S11, dB

Frequency, GHz

0

–40

–30

–20

–10

14 15 16 17 18 19 20 21

S21= 

±400 V

S21= 

±100 V

S21= 

0 V

0

±100

±400

S11,

V
79

Research and Technology

Phase shifts of 50-Ω, eight-element, YBa2Cu3O7-δ

 (0.35 μm)/SrTiO3 (1.0 μm)/LaAlO3 (254 μm) coupled microstrip phase

shifter. Left: Temperature, 40 K; frequency, 16 GHz. Right: Temperature, 77 K; frequency, 16 GHz.

400 200 0 200 400

–10

–8

–6

–4

–2

Bias voltage, V

Magnitude of 

S21, dB

0

100

200

300

400

500

Phase of 

S21, deg

Magnitude

Phase

400 200 0 200 400

–10

–8

–6

–4

–2

Bias voltage, V

Magnitude of 

S21, dB

0

100

200

300

400

Phase of 

S21, deg

Magnitude

Phase

In addition, proof-of-concept tunable local oscillators have been assembled

from thin-film ferroelectric ring resonators combined with gallium arsenide

PHEMT’s (pseudomorphic high electron mobility transistors). These devices

have shown apparent unloaded quality factors up to 15 000 and frequency

tuning ranges of 2 GHz in the Ku-band. In addition, two tunable filters have

been linked to form a tunable diplexer operating at the Ku- and K-bands.

The tunable local oscillator and tunable diplexer are intended to be integrated into a discriminator-locked tunable oscillator currently under development at Lewis.

Bibliography

Van Keuls, F.W., et al.: (YBa2Cu3O7-δ,

Au)/SrTiO3/LaAlO3 Thin Film Conductor/Ferroelectric Coupled Microstripline Phase Shifters for Phased Array

Applications. Appl. Phys. Lett., vol. 71,

Nov. 1997, p. 3075.

Subramanyam, G.; Van Keuls, F.; and

Miranda, F.A.: A K-Band Tunable Microstrip Bandpass Filter Using a Thin Film

Conductor/Ferroelectric/Dielectric

Multilayer Configuration. IEEE Microwave Guided Wave Letters, vol. 8, no. 2,

1998, pp. 78–80.

Subramanyam, G.; Van Keuls, F.; and

Miranda, F.A.: Novel K-Band Tunable

Microstrip Bandpass Filter Using a Thin

Film HTS/Ferroelectric/Dielectric Multilayer Configuration. IEEE MTT–International Microwave Symposium Digest,

vol. 2 (NASA TM—1998-207940. Available online: http://gltrs.grc.nasa.gov/

cgi-bin/GLTRS/browse.pl?1998/

TM-1998-207940.html), 1998, pp.

1011–1014.

Romanofsky, R.R.; Van Keuls, F.W.; and

Miranda, F.A.: Cryogenic GaAs PHEMT/

Ferroelectric Ku-Band Tunable Oscillator. J. de Physique IV, vol. 8, no. 3

(NASA TM—1998-206967. Available

online: http://gltrs.grc.nasa.gov/cgibin/GLTRS/browse.pl?1998/TM-1998-

206967.html), 1998, pp. 171–174.

Schematic of two coupled microstrip phase shifters.

All dimensions are in micrometers. Top: 25-Ω,

single-element phase shifter; S = 12.7 μm, W =

76.2 μm. Bottom: 50-Ω, eight-element phase

shifter; S = 7.2 μm, W = 25 μm.

124

124

S

W

W

660

737

10 000

470

S

84 211 414

W
80

A traveling-wave tube (TWT) is an electron beam device that is used to

amplify electromagnetic communication waves at radio and microwave

frequencies. TWT’s are critical components in deep-space probes, geosynchronous communication satellites, and high-power radar systems.

Power efficiency is of paramount importance for TWT’s employed in deepspace probes and communications satellites. Consequently, increasing the

power efficiency of TWT’s has been the primary goal of the TWT group at

the NASA Lewis Research Center over the last 25 years.

Technique Developed for Optimizing Traveling-Wave Tubes

Traveling-wave tube.

An in-house effort produced a

technique (ref. 1) to design TWT’s

for optimized power efficiency. This

technique is based on simulated

annealing, which has an advantage

over conventional optimization

techniques in that it enables the

best possible solution to be

obtained (ref. 2). A simulated

Radiofrequency 

input

Heater

Focus

electrode

Cathode

Radiofrequency 

output

Multistage

depressed

collector

Magnetic focusing

Anode

Slow-wave circuit

Electron beam

Electron gun

Miranda, F.A., et al: Tunable Microwave Components for Ku- and K-Band Satellite

Communications. Integrated Ferroelectronics, vol. 22, no. 789 (NASA TM—1998-

206963. Available online: http://gltrs.grc.nasa.gov/cgi-bin/GLTRS/browse.pl?1998/

TM-1998-206963.html), 1998.

Van Keuls, F.W., et al.: Influence of the Biasing Scheme on the Performance of Au/

SrTiO3/LaAlO3 Thin Film Conductor/Ferroelectric Tunable Ring Resonators. Integrated Ferroelectrics, vol. 22, no. 883 (NASA TM—1998-206641. Available online:

http://gltrs.grc.nasa.gov/cgi-bin/GLTRS/browse.pl?1998/TM-1998-206641.html),

1998.

Van Keuls, F.W.; Romanofsky, R.R.; and Miranda, F.A.: Several Microstrip-Based

Conductor/ Thin Film Ferroelectric Phase Shifter Designs Using (YBa2Cu3O7-δ,Au)/

SrTiO3/LaAlO3 Structures. Integrated Ferroelectrics, vol. 22, no. 893 (NASA TM—

1998-206964. Available online: http://gltrs.grc.nasa.gov/cgi-bin/GLTRS/

browse.pl?1998/TM-1998-206964.html), 1998.

Van Keuls, F.W., et al.: A Ku-Band Gold/BaxSr1-xTiO3/LaAlO3 Conductor/Thin Film

Ferroelectric Microstripline Phase Shifter for Room Temperature Communications

Applications. Microwave and Optical Technology Letters, vol. 20, no. 1, 1999,

pp. 53–56.

Lewis contacts: Dr. Frederick

W. Van Keuls, (216) 433–9704,

Frederick.W.Vankeuls@grc.nasa.gov;

Robert R. Romanofsky,

(216) 433–3507,

Robert.R.Romanofsky@grc.nasa.gov;

and Dr. Félix A. Miranda, (787) 850–

9381, FA_MIRANDA@cuhac.upr.clu.edu,

or (216) 433–6589,

Felix.A.Miranda@grc.nasa.gov

Author:

Dr. Frederick W. Van Keuls

Headquarters program office:

OSS, DDF

Programs/Projects:

Low-cost tracking ground terminals,

low-phase noise receivers
81

Research and Technology

annealing algorithm was created and integrated into the NASA TWT

computer model (ref. 3). The new technique almost doubled the computed conversion power efficiency of a TWT from 7.1 to 13.5 percent (ref. 1).

The algorithm can be readily modified so that any calculable TWT characteristic can be optimized with respect to any combination of input design

parameters. This has enabled TWT’s to be optimally designed for important characteristics for which there were previously no design optimization

methodologies. For example, in one version, the technique has been used

to design a TWT for optimized overall power efficiency, which is extremely

important for TWT’s used on deep-space probes and communications

satellites. This design incorporates both the conversion power efficiency

from the electron beam to the microwave signal and the energy collected

from the “spent” beam in the multistage-depressed collector (see the

illustration). In another version, the technique was used to optimize the

power over a wide frequency bandwidth instead of at just a single frequency. This will substantially increase the data transmission volume

capability of TWT’s. In yet another version, the technique was used to

design a TWT with both high conversion power efficiency and low signal

distortion.

References

1. Wilson, J.D.: Simulated Annealing Algorithm for Optimizing RF Power Efficiency

in Coupled-Cavity Traveling-Wave Tubes. IEEE Trans. Electron Devices, vol. 44,

no. 12, Dec. 1997, pp. 2295–2299.

2. Kirkpatrick, S.: Optimization by Simulated Annealing. Science, 1983, pp. 671–680.

3. Wilson, J.D.: Revised NASA Axially Symmetric Ring Model for Coupled-Cavity

Traveling-Wave Tubes. NASA TP–2675, 1987.

Lewis contact:

Dr. Jeffrey D. Wilson, (216) 433–3513,

Jeffrey.D.Wilson@grc.nasa.gov

Author: Dr. Jeffrey D. Wilson

Headquarters program office: OSS

Programs/Projects:

Deep Space Network Upgrade.

Special recognition:

One of ten finalists for 1997 Rappaport

Prize for the best paper published in

IEEE Transactions on Electron Devices

The Cassini mission to Saturn is an international venture with participation

from NASA, the European Space Agency, and the Italian Space Agency. The

Cassini spacecraft was launched from Cape Canaveral in October 1997 and

is scheduled to arrive at Saturn in July 2004. After arrival, the spacecraft will

orbit Saturn about 60 times over a period of 4 years. During this time, the

Cassini Radio Science Subsystem will be used to investigate the atmosphere and rings of Saturn and the atmosphere of its largest moon, Titan—

which is larger than Mercury and is the only moon in our solar system with

a dense atmosphere.

A critical component in Cassini’s Radio Science Subsystem is a travelingwave tube (TWT) that was designed at the NASA Lewis Research Center

and built by Hughes Electronic Dynamics Division (ref. 1). This TWT will

amplify downlink microwave signals at a frequency of 32 GHz for the Deep

Space Network and will be involved in a number of experiments. These

include occultation experiments in which the microwave signal will be

beamed through rings and atmospheres toward Earth. Researchers will

analyze the received signals to determine the sizes and distributions of

the particles in the rings and the structure and composition of the atmosCassini Radio Science Experiments on Saturn and Titan

Preserved Because of Lewis Analysis

pheres. The Radio Science Subsystem also will also be used to

more accurately determine the

mass and size of Saturn and its

moons, to investigate the solar

corona, and to search for gravity

waves from outside the solar

system.

During tests of the TWT in 1995,

the Jet Propulsion Laboratory (JPL)

discovered an unexpected anomalous signal at a frequency of

35 GHz. Curiously, this signal only

occurred at power levels below

that at which the TWT was to be

operated. Mission planners were

very concerned that a small change

in operating conditions might
82

cause the spurious signal to occur at higher power levels and interfere

with detection of the desired signal at 32 GHz. Because the Jet Propulsion

Laboratory was unable to determine the cause of the anomalous signal

experimentally, an in-house effort was initiated at Lewis to try to determine the cause of the anomalous signal with Lewis-developed computer

analysis techniques.

Lewis’ extensive computer analysis determined that the anomalous signal

is an intermodulation product of the 32-GHz signal with a 67-GHz oscillation (ref. 2). The oscillation is induced by coupling to a second harmonic

beam current, which is significantly high only at low power levels. The

analysis indicated that possible changes in operating conditions would

not cause the anomaly to be a problem. The analysis convinced the Jet

Propulsion Laboratory to not remove the Radio Science Subsystem from

the Cassini spacecraft, and the experiments will proceed as originally

planned.

References

1. Curren, A.N., et al.: The Cassini

Mission Ka-Band TWT. 1994 IEEE

IEDM Technical Digest, 1994,

pp. 783–786.

2. Dayton, J.; Wilson, J.; and Kory, C.:

Computer Analysis of the Spectrum

Anomaly in the 32-GHz TWT for the

Cassini Mission. NASA TP—1999-

208691.

Lewis contact:

Dr. Jeffrey D. Wilson, (216) 433–3513,

Jeffrey.D.Wilson@grc.nasa.gov

Author: Dr. Jeffrey D. Wilson

Headquarters program office: OSS

Programs/Projects:

Cassini Mission to Saturn

A low cost, small size and mass, low heater power, durable high-performance barium dispenser thermionic cathode has been developed that

offers significant advancements in the design, manufacture, and performance of the electron sources used in vacuum electronic devices—such as

microwave (and millimeter wave) traveling-wave tubes (TWT’s)—and in

display devices such as high-brightness, high-resolution cathode ray tubes

(CRT’s). The lower cathode heater power and the reduced size and mass

of the new cathode are expected to be especially beneficial in TWT’s for

deep space communications, where future missions are requiring smaller

spacecraft, higher data transfer rates (higher frequencies and radiofrequency output power), and greater electrical efficiency. Also expected to

benefit are TWT’s for commercial and government communication satellites, for both low and geosynchronous Earth orbit, with additional benefits

offered by lower cost and potentially higher cathode current loading.

A particularly important TWT application is in the microwave power module (MPM), which is a hybrid microwave (or millimeter wave) amplifier

consisting of a low-noise solid-state driver, a vacuum power booster (small

TWT), and an electronic power conditioner integrated into a single compact package. The attributes of compactness and potentially high electrical

efficiency make the MPM very attractive for many commercial and government (civilian and defense) applications in communication and radar

systems. The MPM is already finding application in defense electronic

systems and is under development by NASA for deep space communications. However, for the MPM to become competitive and commercially

successful, a major reduction in cost must be achieved.

Novel Low-Cost, Low-Power Miniature Thermionic Cathode

Developed for Microwave/Millimeter Wave Tube and

Cathode Ray Tube Applications

The new cathode is expected to

make an important contribution to

lowering the cost of MPM’s and, in

addition, offer significant improvements in the electrical efficiency

and thermal management, as well

as potential reductions in size and

mass. Barium dispenser cathodes,

with their higher emission current

density capability and greater

durability, offer brightness, resolution, and lifetimes superior to the

alkaline-earth oxide-coated cathodes most commonly used in

CRT’s. Immediate applications for

CRT’s are in high-end monitors

(such as those for medical imaging,

computer-aided design (CAD), and

air traffic control), in CRT-based

projection displays, and in highdefinition television receivers.

FDE Associates of Beaverton,

Oregon, developed the miniature

barium dispenser cathode under

a Small Business Innovation
83

Research and Technology

Research (SBIR) contract with NASA Lewis Research Center. The design

and fabrication are based on practices used in the production of CRT’s,

which is one of the most competitive and efficient manufacturing operations in the world today. The approach used in the design and manufacture

of thermionic cathodes and electron guns for CRT’s has been optimized for

fully automated production, standardization of parts, and minimization of

costs. In addition, this approach offers a number of potentially significant

benefits for the production of similar components in microwave tubes,

among which are low cost, reductions in size and mass, and significantly

lower cathode heater power. Although CRT’s employ primarily oxide-coated

cathodes, CRT manufacturing technology is equally applicable to barium

dispenser cathodes.

FDE is presently using the CRT approach to develop and produce a number

of cathode sizes ranging from 0.050- to 0.0250-in. in diameter, with corresponding heater powers ranging from <1.0 to <3.5 W, respectively. In

addition to miniature TWT’s, the 0.050-in. cathode is applicable to highresolution, high-performance CRT’s that require a high brightness electron

source. Low cathode heater power results from a highly efficient thermal

design and a novel process for thinning and polishing the cathode support

sleeve to reduce heat conduction and radiation losses. The barium source

is a 4:1:1-impregnated, die pressed, and sintered tungsten alloy pellet for

long life and enhanced electron emission ( >2 A/cm2), although other

compositions, such as the osmiumcoated tungsten pellet are just as

feasible. Other basic design features include cylindrical mounting

(for ruggedness), ceramic disk

mounting (standardized drop-in

mounting into electron guns), a

slip-in cataphoretic-coated heater

(low cost and automated production), a heater enclosure separate

from the main support sleeve (for

power efficiency and enhanced

heater life), a nearly all-welded

design with limited brazing (automatic assembly and low cost), and

the use of simple shapes, drawn or

stamped parts and parts joined by

mechanical capture (crimping).

Emphasis is on standardization of

parts and the design of tooling and

assembly procedures for automated production. Flexibility in

manufacturing allows easy substitution of other cathode sizes.

Find out more on the World Wide

Web: http://www.fdeassc.com

Lewis contact:

Edwin G. Wintucky, (216) 433–3510,

Edwin.G.Wintucky@grc.nasa.gov

FDE Assoc. contact:

Bernard K. Vancil, (503) 628–0703,

Bernie@fdeassc.com

Author: Edwin G. Wintucky

Headquarters program office: OSS

Programs/Projects:

Space Communications

Support sleeve and ceramic disk mounting for 0.065-in.

cathode scale.
84

Future rapid deployment of low- and medium-Earth-orbit satellite constellations that will offer various narrow- to wide-band wireless communications services will require phased-array antennas that feature wide-angle

and superagile electronic steering of one or more antenna beams. Antennas, which employ monolithic microwave integrated circuits (MMIC), are

perfectly suited for this application. Under a cooperative agreement, an

MMIC-based, K-band phased-array antenna is being developed with 50/50

cost sharing by the NASA Lewis Research Center and Raytheon Systems

Company. The transmitting array, which will operate at 19 gigahertz (GHz),

is a state-of-the-art design that features dual, independent, electronically

steerable beam operation (±42°), a stand-alone thermal management, and

a high-density tile architecture. This array can transmit 622 megabits per

second (Mbps) in each beam from Earth orbit to small Earth terminals.

The weight of the total array package is expected to be less than 8 lb. The

tile integration technology (flip chip MMIC tile) chosen for this project

K-Band Phased Array Developed for Low-Earth-Orbit

Satellite Communications

Manifold

clamping

ring

Radiofrequency

cable (2)

Patch antenna

elements

(256)

Quad transmit MMIC

subassemblies (64)

Direct-current power/logic

(control) commands

cable (4)

Manifold assembly

Manifold mounting

ring

K-band phased-array antenna assembly. (Copyright Raytheon TI Systems; used with permission.)

represents a major advancement

in phased-array engineering and

holds much promise for reducing

manufacturing costs.

The transmit phased-array antenna

has passed the critical design

review and is being fabricated, with

delivery anticipated in late 1999.

Soon after the array is completed,

Lewis plans to integrate it into a

flight experiment to test its operation in the space environment. The

array will be a critical component

of the Direct Data Distribution (D3)
85

Research and Technology

Assembled K-band phased-array antenna. (Copyright Raytheon TI Systems; used

with permission.)

flight experiment on a future space

shuttle mission, as early as 2001.

Incorporation and proper design of

significant mechanical, thermal,

and control interfaces required for

safe and successful operation on

a shuttle mission will provide

significant leverage for using the

antenna in future communications

spacecraft. Phased-array antennas

that are capable of rapid direct

downlinking of large volumes of

data from various space platforms,

such as the one being developed

and described here, are high on

the priority list.

For more information, visit Lewis’

Communications Technology Division

on the World Wide Web:

http://ctd.grc.nasa.gov

Lewis contact:

Godfrey Anzic, (216) 433–3570,

Godfrey.Anzic@grc.nasa.gov

Author: Godfrey Anzic

Headquarters program office: OSS

Programs/Projects:

Space Communications, D3
86

Conformal, transparent printed antennas have advantages over conventional antennas in terms of space reuse and aesthetics. Because of their

compactness and thin profile, these antennas can be mounted on video

displays for efficient integration in communication systems such as palmtop computers, digital telephones, and flat-panel television displays. As

an array of multiple elements, the antenna subsystem may save weight by

reusing space (via vertical stacking) on photovoltaic arrays or on Earthfacing sensors. Also, the antenna could go unnoticed on automobile windshields or building windows, enabling satellite uplinks and downlinks or

other emerging high-frequency communications.

Conformal, Transparent Printed Antenna Developed for

Communication and Navigation Systems

Transparent, electromagnetically coupled ring-slot optically transparent conducting (OTC) antenna.

Top: Shown placed on top of a U.S. stamp. Bottom: Attachment details and performance data.

CD-97-75353

Transmitted

 signal, dB –20.0

–30.0

20.0 21.5 23.0

–10.0

0.0

Frequency, F, GHz

Return loss

Magnitude, dB

–80.0

90 120

–60.0

–40.0

–20.0

150 180 210 240 270

Angle, deg

H-Plane radiation pattern

F = 20.55 GHz

Transparent conductor

Slot line ring

Glass substrate

Transparent

microstrip feed

Recently, a class of transparent

patch and slot antennas consisting

of an ultrathin film of electrically

conductive material deposited on

glass or plastic substrates has been

developed jointly by the NASA

Lewis Research Center and the

Federal Data Corporation. A prototype antenna has been demonstrated with the antenna fabricated
87

Research and Technology

Transparent, rectangular microstrip optically transparent conducting (OTC) patch antenna.

Top: Shown placed on top of a U.S. stamp. Bottom: Attachment details and performance data. Reflected signal,

S11, dB

23.0

Frequency, F, GHz

Return loss

Magnitude, dB

90 120 150 180 210 240 270

Angle, deg

H-Plane radiation pattern

CD-97-75352

Transparent patch

Glass substrate

Transparent

ground plane

–40.0

–20.0

0.0

–80.0

–60.0

–40.0

–20.0

19.0 20.0

F = 20.7 GHz

on either glass or plastic substrates. The multistep liftoff lithographic technique is used to fabricate antennas are on glass substrates. Two of the

tested transparent antennas—a slot ring and a rectangular patch antenna

are shown along with their performance data in the figures. The antennas

have demonstrated very broadband characteristics, good impedance

matching, and radiation patterns for frequencies ranging from 2 to 30

gigahertz (GHz). We believe that these antennas can be further developed

to operate in the Federal Communication Commission’s (FCC) newly

opened, high-frequency bands above 30 GHz. Finally, a 2-by-2 array has

been fabricated and is currently being characterized.

Lewis contact:

Dr. Richard Q. Lee, (216) 433–3489,

Richard.Q.Lee@grc.nasa.gov

Authors: Dr. Richard Q. Lee and

Dr. Rainee N. Simons

Headquarters program office:

OA (DDF)

Programs/Projects: Deep space and

commercial wireless communications
88

Electro-optic, polymer-based integrated optic devices for high-speed

communication and computing applications offer potentially significant

advantages over conventional inorganic electro-optic crystals. One key

area of integrated optical technology—primary processing and fabrication—

may particularly benefit from the use of polymer materials. However, as

efforts concentrate on the miniaturization of electro-integrated circuit

pattern geometries, the ability to etch fine features and smoothly sloped

sidewalls is essential to make polymers useful for electro-integrated circuit

applications. There are many existing processes available to etch polymer

materials, but they all yield nearly vertical sidewalls. Vertical sidewalls are

too difficult to reliably cover with a metal layer, and incomplete metalization degrades microwave performance, particularly at high frequency.

However, obtaining a very sloped sidewall greatly improves the deposition

of metal on the sidewall, leading to low-loss characteristics, which are

essential to integrating these devices in high-speed electro-optic modulators.

The NASA Lewis Research Center has developed in-house an inexpensive

etching technique that uses a photolithography method followed by a

simple, wet chemical etching process to etch through polymer layers. In

addition to being simpler and inexpensive, this process can be used to

fabricate smoothly sloped sidewalls by using a commercial nonerodible

mask: Spin-On-Glass. A commercially transparent material, SpinOn-Glass, uses processes and equipment similar to that for photoresist

techniques.

Innovative, Inexpensive Etching Technique Developed for

Polymer Electro-Optical Structures

This new etching technique offers

a simpler fabrication process and

a cost-effective method of generating the moderately sized patterns

typically encountered in low and

medium density interconnection

applications. In addition, it yields

high-resolution patterns and

smooth sidewalls, resulting in low

optical scattering losses. With this

technique, measured microwave

losses have been found to be very

low. The new technique offers lowloss performance comparable to

that found with conventional

reactive ion etching techniques,

which have considerably higher

costs.

Lewis contact:

Hung D. Nguyen, (216) 443–6590,

Hung.D.Nguyen@grc.nasa.gov

Author: Hung D. Nguyen

Headquarters Program Office:

OSS (ATMS)

Program/Projects:

Space Communications Project

NASA Lewis Research Center’s Communications Technology Division has

an ongoing program in the development of efficient channel coding

schemes for satellite communications applications. Through a university

grant, as a part of this research, the University of Toledo is investigating

the performance of turbocodes, which use parallel concatenation of

nonsystematic convolutional encoders with an interleaver. The errorcorrecting capacity of these codes is close to the Shannon limit. The

research emphasis is on the development of low-complexity, but higher

rate (greater than one half), turbocodes and on the iterative decoding of

block codes.

Results of extensive computer simulations of higher rate turbocodes

(2/3, 4/5, 5/6, 10/11, and 15/16) show that their performance is greatly

affected by the proper selection of code polynomials, the puncturing

pattern, and the interleaver selection. In addition, some types of turboHigh-Performance, Low-Complexity Codes Researched for

Communication Channels

codes tend to have flares or

flooring in the energy-to-noise

versus bit-error-rate (Eb/No versus

BER) curve. Research shows that,

for the same code polynomials

selected, puncturing with the

extended puncturing period

exhibits significant coding gains

since more diverse positions are

selected. The graphs show this gain

for a 5/6 turbocode. Investigations

of flares (or flooring) show that

there are two important parameters related to the interleaver

design. The first is the degree of
89

Research and Technology

uniformity. A nonuniform (pseudorandom) interleaver improves

performance because it can make

the weight distribution of a code

similar to that of a random code

that has a performance close to the

Shannon limit. The other aspect is

the interleaver size. Results show

that higher rate turbocodes with

a large size random interleaver

do not exhibit flare. Additional

research showed that an interleaver block size of 64 by 64 is optimum for a bit error rate of 10–6.

Another part of the research is on

the application of the “turbo”

concept to the block codes. The

method of iterative decoding based

on the Maximum a Posteriori Probability (MAP) algorithm, can be

used to improve the performance

of a simple block code to close to

the theoretical limit. Results show

that comparable performance with

turbocodes can be achieved if a

two-dimensional product code is

used whose component codes are

binary linear block codes such as

the (15,11) extended Hamming

code.

For more information, visit the

Communications Technology Division

on the World Wide Web:

http://ctd.grc.nasa.gov/5650/

5650.html

Lewis contact:

Robert E. Jones, (216) 433–3457,

Robert.E.Jones@grc.nasa.gov

Authors:

Dr. Subhash C. Kwatra and

Robert E. Jones

Headquarters program office: OSS

Programs/Projects:

Space Communications

Performance of (23,31) turbocode. Rate, 5/6; conventional puncturing

pattern.

Improved performance of the turbocode with the extended puncturing

period.

10–1

10–2

10–3

10–4

10–5

10–6

1 468 2 3 5 7 9 10

Bit error rate

Ratio of energy-to-noise to bit-error-rate, Eb/No, dB

0

3

8

12

16

Iteration

10–1

10–2

10–3

10–4

10–5

10–6

12345

Bit error rate

Ratio of energy-to-noise to bit-error-rate, Eb/No, dB

0

2

6

12

16

Iteration
90

The Communications Technology Division at the NASA Lewis Research

Center has an ongoing program to develop advanced switching and

routing technology concepts for future satellite onboard processing systems. Through a university grant as a part of this research, the Cleveland

State University is using a flexible reconfigurable data communications

packet switch emulation test bed to investigate packet switching techniques.

Because of the switching speed and protocol complexity, implementing a

data communications network is a tremendous task. Various alternatives

should be carefully studied and evaluated in the development stage so that

the optimal system configuration can be obtained and implemented later.

Therefore, it is desirable to predict the performance of the network before

it is actually constructed. This is especially true in the case of satellite systems. In the past, theoretical analysis, software simulation, and prototyping were used to evaluate performance. However, each method has its

drawback. There are basic tradeoffs among accuracy, cost, and required

evaluation time. No method is completely satisfactory.

Reconfigurable Data Communications Packet-Switch

Emulation Test Bed Demonstrated

Reconfigurable packet-switch emulation test bed.

FIFO buffer and

status circuit

Data collection

circuit

Status Cell arrival

Dest port

State Statistics

Control signals Control signals Control circuit

User interface C program

PC ISA bus interface

ISA bus interface

Address Data

Address Para Para LFSR

FIFO

count

Traffic generator

In addition to the theoretical

analysis, software simulation, and

prototype, there is one other

possible alternative for performance

evaluation—hardware emulation.

Hardware emulation is similar to

software simulation except that

dedicated hardware modules

replace software routines. The

dedicated hardware can achieve

much better performance and

provide several orders of magnitude in increased speed over

software simulation. When applied

to network evaluation, it can be

used to gather data for a realistic

network configuration and to

evaluate the performance of

different schemes and protocols.

Para Parameter Control signal from the computer hardware

FIFO count First in, first out count Signal that keeps track of the number of generated data packets

Dest port Destination port Final destination of a generated data packet

LFSR Linear Feedback Shift Register Used as a uniform random number generator for generating data packets

ISA Industry Standard Architecture Standard hardware configuration used by 286 to 486 class IBM clone computers
91

Research and Technology

In the past, hardware emulation methods suffered the drawback of high

cost and inflexibility and thus were rarely used. However, recent advances

in field programmable gate array technology make it a feasible, and

possibly superior, approach. The Cleveland State study investigated this

approach and explored the feasibility of using hardware emulation to

evaluate the performance of various asynchronous transfer mode (ATM)

switch configurations and protocols. A reconfigurable test bed was developed utilizing static RAM-based (random access memory based) field

programmable gate array technology. The flexible test bed, which was

constructed to emulate an ATM switch, and consisted of an abstract ATM

switch, a traffic generator, a data collection circuit, and a PC host interface

(see the diagram). All the test bed hardware was described in industrial

standard very high development language (VHDL). It was then synthesized

and mapped into a field programmable gate array board. The construction

and testing of the prototyping system was completed and used to evaluate

several buffer management schemes. Results show that the hardware

emulation was 102 to 105 times faster than software simulation.

For more information, visit the

Communications Technology Division

on the World Wide Web:

http://ctd.grc.nasa.gov/5650/

5650.html

Lewis contact:

Robert E. Jones, (216) 433–3457,

Robert.E.Jones@grc.nasa.gov

Authors:

Dr. Pong P. Chu and Robert E. Jones

Headquarters program office: OSS

Programs/Projects:

Space Communications

NASA Lewis Research Center has developed a digital approach for broadcasting high-fidelity audio (nearly compact disk (CD) quality sound) in

the commercial frequency-modulated (FM) broadcast band. This digital

approach provides a means of achieving high data transmission rates with

low hardware complexity—including low mass, size, and power consumption. Lewis has completed the design and prototype development of a

bandwidth-efficient digital modem

(modulator and demodulator) that

uses a spectrally efficient modulation scheme: 16-ary rectangular

quadrature amplitude modulation, or 16-ary QAM. The digital

Bandwidth-Efficient Wireless Digital Modem Developed

Bandwidth-efficient wireless digital modem with 16-ary quadrature amplitude

modulation (QAM) and digital signal processing (DSP).

Phaselock

loop

Clock Clock

Data

Serialtoparallel

DSP

Lowpass

filter

FM

transmitter

Digitaltoanalog

16-ary QAM modulator

Clock Clock

Data Decision

Phaselock

loop

DSP

Lowpass

filter

FM

receiver

Analogtodigital

16-ary QAM demodulator
92

implementation is based strictly on inexpensive, commercial off-the-shelf

digital signal processing (DSP) hardware to perform up and down conversions and pulse shaping. The digital modem transmits data at rates up to

76 kilobits per second (kbps), which is almost 3 times faster than standard

28.8-kbps telephone modems. In addition, the modem offers improved

power and spectral performance, flexible operation, and low-cost implementation.

The resulting technology, digital audio broadcasting, can provide listeners

with greatly enhanced, near-CD quality sound for FM radio, the consumer

electronics industry with a great commercial opportunity, and radio broadcasters with the ability to provide improved sound and services to listeners.

Other potential uses for this modem include digital wireless data applications (commercial, industrial, or educational), wireless transmission of

point-of-sale data and hospital records, and automation of factory control

signals. Its main benefits are low cost, high throughput, and small board

size, comparable to that of a computer modem.

Lewis designed this bandwidth efficient digital modem under a nonreimbursable Space Act Agreement with Telos System Inc., which is based in

Cleveland, Ohio. Telos Systems Inc.

has explicit potential markets and

applications for the modem in the

commercial FM broadcast band,

both in transmitting data-reduced

digital signals and in broadcasting

high-fidelity audio.

Lewis contact:

Muli Kifle, (216) 433–6521,

Muli.Kifle@grc.nasa.gov

Author: Muli Kifle

Headquarters program office: OSS

Programs/Projects: Commercial

Communications Technology

The Communications Technology Division at the NASA Lewis Research

Center is developing advanced electronic technologies for the space

communications and remote sensing systems of tomorrow. As part of the

continuing effort to advance the state-of-the-art in satellite communications and remote sensing systems, Lewis is developing a programmable

Orthogonal Frequency Division Multiplexing (OFDM) modulator card for

high-data-rate communication links. The OFDM modulator is particularly

suited to high-data-rate downlinks to ground terminals or direct data

downlinks from near-Earth science platforms. It can support data rates up

to 622 megabits per second (Mbps) and high-order modulation schemes

such as 16-ary quadrature amplitude modulation (16-ary QAM) or 8-phase

shift keying (8PSK). High order modulations can obtain the bandwidth

efficiency over the traditional binary phase shift keying (BPSK) or quadrature phase shift keying (QPSK) modulator schemes. The OFDM modulator

architecture can also be precompensated for channel disturbances and

alleviate amplitude degradations caused by nonlinear transponder

characteristics.

The 622-Mbps OFDM modulator contains four orthogonal 155-Mbps

channels. Each channel is based on SiCOM’s (Scottsdale, Arizona; see

http://www.sicom.com) SMC–960A Data Encoder and Pulse Shaper ASIC

(application-specific integrated circuit) chip, which is the result of Small

Business Innovative Research Phase II funding from Lewis.

622-Mbps Orthogonal Frequency Division Multiplexing

Modulator Developed

Find out more about Lewis’ work

with communications satellites on

the World Wide Web:

http://sulu.grc.nasa.gov

Lewis contact:

Nam T. Nguyen, (216) 433–3425,

Nam.T.Nguyen@grc.nasa.gov

Author: Nam T. Nguyen

Headquarters program office:

OSS, SBIR

Programs/Projects:

Near Earth Science, EOS, D3
93

Research and Technology

Emerging multimedia applications and future satellite systems will require

high-speed switching networks to accommodate high data-rate traffic

among thousands of potential users. This will require advanced switching

devices to enable communication between satellites. The NASA Lewis

Research Center has been working closely with industry to develop a stateof-the-art fast packet switch (FPS) to fulfill this requirement.

Recently, the Satellite Industry Task Force identified the need for highcapacity onboard-processing switching components as one of the “grand

challenges” for the satellite industry in the 21st century. In response to this

challenge, future generations of onboard processing satellites will require

low power and low mass components to enable transmission of services in

the 100 gigabit (1011 bits) per second (Gbps) range.

Lewis and Sierra Monolithics developed a state-of-the-art 10-Gbps-per-port

switch. This promising silicon-germanium (SiGe) technology enables the

design of high-speed circuits at very low power consumption. Some of the

benefits of this state-of-the-art technology follow:

• U. S. satellite industry competitiveness will be enhanced through the

development of high-throughput, low-power onboard components to

support the National and Global Information Infrastructures (NII/GII).

• Satellite and terrestrial systems will become fully interoperable.

• Bandwidth efficient systems will improve capacity 100 times over existing

satellite architectures for commercial applications.

• The complexity of Earth stations will be reduced.

Silicon-Germanium Fast Packet Switch Developed for

Communications Satellites

Silicon germanium fast packet switch test board.

The SiGe switch is based on a

16-input by 16-output SiGe crossbar switch that can transmit and

receive data at 10 Gbps per port

with very low power and weight

requirements. The switch has a

robust contention control that is

suitable for satellite onboard

processing applications.

In 1998, Sierra Monolithics simulated and tested a prototype 4-by-4

SiGe switch with favorable results.

The test board, along with the fast

packet switch chip, is shown in the

photograph. A final 16-by-16 switch

is being fabricated and will be

tested and delivered to Lewis in

1999 for further evaluation. Lewis

envisions this technology as suitable for next-generation satellite

systems for NASA missions as well

as commercial applications requiring high capacity and higher data

rates.

Lewis contact:

Gene Fujikawa, (216) 433–3495,

Gene.Fujikawa@grc.nasa.gov

Author: Jorge A. Quintana

Headquarters program office: OSS

Programs/Projects: NII/GII, commercial satellite and terrestrial communications systems

SiGe fast packet switch
94

At the NASA Lewis Research Center, a comprehensive assessment was

made of the predictive capability of the average passage flow model as

applied to multistage axial-flow compressors. This model, which describes

the time-averaged flow field within a typical passage of a blade row

embedded in a multistage configuration, is being widely used throughout

U.S. aircraft industry as an integral part of their design systems.

Turbomachinery and Propulsion Systems

Multistage Turbomachinery Flows Simulated Numerically

Rotor flow-angle deviation.

Overall pressure rise coefficient per stage.

Measured

Predicted

Flow angle

(Increased flow angle

denotes increased

deviation from blade

exit metal angle.)

Blade exit

metal angle

Fully turbulent boundary 

layer initiated at blade 

leading edge

Initially laminar boundary layer 

transitions to turbulent 

aft of blade leading edge

0.25 0.30 0.35 0.40 0.45

Inlet flow coefficient

Pressure coefficient/stage

APNASA (Standard κ−ε) 

Measurements

APNASA (Shih κ−ε)

0.35

0.40

0.45

0.50

0.55

0.60

In this work, detailed data taken within a four and one-half stage large lowspeed compressor were used to assess the weaknesses and strengths of

the predictive capabilities of the average passage flow model. The lowspeed compressor blading is of modern design and employs stator endbends. Measurements were made

with slow- and high-response

instrumentation. The high-response

measurements revealed the velocity components of both the rotor

and stator wakes. From the measured wake profiles, we found that

the flow exiting the rotors deviated

from the rotor exit metal angle to

a lesser degree than was predicted

by the average passage flow model

(see the top figure). This was found

to be due to blade boundary layer

transition, which recently has been

shown to exist on multistage axial

compressor rotor and stator blades

(ref. 1), but was not accounted for

in the average passage model.

Consequently, a model that mimics

the effects of blade boundary layer

transition, Shih k-ε model, was

incorporated into the average

passage model. Simulations that

incorporated this transition model

showed a dramatic improvement in

agreement with data (see the

bottom figure). The altered model

thus improved predictive capability

for multistage axial-flow compressors, and this was verified by

detailed experimental measurement. For more information, see

references 2 and 3.

References

1. Adamczyk, J.J., et al.: Numerical

Simulation of Multistage Turbomachinery Flows. AGARD Propulsion

and Energetics Panel. Tulous, France,

May 1998.

2. Adamczyk, J.J., et al.: Numerical

Simulation of Multistage Turbomachinery Flows. Unsteady Flows in

Turbomachines, VKI–LS–1996–05,

vol. 2, 1996.
95

Research and Technology

3. Shih, T.-H., et al. New K-ε Eddy Viscosity Model for High Reynolds Number

Turbulent Flows. Comp. Fluids, vol. 24, no. 3, 1995, pp. 227–238.

Lewis contacts:

Dr. Michael D. Hathaway, (216) 433–6250, Michael.D.Hathaway@grc.nasa.gov; and

Dr. John J. Adamczyk, (216) 433–5829, John.J.Adamczyk@grc.nasa.gov

Authors: Dr. Michael D. Hathaway, Dr. John J. Adamczyk, Aamir Shabbir, and

Steven R. Wellborn

Headquarters program office: OAT

Programs/Projects: Propulsion Systems R&T, P&PM, AST, HSR, SGE

The overall performance of a centrifugal compressor depends on the

performance of the impeller and diffuser as well as on the interactions

occurring between these components. Accurate measurements of the

flow fields in each component are needed to develop computational

models that can be used in compressor design codes. These measurements must be made simultaneously over an area that covers both components so that researchers can understand the interactions occurring

between the two components.

Optical measurement techniques are being used at the NASA Lewis

Research Center to measure the velocity fields present in both the impeller and diffuser of a 4:1 pressure ratio centrifugal compressor operating

at several conditions ranging from design flow to surge. Laser Doppler

Velocimetry (LDV) was used to measure the intrablade flows present in

the impeller, and the results were compared with analyses obtained from

two three-dimensional viscous codes. The development of a region of low

throughflow velocity fluid within this high-speed impeller was examined

and compared with a similar region first observed in a large low-speed

centrifugal impeller at Lewis.

Optical Flow-Field Techniques Used for Measurements

in High-Speed Centrifugal Compressors

Contours of impeller throughflow velocity from LDV measurements at 70 percent

of the blade chord, where Vqm is the quasi-meridional, or throughflow, velocity;

Utip is the blade tip speed at the impeller discharge; PS is the impeller blade

pressure surface; and SS is the suction surface. This figure is shown in color in

the online version of this article (http://www.grc.nasa.gov/WWW/RT1998/5000/

5810skoch.htm).

Particle Image Velocimetry (PIV) is

a relatively new technique that has

been applied to measuring the

diffuser flow fields. PIV can collect

data rapidly in the diffuser while

avoiding the light-reflection problems that are often encountered

when LDV is used. The Particle

Image Velocimeter employs a sheet

of pulsed laser light that is introduced into the diffuser in a quasiradial direction through an optical

probe inserted near the diffuser

discharge. The light sheet is positioned such that its centerline is

parallel to the hub and shroud

surfaces and such that it is parallel to the diffuser vane, thereby

avoiding reflections from the solid

surfaces. Seed particles small

enough to follow the diffuser flow

are introduced into the compressor

at an upstream location. A highspeed charge-coupled discharge

(CCD) camera is synchronized to

the laser pulse rate; this allows it

to capture images of seed particle

position that are separated by a

small increment in time. A crosscorrelation of a particle’s position

in two consecutive images provides

an estimate of flow velocity and

direction. Multiple image pairs

obtained in rapid succession at a

particular flow condition provide

enough measurements for statistical significance. PIV provides

Vqm/Utip

PS SS PS SS

Splitter

0.25 0.28 0.31 0.34 0.37 0.40
96

Diffuser absolute velocity field at 88 percent of the passage height obtained using Particle Image Velocimetry.

This figure is shown in color in the online version of this

article (http://www.grc.nasa.gov/WWW/RT1998/5000/

5810skoch.htm).

Axial position, mm

Longitudinal position, mm

Particle

velocity,

m/sec

375.0

366.1

357.1

348.2

339.3

330.4

321.4

312.5

303.6

294.6

285.7

276.8

267.9

258.9

250.0

simultaneous velocity measurements over the entire plane that is illuminated by the light sheet instead of at a single point, as is the case when

LDV is used.

PIV has a further advantage in that the laser light pulse can be triggered

by an external source such as a high-response pressure transducer. This

feature will allow PIV to synchronize flow imaging to physical phenomena

such as rotating stall or stall precursor waves. We hope that this

technique can be used to obtain

images of the flow field during

and just prior to stall.

Bibliography

Skoch, G.J., et al.: Laser Anemometer

Measurements of the Flow Field in a

4:1 Pressure Ratio Centrifugal Impeller.

ASME Paper 97–GT–342, 1997.

Hathaway, M.D., et al.: Experimental

and Computational Investigation of the

NASA Low-Speed Centrifugal Compressor Flow Field. ASME J. Turbomachinery

(ASME Paper 92–GT–216), vol. 115,

no. 3, July 1993, pp. 527–541.

Wernet, M.P.: Digital PIV Measurements

in the Diffuser of a High Speed Centrifugal Compressor. AIAA Paper 98–2777,

1998.

Lewis contacts:

Gary J. Skoch, (216) 433–3396,

Gary.J.Skoch@grc.nasa.gov; and

Dr. Mark P. Wernet, (216) 433–3752,

Mark.P.Wernet@grc.nasa.gov

Author: Gary J. Skoch

Headquarters program office: OAT

Programs/Projects:

Propulsion Systems R&T, P&PM
97

Research and Technology

For the last several years, LeRC–HT, a three-dimensional computational

fluid dynamics (CFD) computer code for analyzing gas turbine flow and

convective heat transfer, has been evolving at the NASA Lewis Research

Center. The code is unique in its ability to give a highly detailed representation of the flow field very close to solid surfaces. This is necessary for an

LeRC–HT: NASA Lewis Research Center General Multiblock

Navier-Stokes Heat Transfer Code Developed

Visualization of the flow and heat transfer over the tip of a high-pressure turbine

rotor blade as computed with the LeRC–HT code. Darkest areas represent undesirable high-heat-flux regions. This figure is shown in color in the online version of

this article (http://www.grc.nasa.gov/RT1998/5000/5820heidmann.htm).

accurate representation of fluid

heat transfer and viscous shear

stresses. The code has been used

extensively for both internal cooling

passage flows and hot gas path

flows—including detailed film

cooling calculations, complex tipclearance gap flows, and heat

transfer. In its current form, this

code has a multiblock grid capability and has been validated for a

number of turbine configurations.

The code has been developed and

used primarily as a research tool

(at least 35 technical papers have

been published relative to the code

and its application), but it should

be useful for detailed design analysis. We now plan to make this

code available to selected users for

further evaluation.

Lewis contacts:

Dr. James D. Heidmann, (216) 433–3604,

James.D.Heidmann@grc.nasa.gov; and

Dr. Raymond E. Gaugler, (216) 433–5882,

Raymond.E.Gaugler@grc.nasa.gov

Authors:

Dr. James D. Heidmann and

Dr. Raymond E. Gaugler

Headquarters program office: OAT

Programs/Projects:

Propulsion Systems R&T, AST, HSR
98

NASA Lewis Research Center’s new, world-class, 60-atmosphere (atm)

combustor research facility, the Advanced Subsonic Combustion Rig

(ASCR), is in operation and producing highly unique research data. Specifically, data were acquired at high pressures and temperatures representative of future subsonic engines from a fundamental flametube configuration with an advanced fuel injector. The data acquired include exhaust

emissions as well as pressure and temperature distributions. Results to

date represent an improved understanding of nitrous oxide (NOx) formation at high pressures and temperatures and include an NOx emissions

reduction greater than 70 percent with an advanced fuel injector at operating pressures to 800 pounds per square inch absolute (psia).

ASCR research is an integral part of the Advanced Subsonic Technology

(AST) Propulsion Program. This program is developing critical low-emission

combustion technology that will result in the next generation of gas turbine

engines producing 50 to 70 percent less NOx emissions in comparison to

1996 International Civil Aviation Organization (ICAO) limits. The results to

date indicate that the AST low-emission combustor goals of reducing NOx

emissions by 50 to 70 percent are feasible.

Research Data Acquired in World-Class, 60-atm Subsonic

Combustion Rig

U.S. gas turbine manufacturers

have started testing the lowemissions combustors at the ASCR.

This collaborative testing will enable the industry to develop lowemission combustors at the high

pressure and temperature conditions of future subsonic engines.

The first stage of the flametube

testing has been implemented.

Four GE Aircraft Engines lowemissions fuel injector concepts,

three Pratt & Whitney concepts, and

two Allison concepts have been

tested at Lewis’ ASCR facility. Subsequently, the flametube was

removed from the test stand, and

the sector combustor was installed.
99

Research and Technology

CD-95-70913

Ceramic liner

Inlet 

air

Water-cooled

flanges

Zirconium oxide

(ZrO) liner

Ceramic liner

Thermocouple

Exhaust

Thermocouple

Thermocouple

Sample probe

Multiple point, lean 

direct injector (LDI)

Pressure rake

Hydrogen torch

Advanced Subsonic Technology flametube combustor.

The testing of low-emissions sector has begun. Low-emission combustors

developed as a result of ASCR research will enable U.S. engine manufacturers to compete on a worldwide basis by producing environmentally

acceptable commercial engines.

Lewis contacts: Dr. Chi-Ming Lee, (216) 433–3413, Chi-Ming.Lee@grc.nasa.gov;

and Dr. Changlie Wey, (216) 433–5379, Changlie.Wey@grc.nasa.gov

Authors: Dr. Chi-Ming Lee and Dr. Changlie Wey

Headquarters program office: OAT

Programs/Projects: AST

The NASA Lewis Research Center, in cooperation with Rocketdyne, the

Boeing Company, tested a novel rocket engine ignition system, called the

combustion-wave ignition system, in its Research Combustion Laboratory.

This ignition system greatly simplifies ignition in rocket engines that have

a large number of combustors. The particular system tested was designed

and fabricated by Rocketdyne for the national experimental spacecraft,

X–33, which uses Rocketdyne’s aerospike rocket engines.

The goal of the tests was to verify the system design and define its operational characteristics. Results will contribute to the eventual successful

flight of X–33. Furthermore, the combustion-wave ignition system, after

it is better understood and refined on the basis of the test results and,

later, flight-proven onboard X–33, could become an important candidate

engine ignition system for our Nation’s next-generation reusable launch

vehicle.

X–33 Combustion-Wave Ignition System Tested

This recent test program successfully proved that the ignition system

design is sound. In addition, it identified the operational characteristics of the ignition system that are

important to the X–33 vehicle as

well as for other future applications.

The program started in 1997, first

with a subscale, prototype system

that used gaseous propellants, and

then with a full-scale, flightlike,

prototype system that used liquid

hydrogen and oxygen propellants to

fully simulate the flight conditions.
100

The general concept of the combustion-wave ignition system is that it

initiates and distributes the flames of its premixed propellants so that they

ignite in multiple locations. It has three basic components: a propellant

premixer, an ignition device mounted to the premixer, and a number of

tubes stemming from the premixer. The premixer mixes the incoming

propellants as they flow into the tubes. The tubes are each terminated at a

location where ignition is desired (for example, in a combustor). Several

considerations need to be taken, however. First, the propellants in the

ignition system should be in a gaseous state and well mixed for the flames

to propagate successfully. Second, the propellant pressure in the system

should be appropriate in order to avoid damage to the system upon flame

initiation. Third, the timing should be optimized for propellant priming and

flame initiation at the premixer for each unique application of the system.

Combustion-wave ignition system testing for the X–33 demonstrator vehicle.

The combustion-wave ignition

system is simple to design and

fabricate, and it offers weight savings, decreased electrical power

consumption, and a precise and

uniform timing for ignition at

multiple locations.

Lewis contact:

Larry C. Liou, (216) 977–7433,

Larry.C.Liou@grc.nasa.gov

Author: Larry C. Liou

Headquarters program office: OAT

Programs/Projects: X–33 Advanced

Technology Demonstrator
101

Research and Technology

At the NASA Lewis Research Center, the launch vehicle gross lift-off weight

(GLOW) was analyzed for solid particle feed systems that use high-energy

density atomic propellants (ref. 1). The analyses covered several propellant

combinations, including atoms of aluminum, boron, carbon, and hydrogen

stored in a solid cryogenic particle, with a cryogenic liquid as the carrier

fluid. Several different weight percents for the liquid carrier were investigated, and the GLOW values of vehicles using the solid particle feed systems

were compared with that of a conventional oxygen/hydrogen (O2/H2)

propellant vehicle.

Atomic propellants, such as boron, carbon, and hydrogen, have an enormous potential for high specific impulse I

sp operation, and their pursuit has

been a topic of great interest for decades. Recent and continuing advances

in the understanding of matter, the development of new technologies for

simulating matter at its most basic level, and manipulations of matter

through microtechnology and nanotechnology will no doubt create a bright

future for atomic propellants and an exciting one for the researchers

exploring this technology.

New technologies in atom isolation and in the physics of material manipulation have led to the discovery and synthesis of materials that can be used

as rocket propellants (refs. 1 to 3). Solid cryogenic propellants storing

atoms of aluminum, boron, carbon, or hydrogen, or other atomic additives,

require a unique propulsion system design where the fuels are stored at

liquid helium temperatures during ground handling and flight. Feeding

atomic propellants from a propellant tank, through a feed system, to a

rocket engine will be a formidable challenge.

Use of Atomic Fuels for Rocket-Powered Launch Vehicles

Analyzed

The performance with atomic H

was the highest of any of the cases

investigated, with its I

sp values

ranging from 600 to nearly 1300

sec. As shown in this figure, the

highest monopropellant performance was 1500 sec delivered at

100 weight percent (wt %) of

atomic hydrogen (H). Because it

is unlikely that we will be able to

store 100 wt % of atomic H, the

lower levels of 10, 15, and 50 wt %

were investigated in these GLOW

analyses.

The figure on the next page illustrates the GLOW values for atomic

H. With atomic H, no GLOW reductions were possible until 15 wt %

of atomic H was used. At the

10-wt % level, the atomic H GLOW

was 4-percent greater than the

O2/H2 baseline. At a 15-wt % level,

the GLOW was reduced by 44 percent. With a 50-wt % atom loading,

the GLOW was reduced by 78 percent.

When helium (He) was added to

the flow with 10-wt % atomic H,

the GLOW was 22- to 106-percent

higher than for the baseline case,

as shown in the previous figure.

The 15-wt % atomic H cases only

exceeded the baseline GLOW

when 40-wt % He was added. For

these cases, the GLOW ranged from

a reduction of 37 percent (with a

10-wt % He addition) to an increase

of 6 percent (with a 40-wt % He

addition). The 50-wt % atomic H

cases were almost unaffected by

the He addition. The GLOW was

72-percent less than the O2/H2 vehicle GLOW, even with the 40-wt %

He addition. In summary, with

atomic H at a 10-wt % level, no

reductions over O2/H2 vehicles

were possible. At 15-wt % atomic

H, there were very significant GLOW

Atomic hydrogen rocket engine performance.

0

40

20

60

80 100

0.00

0.25

0.50

0.75

1.00

Rocket engine specific 

impulse, Isp, sec

1600

1200

Oxygen to fuel ratio

H atom concentration, wt %

0

800

400
102

reductions for all but the 40-wt % He addition case. The

50-wt % atomic H cases showed a strong insensitivity to

He addition.

For more information, visit the SBIR Fuels and Space Propellants

web site: http://www.grc.nasa.gov/WWW/TU/launch/foctopsb.htm

References

1. Palaszewski, B.: Launch Vehicle Performance With Solid Particle

Feed Systems for Atomic Propellants. AIAA Paper 98–3736 (NASA

TM—1998-208498), 1998.

2. Palaszewski, B.: Atomic Hydrogen Propellants: Historical Perspectives and Future Possibilities. AIAA Paper 93–0244 (NASA TM–

106053), 1993.

3. Carrick, P.G., and Tam, S., eds.: Proceedings of the High Energy

Density Materials (HEDM) Contractors’ Conference. USAF Phillips

Laboratory, Report Number PL–TR–95–3039, 1996.

Lewis contact: Bryan A. Palaszewski, (216) 977–7493,

fax (216) 977–7545, Bryan.A.Palaszewski@grc.nasa.gov

Author: Bryan A. Palaszewski

Headquarters program office: OAT

Programs/Projects: ASTP

GLOW, kg 

0

10

20

40

Helium added,

wt % 

5x106

4

3

2

1

0

Propulsion technology

H atom concentration in solid 

particle feed system, wt %

10 15 50

NLS

Comparison of atomic propellant launch vehicle

gross lift-off weight (GLOW) for the National Launch

System (NLS, a conventional O2/H2 vehicle) and

solid particle feed systems.

The NASA/Federal Aviation Administration (FAA) Tailplane Icing Program

(TIP) resulted from a concern about aircraft incidents and accidents due to

ice-contaminated tailplane stall (ICTS). At least 16 accidents resulting in

139 fatalities have been attributed to such stall. The objectives of this

program were to improve understanding of iced tailplane aeroperformance and aircraft aerodynamics

and to develop training aids and

design tools to expand the awareness of ICTS.

TIP was a 4-yr research program

that employed a combination of

facilities. These included NASA

Lewis Research Center’s Icing

Research Tunnel to generate actual

ice shapes, the Ohio State University’s 7310 ft Low Speed Wind

Tunnel to provide tailplane aeroperformance data, and finally Lewis’

icing research aircraft, a modified

DeHavilland DHC–6 Twin Otter, for

tailplane aeroperformance and aircraft aerodynamics measurements.

The highlighted accomplishments

to date follow:

NASA/FAA Tailplane Icing Program Completed

Lewis’ Twin Otter icing research aircraft.
103

Research and Technology

Twin Otter tailplane with artificial ice shape and flow probes. The ice shapes were

cast from ice generated in Lewis’ Icing Research Tunnel.

• Development of a comprehensive database on tailplane

aeroperformance

• Identification of dominant drivers that lead to tailplane stall

• Demonstration of an effective tailplane stall recovery procedure

• Provision of in-flight demonstrations to guest pilots

• Production and distribution of an educational video for pilots

At the conclusion of the flight tests, we felt that TIP had gathered enough

new information that it should be shared. A Guest Pilot Workshop was

developed to demonstrate first hand the unique flying qualities of an aircraft with an ice-contaminated tailplane. An international group representing various facets of the aviation industry—aviation regulatory agencies,

aircraft manufacturers, and aviation media pilots and reporters—were

invited. In total, 15 guest pilots and engineers flew the Twin Otter with

the tailplane ice shape attached. This demonstration program provided a

mutually beneficial forum for the exchange of information between NASA

and the user community. Feedback

from the guest pilots has been very

positive. Each indicated a greater

appreciation of ICTS as a result of

this guest pilot workshop. In addition, a 23-min video, Tailplane

Icing, was produced and distributed. The target audience is primarily pilots who might encounter

in-flight icing. It has been enthusiastically received.

TIP and the Guest Pilot Workshop

are excellent examples of NASA’s

response to flight safety initiatives

and quick release of information to

the user community. TIP is now

complete, but a follow-on program,

TIP 2, has begun to look at other

tailplane configurations and subscale test methods. Learjet and The

Wichita State University are partners in the continued activity.

Find out more about this research in

the Icing Branch’s document archive:

http://icebox.grc.nasa.gov/

DocArchive.html

Lewis contacts:

Thomas P. Ratvasky, (216) 433–3905,

Thomas.P.Ratvasky@grc.nasa.gov; and

Judith Foss Van Zante, (216) 433–3587,

Judith.K.VanZante@grc.nasa.gov

Authors:

Thomas P. Ratvasky and

Judith Foss Van Zante

Headquarters program office: OAT

Programs/Projects: AOS

Special recognition: Six articles in

aviation magazines including Aviation

Week, Flying, and Airline Pilot
104

The main purpose of this investigation was for NASA to help the National

Transportation Safety Board (NTSB) gain better understanding of the events

that led to the loss of Comair Flight 3272 over Monroe, Michigan, on January 9, 1997. In-flight icing was suspected as being the primary cause of this

accident. Of particular interest to the Safety Board was what NASA could

learn about the potential performance degradation of the wing of the

Embraer EMB–120 twin-turboprop commuter aircraft with various levels of

ice contamination.

NASA agreed to undertake (1) ice-accretion prediction computations with

NASA’s LEWICE program to bound the kind of contaminations that the

vehicle may have developed, (2) testing in the NASA Lewis Research

Center’s Icing Research Tunnel to verify and refine the ice shapes developed by LEWICE, (3) a two-dimensional Navier-Stokes analysis to determine the performance degradation that those ice shapes could have

caused, and (4) an examination using three-dimensional Navier-Stokes

codes to study the three-dimensional effects of ice contamination.

National Transportation Safety Board Aircraft Accident

Investigation Supported

Ice accreted during the testing of the EMB–120 wing in Lewis’ Icing

Research Tunnel.

The experimental and twodimensional Navier-Stokes study

showed a possible scenario of

what may have happened during

the accident. Several of the ice

shapes accreted in the Icing

Research Tunnel had a roughleading-edge ice accretion with a

prominent ridge. The most significant finding of the analysis was

that the ridged ice shape produced

lift and drag degradations that could

have caused the kind of control

disruption recorded in Flight 3272.

The National Transportation Safety

Board presented its findings in a

public meeting on August 27,1998.

One of the conclusions was that ice

formation played a significant role

in the accident. Among the safety

recommendations, two were

directed to NASA. The first recommended that NASA continue its

efforts to develop training tools to

educate the flying community

regarding the hazards of in-flight

icing. The second advised NASA to

conduct additional research to

define realistic ice shapes and

understand their effect on aerodynamics to help in defining future

certification and training requirements.

Find out more on the World Wide

Web: http://icebox.grc.nasa.gov/

Lewis contacts:

Andrew L. Reehorst, (216) 433–3938,

Andrew.L.Reehorst@grc.nasa.gov; and

Dr. Mark G. Potapczuk, (216) 433–

3919, Mark.G.Potapczuk@grc.nasa.gov

Author: Andrew L. Reehorst

Headquarters program office: OAT

Programs/Projects: AOS
105

Research and Technology

Turbofan engine-face flow distortion is one of the most troublesome and

least understood problems for designers of modern engine inlet systems

(refs. 1 and 2). One concern is that there are numerous sources of flowfield distortion that are ingested by the inlet or generated within the inlet

duct itself. Among these are (1) flow separation at the cowl lip during

in-flight maneuvering, (2) flow separation on the compression surfaces

due to shock-wave/boundary layer interactions, (3) spillage of the fuselage boundary layer into the inlet duct, (4) ingestion of aircraft vortices

and wakes emanating from upstream disturbances, and (5) strong secondary flow gradients and flow separation induced by wall curvature within

the inlet duct itself. Most developing aircraft (including the B70, F–111,

F–14, Mig–25, Tornado, and Airbus A300) have experienced one or more

of these types of problems, particularly at high Mach numbers and/or

extreme maneuver conditions when flow distortion at the engine face

exceeded the allowable limits of the engine.

Using vane vortex generators to “locally” control the effects of separation

has been the most common method of flow control in inlet ducts. Lowand high-momentum regions are mixed locally in the flow to effectively

spread out the lower momentum fluid and suppress flow separation from

the wall. However, in advanced, inlet serpentine S-duct configurations, this

method often does not achieve significant control and reduction in engine

face distortion. Furthermore, locally used vortex generators can control

separation at one flow condition only (usually the cruise condition)—all

other flow conditions are off-design.

In this study at the NASA Lewis Research Center, vortex generators were

used in an entirely different manner. They were used to “globally” restructure secondary flow to increase inlet total pressure recovery and decrease

engine face distortion (as first proposed in refs. 3 and 4). Such vortex generator installations can be optimized for optimum system levels of the inlet

total pressure recovery and engine face distortion level over a wide range

of inlet operating conditions. Thus, flow separation is not prevented unless

this improves the overall inlet system as measured by engine face flow

characteristics. With this global method, computational fluid dynamics

(CFD) and design of experiments (DOE) optimization procedures can be

used to design vortex generator installations that encompass a wide variety

of inlet operating conditions.

Work done under the NASA/MOD cooperative Joint Aeronautical Program

(refs. 5 and 6) led to an equivalency-of-flow principle: It does not matter

how vorticity is produced; all that matters is the overall vorticity strength

and distribution and how it interacts with the inlet secondary flow field.

Consequently, inlet flow-control research shifted from concerns for vane

generator geometry and preventing flow separation to establishing the

proper overall vorticity signature to maximize inlet total pressure recovery

and minimize engine face distortion. This equivalency principle simplifies

the modeling of vortex generators by full Navier-Stokes CFD analyses, and

it gives excellent performance comparisons with experimental data even

for coarse grid calculations (ref. 7).

Microelectrical Mechanical Systems Flow Control Used to

Manage Engine Face Distortion in Compact Inlet Systems

The current development strategy

for combat aircraft emphasizes

reducing life-cycle cost without

compromising aircraft performance

and survivability. Because this

strategy has been extended to the

component level, advanced S-duct

inlet configurations are being made

more compact (or shorter) to

reduce weight and volume (and

life-cycle cost). However, these

compact S-ducts are typified by

high distortion and low pressure

recovery because of the extreme

wall curvature and strong secondary flow gradients. Flow-control

methods are being developed for

these compact ducts to improve

their performance. In addition,

microelectrical mechanical systems

(or MEMS-based) actuators are

being evaluated as minimally

intrusive, lightweight, controllable

flow-control effectors to replace

fixed-vane vortex generators. However, before MEMS technology can

manage inlet flow fields successfully, much basic and applied aerodynamic research must be done.

This study used a full Navier-Stokes

analysis to examine MEMS flow

control of compact inlets, focusing

on developing a workable installation. Instead of examining a particular MEMS device, it defined a

device that could manage the

entire inlet flow field effectively.

This reference MEMS device has a

vane height between 2 and 4 mm

above the duct walls (about the

height of the boundary layer

momentum thickness). The objectives of this study were to

(1) Enable the MEMS reference

vane generator installations to

manage the entire inlet flow field

by controlling the secondary flow in

the thin layer adjacent to the wall.
106

(2) Use DOE procedures and full Navier-Stokes analysis to determine

design rules for the MEMS installations, where the grid defines the individual vanes.

(3) Validate that CFD modeling can be used effectively to design MEMS

installations by using a full Navier-Stokes analysis to repeat the DOE design

matrix, where the individual vanes are modeled.

These CFD experiments were conducted using the Defense Evaluation

Research Agency DERA/M2129 inlet S-duct (ref. 8) and the Lockheed

Martin Tactical Aircraft Systems (LMTAS) full Navier-Stokes FALCON code

(ref. 7), which includes a vortex generator model. In part, this study, which

is part of The Technical Cooperative Program (TTCP), is a validation of the

vortex generator model in the FALCON code.

The figure shows particle traces emanating just below the height of the

MEMS reference vane generator installation, a co-rotating installation.

Co-rotating installations interact to create vortices that merge quickly to

form an overall single secondary flow pattern that remains in the wall

boundary layer. This induced flow pattern prevents pairs of counter-rotating

vortices from forming in the S-duct and, consequently, prevents the damaging effects of vortex lift-off (separation) on engine face distortion.

Particle traces showing the secondary flow field induced by the MEMS-scale

reference generator installation.

References

1. Advisory Group for Aerospace Research and Development (AGARD):

Engine Response to Distorted Inflow

Conditions. AGARD CP–400, 1987.

2. E.L. Goldsmith and J. Seddon, eds.:

Practical Intake Aerodynamic Design.

Blackwell Scientific Publications,

Oxford, 1993.

3. Anderson, B.H., et al.: Study of Vortex

Flow Control of Inlet Distortion. J.

Propuls. Pow., vol. 8, no. 6, Nov.—

Dec. 1992, pp. 1266–1272.

4. Anderson, B.H.; and Gibb, J.: Study

on Vortex Generator Flow Control for

the Management of Inlet Distortion.

J. Propuls. Pow., vol. 9, no. 3, May—

June 1993, pp. 420–430.

5. Anderson, B.; and Gibb, J.: VortexGenerator Installation Studies on

Steady-State and Dynamic Distortion. AIAA Paper 96–3279, 1996.

6. Gibb, J.; and Anderson, B.H.: Vortex

Flow Control Applied to Aircraft

Intake Ducts. Proceedings of the

High Lift and Separation Control

Conference, Royal Aeronautical

Society, U.K., 1995, pp. 14.1 to

14.11.

7. Bender, E.E.; Anderson, B.H.; and

Yagle, P.T: Vortex Generator Modeling for Navier-Stokes Codes.

Symposium on Flow Control of Wall

Bounded and Free Shear Flow, 1999

Joint ASME/JSME Fluids Engineering

Conference, ASME Paper FEDSM99–

6919, 1999.

8. Willmer, A.C.; Brown, T.W.; and

Goldsmith, E.L.: Effects of Intake

Geometry on Circular Pitot Intake

Performance at Zero and Low Forward Speeds. Aerodynamics of

Power Plant Installation, AGARD

CP–301, Paper 5, 1981, pp. 51–56.

Lewis contact:

Bernhard H. Anderson, (216) 433–5822,

Bernhard.H.Anderson@grc.nasa.gov

Authors:

Bernhard H. Anderson and

Daniel N. Miller

Headquarters program office: OAT

Programs/Projects:

Propulsion Systems R&T, TTCP
107

Research and Technology

A Two-Dimensional Bifurcated (2DB) Inlet was successfully tested in NASA

Lewis Research Center’s 10- by 10-Foot Supersonic Wind Tunnel. These

tests were the culmination of a collaborative effort between the Boeing

Company, General Electric, Pratt & Whitney, and Lewis. Extensive support

in-house at Lewis contributed significantly to the progress and accomplishment of this test. The results, which met or exceeded many of the HighSpeed Research (HSR) Program goals, were used to revise system studies

within the HSR Program. The HSR Program is focused on developing lownoise, low-polluting, high-efficiency supersonic commercial aircraft. A supersonic inlet is an important component of an efficient, low-noise vehicle.

The Two-Dimensional Bifurcated (2DB) Inlet model is a mixed-compression inlet that was designed to efficiently and reliably convert the high

supersonic airflow of a High-Speed Civil Transport (HSCT) into highpressure, subsonic flow for a jet engine. This inlet concept is different from

the external compression inlets used on most supersonic fighter aircraft.

Testing began in October 1996 with a “cold-pipe” test. A cold pipe is a

large valve used as a simple simulation of a jet engine. A year later, the

cold-pipe hardware was replaced with a J85 turbojet engine. Recent tests

characterized the operability and dynamics of the inlet installed in front of

the J85 Engine. Testing was conducted over a range of supersonic Mach

numbers as well as at some subsonic flow conditions.

When compared with an external compression inlet, a mixed-compression

inlet has generally better efficiency but greater natural flow instability. This

potential instability can be controlled through the proper choice of variable

geometry. Features of this model included remotely variable ramp geometry, multiple bleed compartments, and bypass valves. Therefore, variable

geometry features were needed to develop a robust control for this mixedcompression inlet over the entire High-Speed Civil Transport flight regime.

Two-Dimensional Bifurcated Inlet/Engine Tests Completed

in 10- by 10-Foot Supersonic Wind Tunnel

Two-Dimensional Bifurcated Inlet and J85 Engine installed in

Lewis’ 10- by 10-Foot Supersonic Wind Tunnel.

Back end of installation showing the nozzle leaves of the

J85 Engine.

Operation across variations in

normal shock location, inlet

unstart/restart, and angle of attack

was demonstrated both with and

without the J85 Engine installed.

This program is an example of a

successful industry/NASA partnership. The test series was responsive

to the needs of the HSR program.

The end result was a test that

brought efficient, robust mixedcompression inlet concepts one

step closer to reality.

Find out more on the World Wide

Web:

http://www.grc.nasa.gov/WWW/HSR

http://www.grc.nasa.gov/WWW/inlet

Lewis contacts:

John D. (Dave) Saunders, (216) 433–

6278, John.D.Saunders@grc.nasa.gov;

and David J. Arend, (216) 433–2387,

David.J.Arend@grc.nasa.gov

Author: John D. Saunders

Headquarters program office: OAT

Programs/Projects: HSR, HSCT

C-98-303 C-98-131
108

0.0 0.1 0.2 0.3 0.5 0.6 0.7

Mach number

The WIND code is a general-purpose, structured, multizone, compressible

flow solver that can be used to analyze steady or unsteady flow for a wide

range of geometric configurations and over a wide range of flow conditions. WIND is the latest product of the NPARC Alliance, a formal partnership between the NASA Lewis Research Center and the Air Force Arnold

Engineering Development Center (AEDC). WIND Version 1.0 was released

in February 1998, and Version 2.0 will be released in February 1999.

The NPARC (National Program for Applications-Oriented Research in

CFD) Alliance was established in 1992 with the goal of providing an

applications-oriented computational fluid dynamics (CFD) system, primarily for aerospace flow simulation. This alliance is committed to the longrange maintenance and improvement of this capability, with teams

focused on user support, code development, and validation. The unique

talents and capabilities of its partners, plus the experience and insight of

its government, industrial, and academic customers, are used to ensure

that the alliance’s efforts are cost-effective and responsive to users’ needs.

The WIND code represents a merger of the capabilities of three existing

computational fluid dynamics codes—NPARC (the original NPARC Alliance

flow solver), NXAIR (an Air Force code used primarily for unsteady store

separation problems), and NASTD (the primary flow solver at McDonnell

Douglas, now part of Boeing). The development effort began in early 1997

using the NASTD code as the basis, and was carried out jointly between

Lewis, AEDC, and Boeing. Recently, the Air Force Research Laboratory

joined the effort. Funding for the work has been provided by NASA’s HighSpeed Research (HSR) Program, by AEDC, by Boeing, and by the Air Force’s

High Performance Computing Modernization Program.

WIND Flow Solver Released

Mach number contours for flow through a subsonic diffuser with turning. This figure is shown in color in the online version of

this document (http://www.grc.nasa.gov/WWW/RT1998/5000/5850towne.htm).

A variety of physical models are

available in the WIND code,

allowing users to choose those

appropriate to the problem being

analyzed. These include

• Navier-Stokes, thin-layer NavierStokes, parabolized Navier-Stokes,

or Euler equations

• Algebraic, one-equation, and

two-equation turbulence models

• Perfect gas, frozen chemistry,

equilibrium air, and finite rate

chemistry models

• Abutting and/or overlapping grid

zones

In addition, a variety of numerical

models are available, including

• Explicit, scalar implicit, or block

implicit solution operators

• Central, Coakley upwind, Roe

upwind, and physical upwind

explicit differencing—first to fifth

order

• Implicit and explicit boundary

conditions
109

Research and Technology

• Second- and fourth-order explicit smoothing, boundary damping, and

TVDa flux limiting

• Time step specification via a CFLb number, Global Newton iteration, and

Runge-Kutta schemes

• Convergence acceleration using grid sequencing, local CFL numbers, and

ramped CFL numbers

The WIND code is available to all U.S.-owned companies, public and

private universities, and Government agencies. However, only U.S. citizens

and resident aliens may access the software. Instructions for obtaining the

code are available on the NPARC Alliance home page on the World Wide

Web or from the NPARC Alliance User Support team.

_________________________

aTotal Variation Diminishing.

bCourant-Friedrichs and Lewy.

Additional information about the

NPARC Alliance and the WIND

code is available at the following

WWW sites:

NPARC Alliance:

http://www.arnold.af.mil/nparc

WIND documentation: http://

www.grc.nasa.gov/WWW/winddocs

WIND validation:

http://www.grc.nasa.gov/WWW/wind/

valid/validation.html

Lewis contact:

Dr. Charles E. Towne, (216) 433–5851,

Charles.E.Towne@grc.nasa.gov

AEDC contact: NPARC Alliance

User Support, (931) 454–7455,

nparc-support@info.arnold.af.mil

Author: Dr. Charles E. Towne

Headquarters program office: OAT

Programs/Projects: HSR, AST, HPCCP,

Propulsion Systems Base R&T

About half of all scientific and commercial spacecraft spin during some part

of their mission. Although spinning has many benefits (increasing stability,

controlling the location of liquid propellants, and distributing solar heat

loads), it also creates problems because a precession (or wobble) motion

is unavoidable. For modern spacecraft, by far the largest source of wobble

is energy dissipation in the liquid of partially filled tanks. The liquid’s

energy dissipation cannot, however, be quantified adequately by any

ground test. Current analytical models are also inadequate because

fundamental data on fluid motion in low gravity are needed to validate

them. Consequently, spacecraft attitude-control systems are designed and

operated very conservatively. Nonetheless, spacecraft often still perform

poorly in orbit, and some have been lost because of a rapid increase of the

wobble rate. The Liquid Motion Experiment (LME) was designed to provide

spacecraft designers accurate data on the wobble dynamics of spacecraft

that contain large quantities of mobile liquids. LME, which was flown on

the space shuttle mission STS–84, was built under contract to the NASA

Lewis Research Center by the Southwest Research Institute of San Antonio,

Texas. Major accomplishments for 1998 include reduction of the flight data

and publication of the experimental results.

LME was essentially a spin table that created a realistic nutation motion

of scale-model tanks containing liquid. Two spherical and two cylindrical

transparent tanks were tested simultaneously, and three sets of such tanks

were employed to vary liquid viscosity, fill level, and propellant management device (PMD) design. All the tanks were approximately 4.5 in. in

diameter. The primary test measurements were the radial and tangential

torques exerted on the tanks by the liquid. These torques could not be

Results of Liquid Motion Experiment Analyzed

measured on the ground because

of the masking effects of gravity.

Liquid resonant frequencies and

viscous damping coefficients were

determined by sine sweep tests.

The observed resonance frequencies for cylindrical tanks depended

on the fill level. These frequencies

were in the range of 0.73 to 0.78

times the spin rate for resonances

about the radial axis and were in

the range of 0.55 to 75 times the

spin rate for resonances about

the tangential axis. The available

analytical model predicted only

one resonance for a given fill level,

and this prediction agreed rather

closely with the tangential axis

resonance frequencies observed in

the tests. The resonances for the

spherical tanks, which were in the

range of 0.74 to 77 times the spin

rate, did not vary significantly

between the tangential and radial

axes. The PMD’s sometimes

enhanced the resonances and
110

0.00 0.20 0.40 0.60 0.80 1.00

Tank fill fraction

0.4

0.6

0.8

Resonance frequency/spin rate ratio, λ/Ω0

Analytical model

1

10

1

10

Radial

Radial

Tangential

Tangential

Torque Viscosity of liquid,

centipoise

Solid symbols denote tests at 20 rpm

Open symbols denote tests at 14.1 rpm

energy dissipation rates and sometimes decreased them. This inconsistency points out our need to better understand the effects of PMD on

liquid motion as a function of PMD and tank design. Energy dissipation

rates were determined by sine dwell tests. The LME energy dissipation

rates varied from 0.3 to 0.5 times the estimates obtained from previous

ground tests and spacecraft flight data.

Resonant frequencies for bare cylindrical tanks at

high spin rates (20 and 14.1 rpm spin).

Find out more on the World Wide

Web: http://space-power.grc.nasa.gov/

ppo/lme/lme.html

Bibliography

Chato, D.J., et al.: Liquid Motion

Experiment Flight Test Results. NASA

TM—1998-208479, 1998.

Deffenbaugh, D.M.; Dodge, F.T.; and

Green, S.T.: Liquid Motion in a Rotating

Tank Experiment (LME). NASA CR—

1998-208667, 1998.

Lewis contact:

David J. Chato, (216) 977–7488,

David.Chato@grc.nasa.gov

Author: David J. Chato

Headquarters program office: OSS

Programs/Projects:

Microgravity Science, HEDS, In-STEP,

Space Transportation R&T

The NASA Lewis Research Center, in conjunction with Boeing North

American, has been supporting the High Energy Upper Stage (HEUS)

program by performing feasibility studies at Lewis’ Supplemental Multilayer Insulation Research Facility (SMIRF). These tests were performed to

demonstrate the feasibility of chilling and filling a tank with liquid hydrogen in under 5 minutes.

The goal of the HEUS program is to release a satellite from the shuttle

cargo bay and then use a cryogenic (high-energy) upper stage to allow

the satellite to achieve final orbit. Because of safety considerations, the

propellant tanks for the upper stage will be launched warm and dry. They

will be filled from the shuttle’s external tank during the mission phase

after the solid rocket boosters have jettisoned and prior to jettison of the

external tank. Data from previous shuttle missions have been analyzed to

ensure that sufficient propellant would be available in the external tank

to fill the propellant tank of the proposed vehicle upper stage. Because of

mission time-line considerations, the propellant tanks for the upper stage

will have to be chilled down and filled in approximately 5 minutes.

An existing uninsulated flight weight test tank was installed inside the

vacuum chamber at SMIRF, and the chamber was evacuated to the

10–5 torr range to simulate space vacuum conditions in the cargo bay

Rapid Chill and Fill of a Liquid Hydrogen Tank Demonstrated

with the doors open. During prerun

operations, the facility liquid hydrogen (LH2) supply piping was prechilled with the vent gas bypassing

the test article. The liquid hydrogen supply dewar was saturated at

local ambient pressure and then

pressurized with ambient temperature gaseous helium to the test

pressure. A control system was

used to ensure that the liquid

hydrogen supply pressure was

maintained at the test pressure.

During the initial phase of the test,

the test tank vent valve was open.

Despite the open valve, pressure in

the test article rose rapidly until it

approached the supply dewar

pressure. During this initial pressure rise, the warm test tank walls

were sprayed with liquid. As the
111

Research and Technology

0 50 100 150 200 250 300

0

100

200

300

400

500

600

Temperature, °R

Bottom of tank 

Top of tank 

Elapsed time, sec

walls reached near liquid temperatures, the quantity of gas evolved

began to decrease and test tank

pressure began to drop. When the

test tank pressure had fallen close

to atmospheric pressure, the tank

vent valve was closed and the fill

process was allowed to continue.

The purpose of closing the vent

valve was to demonstrate the

no-vent fill process, which conserves propellant and yields a pressurized propellant tank when the

fill is complete. See the graphs for

the pressure, fill level, and temperature traces for a typical test.

Four demonstrations of the rapid

chill and fill process were completed at the SMIRF facility. All tests

resulted in greater than 85-percent

fills obtained in under 5 minutes.

Some parametric studies were

performed to provide a relationship

between the liquid hydrogen supply dewar pressure and the fill

efficiency. The timing of the test

tank vent-valve closure was also

studied. Testing at the SMIRF facility demonstrated the feasibility of

the rapid chill and fill technique

and reduced the technology risks

for the High Energy Upper Stage

program.

Lewis contact:

Maureen T. Kudlac, (216) 977–7476,

Maureen.T.Kudlac@grc.nasa.gov

Author: Maureen T. Kudlac

Headquarters program office: OSF

Programs/Projects: HEUS

Typical tank pressure and liquid level versus elapsed time.

Typical tank temperatures versus elapsed time.

Tank pressure 

Percent full 

Elapsed time, sec

0

10

20

30

40

50

0

20

40

60

80

100

Pressure, psia

Percent full

0 50 100 150 200 250 300

Tank 

vented

Tank chill down

Vent valve closed

No-vent

fill
112

Tests conducted at the NASA Lewis Research Center’s Supplemental

Multilayer Insulation Research Facility (SMIRF) demonstrated that a hybrid

thermal control system could eliminate boiloff of cryogenic propellants.

This is significant because of the substantial mass and cost savings that

could be achieved for any long-duration space mission that requires

cryogenic propellants.

With long-duration cryogenic storage, propellants will boil off because of

the environmental heating of the tank. To accommodate these losses, extra

propellant is required along with larger propellant tanks. Analysis of Mars

mission scenarios using space-transfer cryogenic stages showed that significant savings in propellant mass and tank size could be achieved if it

Zero Boiloff Storage of Cryogenic Propellants Achieved at

Lewis’ Supplemental Multilayer Insulation Research Facility

Test article prior to installation into the vacuum chamber. Mated to the cryocooler

is a condenser (not shown) which was designed to transmit heat entering the

tank by condensing hydrogen vapor. This extended into the ullage of the liquid

hydrogen tank.

were possible to eliminate or

significantly reduce propellant

boiloff.

Engineers and technicians at NASA

Lewis designed, built, and tested a

hybrid thermal control system to

eliminate or significantly reduce

cryogenic propellant boiloff. The

system consists of an active cryocooling system using a cryocooler

in addition to the traditional passive thermal insulation, as shown

in the photo.

Testing conducted over the summer of 1998 compared the performance of the hybrid system to

that of the traditional passive-only

insulation solution. The hardware

consisted of a pressurized 50-ft3

tank insulated with 34 layers of

multilayer insulation (MLI), a condenser, and a Gifford/McMahon

Cryocooler that has a cooling

capacity of 15 to 17.5 watt (W).

Liquid hydrogen was the test fluid.

The test tank was installed into a

vacuum chamber, simulating space

vacuum.

First, the passive insulation system

was tested to determine the boiloff

under steady-state conditions. The

measured steady-state boiloff rate

was approximately 0.12 kg/hr,

equating to a heating rate of

approximately 14.5 W, slightly

higher than the prediction of 13 W

and the historic data of 11.5 W

(1978 tests). Even so, the heating

rate was less than the cryocooler

capacity (15 to 17.5 W).

Next, a series of tests was conducted with the cryocooler in operation. Late into these tests, after

the chamber vacuum pressure

finally inched upward to approximately 8.5310–5 torr, which was

still a full one-half decade worse

Insulated

tank

Cryocooler
113

Research and Technology

than that of our previous test, we achieved zero boiloff (0.00 kg/hr). At this

time, the test tank pressure and temperatures decreased. This tank pressure decrease was approximately 0.07 psi/hr, and the condenser temperature decrease was approximately 0.12 K/hr. The testing was terminated

when the tank pressure dropped near atmospheric pressure. This was to

prevent air from being accidentally drawn into the tank.

These tests confirmed that it is possible to achieve zero boiloff with cryogenic propellant storage using a hybrid thermal control system, and they

provided the foundation for the more detailed analysis and testing needed

to support future long-duration space missions.

Lewis contact:

David W. Plachta, (216) 977–7126,

David.W.Plachta@grc.nasa.gov

Author: David W. Plachta

Headquarters program office: OSF

(Advanced Projects), AST at MSFC

Programs/Projects:

Propulsion Systems R&T, AST

Researchers at the NASA Lewis Research Center have devised an objectoriented architecture for integrating engineering application software. The

architecture provides an environment in which the complete engineering

design and analysis process can be conducted. It eliminates the need to

repeatedly adapt graphical user interfaces (and other tools) to various

experimental and analytical information sources, and it brings those

sources into a common domain to facilitate the easy (and, eventually,

automatic) sharing of engineering content. This architecture has been

implemented in the C++ language, and a graphical user interface interacting with the architecture has been developed with a portable tool kit.

The structure, depicted in the figure, consists of an application object

(labeled PacAppl in the figure) from which emanates (1) a tree of data

identifications (the Pid items to the right), (2) a tree of data configurations

(the PacCfg items extending toward the lower left), and (3) a map of

available operations (the Op items to the upper left). The structure of the

Engineering Application Integration Architecture Devised

Generic engineering application integration architecture.

PacAppl

PacCfg

PacCfg PacCfg PacCfg

PacCfg PacCfg PacCfg

Par: C/Inl/Cfd

Par: A/Inl/Cfd Par: E/Duc/Cfd

Par: B/Inl/Cfd Par: D/Duc/Cfd Par: F/Noz/Cfd

Par: E/Duc/Cfd Par: E/Duc/Cfd Par: E/Duc/Cfd

Par: F/Noz/Cfd

Par: A/Inl/Cfd Par: A/Inl/Cfd Par: A/Inl/Cfd

Op: Kill

Op: Init Op: Run

Pid: Cfd

Pid: Duc Pid: Inl Pid: Noz

Pid: D Pid: E Pid: A Pid: B Pid: C Pid: F

data is revealed by the structure of

the identification tree, eliminating

the need to replicate data within

configurations merely to reveal

structure. Missing data elements of

a configuration are then inherited

from the most immediate precursor. Encapsulated operations, as

well as data identifications, operate

within the context of a current data

configuration and, thus, are able to

refuse operations not supported by

the state of the data.

All the object classes of the architecture derive from a common
114

base class that can provide descriptions of the objects when desired.

Names, units, annotations, access controls, and the like can be attached

on an object-by-object basis. Key among these is automatic value change

histories and a text description, which can be used as a basis for engineering journals.

More information about the engineering application integration architecture is

available on the World Wide Web: http://www.grc.nasa.gov/WWW/price000/

Lewis contact:

Dr. William H. Jones, (216) 433–5862,

William.H.Jones@grc.nasa.gov

Author: Dr. William H. Jones

Headquarters program office: OAT

Programs/Projects:

Information Systems Base R&T

The engineering research and design requirements of today pose great

computer-simulation challenges to engineers and scientists who are called

on to analyze phenomena in continuum mechanics. The future will bring

even more daunting challenges, when increasingly complex phenomena

must be analyzed with increased accuracy. Traditionally used numerical

simulation methods have evolved to their present state by repeated incremental extensions to broaden their scope. They are reaching the limits of

their applicability and will need to be radically revised, at the very least, to

meet future simulation challenges.

At the NASA Lewis Research Center, researchers have been developing a

new numerical framework for solving conservation laws in continuum

mechanics, namely, the Space-Time Conservation Element and Solution

Element Method, or the CE/SE method. This method has been built from

fundamentals and is not a modification of any previously existing method.

It has been designed with generality, simplicity, robustness, and accuracy

as cornerstones.

The CE/SE method has thus far been applied in the fields of computational

fluid dynamics, computational aeroacoustics, and computational electromagnetics. Computer programs based on the CE/SE method have been

developed for calculating flows in one, two, and three spatial dimensions.

Results have been obtained for numerous problems and phenomena,

including various shock-tube problems, ZND detonation waves, an implosion and explosion problem, shocks over a forward-facing step, a blast

wave discharging from a nozzle, various acoustic waves, and shock/

acoustic-wave interactions. The method can clearly resolve shock/acousticwave interactions, wherein the difference of the magnitude between the

acoustic wave and shock could be up to six orders. In two-dimensional

flows, the reflected shock is as crisp as the leading shock. CE/SE schemes

are currently being used for advanced applications to jet and fan noise

prediction and to chemically reacting flows.

Key features and advantages of the CE/SE method follow.

• Space and time are treated in a unified fashion. The space-time domain

is discretized into Solution Elements (SE’s), within which the numerical

approximation is a simple (linear, for example) function of space and time.

Conservation Elements (CE’s) that fill the space-time domain without

overlap are also defined.

Space-Time Conservation Element and Solution Element

Method Being Developed

• The main emphasis is on solving

the integral form of the conservation law in the space-time domain,

although the differential form is

also considered. Conservation of

space-time flux is enforced for each

CE. Because the CE’s fill space-time

without overlap, and because the

flux through any face of a CE is

uniquely defined, this local conservation of flux ensures that spacetime flux is conserved globally in

the space-time domain.

• The method is very simple. It

uses only the simplest of approximation techniques. In addition, no

knowledge of the properties of the

solution, such as the characteristics

or the shock-wave profile, is used.

• The flux-based nature of the

method leads to the use of fluxbased boundary conditions. The

nonreflecting boundary conditions

needed for practical computations

on unbounded spatial domains

are remarkable for their simplicity,

needing only a few lines in the

computer program. These conditions allow even shock waves

to pass out of the domain with

no noticeable reflection at the

boundary.

• The CE/SE method is genuinely

multidimensional; that is, no directional splitting is employed in modeling multidimensional problems.
115

Research and Technology

Snapshots in time (nondimensionalized) of the implosion/explosion of a hexagonal shock in a square

box. Pressure ratio, 10.

2

1

0

–1

–2

–2 –1 0

Time = 0.384

y

1 –2 –1 0

Time = 0.488

1 –2 –1 0

Time = 0.384

1

2

1

0

–1

–2

–2 –1 0

Time = 1.172

y

1 –2 –1 0

Time = 1.288

1 –2 –1 0

Time = 2.485

1

2

1

0

–1

–2

–2 –1 0

Time = 3.188

y

1 –2 –1 0

Time = 3.457

1 –2 –1 0

x

x

x

Time = 4.175

1

2

2

2

• The two- and three-dimensional spatial meshes employed by the present

method are built from triangles and tetrahedrons, respectively. Triangles

and tetrahedrons, respectively, are also the simplest building blocks for

two- and three-dimensional unstructured meshes. Thus, the multidimensional schemes can be directly applied on unstructured meshes, which are

the only efficient way to deal with complex geometries.

• The simplicity of the CE/SE schemes makes the computer programs easy

to vectorize and parallelize.

Find out more about this research on

the World Wide Web:

http://www.grc.nasa.gov/WWW/

microbus/

Bibliography

Chang, S.C., et al.: The Method of

Space-Time Conservation Element and

Solution Element—A New Paradigm for

Numerical Solution of Conservation
116

Laws. To appear in Computational Fluid Dynamics Review, M.M. Hafez and

K. Oshima, eds., John Wiley and Sons, West Sussex, U.K.

Chang, S.-C.: The Method of Space-Time Conservation Element and Solution

Element—A New Approach for Solving the Navier-Stokes and Euler Equations.

J. Comput. Phys., vol. 119, 1995, pp. 295–324.

Lewis contact:

Dr. Philip C.E. Jorgenson, (216) 433–5386, Philip.C.Jorgenson@grc.nasa.gov

Authors: Dr. Sin-Chung Chang, Dr. Ananda Himansu, Dr. Philip C.E. Jorgenson,

Dr. Ching-Yuen Loh, Dr. Xiao-Yen Wang, and Dr. Sheng-Tao Yu

Headquarters program office: OAT

Programs/Projects: Propulsion Systems R&T, AST
117

Research and Technology

Advanced materials are being developed for use in high-temperature gas

turbine applications. For these new materials to be fully utilized, their

deformation properties, their nondestructive evaluation (NDE) quality

and material durability, and their creep and fatigue fracture characteristics

need to be determined by suitable experiments. The experimental findings

must be analyzed, characterized, modeled and translated into constitutive

equations for stress analysis and life prediction. Only when these ingredients—together with the appropriate computational tools—are available, can

durability analysis be performed in the design stage, long before the

component is built.

For example, for NDE information to be useful in structural characterization

and modeling, the NDE data format must be compatible with microstructural and structural models currently being developed (refs. 1 and 2). In

addition, qualitative and quantitative computer analysis tools based on

NDE imaging modalities must be developed to enhance the usefulness

of NDE applications. Qualitative tools include two- and three-dimensional

visualization methods. Quantitative tools include segmentation methods

that can send output to commercial finite element, micromechanical,

and/or continuum damage model software (refs. 3 and 4) for evaluation

of composite materials and components. Linking NDE data with engineering analysis methods will provide the engineering community the

great ability to do a complete structural analysis on as-manufactured or

as-inspected components rather than solely on as-designed components.

With this capability, extensive effect-of-defect studies can be performed

to determine the effect of manufacturing anomalies or in-service component degradation on part performance. This will provide an NDE capability

for engineering structures and condition-based maintenance.

One of the many structural components being evaluated by the NDE

group at the NASA Lewis Research Center is the flywheel system. It is

being considered as an energy storage device for advanced space

vehicles. Such devices offer advantages over electrochemical batteries

in situations demanding high power delivery and high energy storage per

unit weight. In addition, flywheels have potentially higher efficiency and

longer lifetimes with proper motor-generator and rotor design (ref. 5).

Flywheels made of fiber-reinforced polymer composite material show

great promise for energy applications because of the high energy and

power densities that they can achieve (ref. 6) along with a burst failure

mode that is relatively benign in comparison to those of flywheels made

of metallic materials (ref. 7).

However, the big challenge of developing materials for high-energy flywheel systems that can withstand the stresses caused by high rotational

speeds remains eminent. For example, a typical 16-in. steel or titanium

flywheel can shatter at 20 000 rpm. However, flywheels made out of composites show good promise since they can store 20 to 30 times more

Nondestructive Evaluation Correlated With

Finite Element Analysis

power per unit/weight than leadacid batteries, can be recharged

repeatedly, and can handle a wide

range of temperature variations.

They can also be respun thousands

of times to regain their stored

energy.

Therefore, to help improve durability and reduce structural uncertainties, we are developing a comprehensive analytical approach to predict the reliability and life of these

components under these harsh

loading conditions. The combination of NDE and two- and threedimensional finite element analyses (e.g., stress analyses and

fracture mechanics) is expected

to set a standardized procedure to

accurately assess the applicability

of using various composite materials to design a suitable rotor/

flywheel assembly. Following the

figures is a brief description of

some preliminary analytical results.

The figures represent typical NDEfinite element results. The top left

figure shows a computed tomography (CT) scan for a rotor made of

polymer matrix composite. This

scan illustrates the defects in the

rotor due to centrifugal loading

(spun at 34 000 rpm); a crack

along the circumferential direction

is very obvious. The bottom left

figure shows the finite element

models for the rotor assembly. This

assembly consists of the composite

rotor (Hexel’s IM–7 and AS4D) and

an aluminum hub (7075–T6). Two

and three-dimensional models

were generated (ref. 8) because of

analysis requirements. Results

obtained via finite element analyses (ref. 9) are shown in the figures

Structures and Acoustics Division
118

Radial

stress,

psi

324

–110

–543

–977

–1410

–1844

–2278

–2711

–3145

–3578

–4012

–4446

–4879

–5313

–5746

–6180

Y

Z

X

Location of

crack initiation Crack detected 

by CT scan

Radial

stress,

psi

13 149

11 497

9 844

8 192

6 539

4 887

3 234

1 582

–71

–1 723

–3 376

–5 028

–6 681

–8 333

–9 986

–11 638

Y

Z X

Y

Z

X

Left: Computer tomograph (CT) of a flywheel. Right: Radial stresses in a rotor.

Left: Three-dimensional finite element model; 6747 nodes, 5424 hexagonal elements. Right: Radial

stresses in a rotor-hub assembly. These figures are shown in color in the online version of this article

(http://www.grc.nasa.gov/WWW/RT1998/5000/5920aziz.htm).
119

Research and Technology

on the right, where the stress levels depend on the applied loading. Furthermore, the bottom right figure illustrates the radial stress distribution

in the rotor whereby the tensile stresses near the middle indicate the

presence of the crack as detected by the computed tomography scanning

shown in the top figure. Although these results are considered to be

preliminary, they reflect a step forward in correlating NDE findings with

finite element analysis.

For more information, visit Lewis’ Life Prediction Branch on the World Wide

Web: http://www.grc.nasa.gov/WWW/LPB/

References

1. Baaklini, G.Y., et al.: Structural Characterization of Metal Matrix Composites

Using NDE Data. HITEMP Review 1997, NASA CP–10192, Vol. II, 1997, paper

no. 41, pp. 1–15. (Available from NASA Lewis‘ Subsonic Systems Office.)

2. Baaklini, G.Y., et al.: Nondestructive Evaluation of Titanium Alloy MMC Rings for

Gas Turbine Engines. MMC Life System Development (Phase I)—A NASA/Pratt &

Whitney Life Prediction Cooperative Program. E.V. Zaretsky, ed., NASA RP–1361,

1996, pp. 61–73.

3. Wilt, T.E.: Arnold, S.M.; and Goldberg, R.: Micromechanics Analysis Code, MAC

Features and Applications. HITEMP Review 1997, NASA CP–10192, Vol. II, 1997,

paper no. 30, pp. 1–13. (Available from NASA Lewis‘ Subsonic Systems Office.)

4. Arnold, S.M.; and Wilt, T.E.: Deformation and Life Prediction of Circumferentially

Reinforced SCS–6/Ti–15–3 Ring. MMC Life System Development (Phase I)—

A NASA/Pratt & Whitney Life Prediction Cooperative Program. E.V. Zaretsky, ed.,

NASA RP–1361, 1996, pp. 51–60.

5. Ashley, S.: Flywheels Put a New Spin on Electric Vehicles. Mech. Engrg., vol. 115,

no. 10, Oct. 1993, pp. 44–51.

6. Olszewski, M.; et al.: On the Fly or Under Pressure. Mech. Engrg., vol. 110, no. 6,

Jun. 1988, pp. 50–58.

7. Coppa. A.P.: Flywheel Containment

and Safety Considerations. An

Assessment of Integrated Flywheel

System Technology. NASA CP–2346,

1984, pp. 243–264.

8. The MacNeal-Schwendler Corporation: MSC/PATRAN Graphics and

Finite Element Package. Vols. I

and II, MacNeal-Schwendler,

Costa Mesa, CA, 1997.

9. MARC Analysis Research Corporation: MARC General Purposes Finite

Element Analysis program. Vol. A:

User Information manual, and Vol. F:

Theoretical Manual. MARC Analysis

Research Corporation, Palo Alto,

CA, 1996.

Lewis contact:

Ali Abdul-Aziz, (216) 433–6729,

Ali.Abdul-Aziz@grc.nasa.gov

Authors:

Ali Abdul-Aziz and George Y. Baaklini

Headquarters program office: OAT

Programs/Projects:

Aeronautics Base R&T, HITEMP, ACESE

Micromechanics-based modeling of composite material behaviors requires

an accurate assessment of the constituent properties and behaviors. For

the specific case of continuous-fiber-reinforced metal matrix composites

(MMC’s) manufactured from a foil/fiber/foil process, much emphasis has

been placed on characterizing foil-based matrix materials that have been

fabricated in the same way as the composite. Such materials are believed

to yield mechanical properties and behaviors that are representative of the

matrix constituent within the composite (in situ matrix). Therefore, these

materials are desired for micromechanics modeling input. Unfortunately,

such foils are extremely expensive to fabricate and procure because of the

labor-intensive rolling process needed to produce them. As a potential

solution to this problem that would maintain appropriately representative

in situ properties, the matrix constituent could be characterized with sheetbased materials, which are considerably less expensive to manufacture

Deformation Behaviors of HIPped Foil Compared With Those

of Sheet Titanium Alloys

than foils, are more readily procured, and result in fewer plies to

obtain a desired panel thickness.

The critical question is, however,

does the consolidated sheet

material exhibit the same properties and behaviors as do the

consolidated foils? Researchers at

NASA Lewis Research Center’s Life

Prediction Branch completed a

detailed experimental investigation

to answer this question for three

titanium alloys commonly used in

metal matrix composite form.
120

Time, min

0 400 800 1200 1600

Stress, MPa

200

250

300

350

400

450

500

550

Foil-based

Sheet-based

1100

Cost,

$/lb

90

Time, min

0 20 40 60 80 100

Creep strain, m/m

0.00

0.01

0.02

0.03

0.04

0.05

0.06 Creep stress = 552 MPa Strain hold = 0.007 m/m

The experimental investigation compared the 427 °C (800 °F) mechanical properties and deformation behaviors of three HIPped (Hot Isostatic

Pressed) foil and HIPped sheet titanium matrix materials commonly used

in silicon-carbide- (SiC-) reinforced titanium composites. The alloys investigated were Ti-15V-3Cr-3Al-3Sn (Ti-15-3), Ti-15Mo-3Nb-3Al-0.2Si (Ti-21S),

and Ti-6Al-4V (Ti-6-4). Elastic properties, creep deformation, and stress

relaxation were examined along with the microstructural features before

and after deformation. Differences in behavior were judged on the basis

of statistical significance, where both a univariate analysis of variance

(ANOVA) and a multivariate analysis of variance (MANOVA) were used and

a two-parameter Norton-Bailey power law relationship was employed. In

general, the HIPped foil and sheet were found to be significantly different

in creep and stress relaxation at the 95-percent confidence level, with the

only exceptions being Ti-15-3 in creep and Ti-21S in stress relaxation.

Influencing this conclusion was the fact that the behaviors for any one

alloy/product-form combination tended to be tightly grouped and exhibited relatively low sample-to-sample functional deviations. Of the three

Creep and stress relaxation behavior of HIPped foil and sheet Ti-6-4 matrix at

427 °C (800 °F).

alloys, only the Ti-15-3 differed

significantly within the microstructure in comparison to the foil and

sheet forms. Furthermore, this

was the only alloy to exhibit differences between the pre- and postdeformation states. This resulted

from the metastable condition of

the Ti-15-3 (even after a stabilization heat treatment). At 427 °C

(800 °F), this alloy tends to

experience a notable degree of

deformation-assisted α-Ti precipitation within the β matrix. Although

the strictly interpreted statistical

analysis indicated that the foils

and sheets exhibited significantly

different behaviors, often a more

practical engineering assessment

and interpretation of the behaviors

suggested that the sheet material

could, in fact, be substituted for the

foil, depending on the intended

application of the data.

Lewis contacts:

Michael G. Castelli, (216) 433–8464,

Michael.G.Castelli@grc.nasa.gov; and

Dr. Bradley A. Lerch, (216) 433–5522,

Bradley.A.Lerch@grc.nasa.gov

Author: Michael G. Castelli

Headquarters program office: OAT

Programs/Projects:

Propulsion Systems R&T, HITEMP

High-performance polymer matrix composites (PMC’s) continue to be the

focus of a number of research efforts aimed at developing cost-effective,

lightweight material alternatives for advanced aerospace and aeropropulsion applications. These materials not only offer significant advantages in

specific stiffness and strength over their current metal counterparts, but

they can be designed and manufactured to eliminate joints and fasteners

by combining individual components into integral subassemblies, thus

making them extremely attractive for commercial applications. With much

emphasis on the low-cost manufacturing aspects of advanced composite

structures, there is heightened interest in high-performance sheet-molding

compounds (SMC’s). SMC’s effectively reduce the costs associated with

Thermomechanical Fatigue Durability of T650–35/PMR–15

Sheet-Molding Compound Evaluated

component production that uses

prepregs, where variable costs

are generally associated with labor,

secondary processes, and scrap.

With compression molding, SMC’s

can be fabricated into complex

shapes facilitating the use of

simple charge patterns, part consolidation, and molded-in inserts,

which reduce labor, equipment,

and operation costs for preparatory

and secondary processes.
121

Research and Technology

Time, min

0 10 20 30 40 50

Temperature, °C

0

50

100

150

200

250

Stress, MPa

0

50

100

150

200 Temperature

Stress

Idle

Idle

Maximum

Maximum

Redline

Redline

Researchers at the NASA Lewis Research Center, in cooperation with the

Allison Advanced Development Company, completed an investigation

examining the use of T650–35/PMR–15 SMC for a midstage inner-vane

endwall application within a gas turbine engine compressor. This component resides in the engine flow path and is subjected not only to high

airflow rates, but also to elevated temperatures and pressures. This

application is unique in that it represents a very aggressive use of highperformance SMC’s, raising obvious concerns related to durability and

property retention in the presence of microstructural damage. Therefore,

it was necessary to evaluate the fatigue behavior and damage tolerance

of this material subjected to a representative thermomechanical fatigue

(TMF) mission-cycle loading spectrum.

Damage progression was tracked through changes in the macroscopic

deformation and elastic stiffness in the loading direction. Additional

properties, such as the glass transition temperature and dynamic mechanical response also were examined. The fiber distribution orientation was

characterized through a detailed quantitative image analysis, and material

durability and damage tolerance were quantified on the basis of residual

static tensile properties after a prescribed number of TMF missions.

Detailed microstructural examinations used optical and scanning electron microscopy to characterize the local damage.

Thermomechanical fatigue mission cycle used for the T650–35/PMR–15 SMC

inner-vane endwall application, showing idle, maximum, and redline (overload)

conditions for temperature and stress.

Results indicate that the imposed

TMF missions had only a modest

effect on material durability as

measured by the mechanical

properties. Some microstructural

damage was observed subsequent

to 100 hours of TMF cycling. It

consisted primarily of fiber debonding and transverse cracking local

to predominantly transverse fiber

bundles. No statistically significant degradation occurred in the

residual tensile properties. Some

of the more aggressive TMF scenarios examined, however, did

promote notable creep damage

and excessive strain accumulation

that led to rupture. In some cases

this creep behavior occurred at

temperatures in excess of 150 °C

below commonly cited values for

the glass transition temperature. As

a result, thermomechanical exploratory creep tests were conducted.

These revealed that the SMC was

subject to time-dependent deformation at stress/temperature

thresholds of 150 MPa/230 °C and

170 MPa/180 °C.

Bibliography

Castelli, M.G.; Sutter, J.K.; and Benson,

D.: Durability and Damage Tolerance

of a Polyimide/Chopped Fiber Composite Subjected to Thermomechanical

Fatigue Missions and Creep Loadings.

Time Dependent and Nonlinear Effects

in Polymers and Composites. ASTM

STP–1357, R.A. Schapery and C.T. Sun,

eds., American Society for Testing and

Materials, 1999.

Lewis contacts:

Michael G. Castelli, (216) 433-8464,

Michael.G.Castelli@grc.nasa.gov; and

Dr. James K. Sutter, (216) 433-3226,

James.K.Sutter@grc.nasa.gov

Author: Michael G. Castelli

Headquarters program office: OAT

Programs/Projects:

Propulsion Systems R&T, HITEMP
122

100

10–1

10–2 100 102 104 106

Reduced strength, 

sr*

Stress rate, s, MPa/sec

Conventional

test rate

Material

GN10

GN10 composite

NCX34

SN252

NC132

NC203

96 wt % alumina

Test

temperature, °C

1300

1300

1200

1371

1100

1300

1000

100

10–1

10–2 100 102 104 106

Reduced strength, 

sr*

Stress rate, s, MPa/sec

Conventional

test rate

Material

A246

AS440

NT154

Kryptonite

AS800

SNW1000

a-SiC

Test

temperature, °C

1200

1100

1200

1200

1200

1300

1371

The accurate determination of inert strength is important in reliable life

prediction of structural ceramic components. At ambient temperature, the

inert strength of a brittle material is typically regarded as free of the effects

of slow crack growth due to stress corrosion. Therefore, the inert strength

can be determined either by eliminating active species, especially moisture,

with an appropriate inert medium, or by using a very high test rate.

However, at elevated temperatures, the concept or definition of the inert

strength of brittle ceramic materials is not clear, since temperature itself is

a degrading environment, resulting in strength degradation through slow

crack growth and/or creep. Since the mechanism to control strength is

rate-dependent viscous flow, the only conceivable way to determine the

inert strength at elevated temperatures is to utilize a very fast test rate that

either minimizes the time for or eliminates slow crack growth. Few experimental studies have measured the elevated-temperature, inert (or “ultra”-

fast fracture) strength of advanced ceramics. This is, in part, because conventional test frames are incapable of very high rate testing. In addition,

data acquisition systems are inadequate, and there are safety concerns.

Commonly, a maximum test rate of about 30 MPa/sec has been used to

determine the elevated-temperature strength of ceramic materials. The

strength that is determined at this test rate has been called the fast-fracture

strength, implying that it is the maximum attainable, or ultimate, strength

of the material at the temperature.

At the NASA Lewis Research Center, an experimental study was initiated

to better understand the “ultra”-fast fracture strength behavior of advanced

ceramics at elevated temperatures. Fourteen advanced ceramics—one

alumina, eleven silicon nitrides, and two silicon carbides—have been tested

to date using constant stress-rate

(dynamic fatigue) testing in flexure

with a series of stress rates including the “ultra”-fast stress rate of

33 000 MPa/sec with digitally controlled test frames. The experimental results for these 14 advanced

ceramics indicate that, notwithstanding possible changes in flaw

populations as well as in flaw configurations because of elevated

temperatures, the strength at

33 000 MPa/sec approached the

room-temperature strength or

reached a higher value than that

determined at the conventional

test rate of 30 MPa/sec. The graphs

summarize the test results for the

14 ceramics, where the reduced

strength is an elevated-temperature

strength normalized with respect

to the corresponding roomtemperature strength. On the

basis of the experimental data, it

can be stated that the elevatedtemperature, inert strength of an

advanced ceramic material can be

defined as the strength where no

“Ultra”-Fast Fracture Strength of Advanced Structural

Ceramic Materials Studied at Elevated Temperatures

Summary of reduced strength as a function of stress rate for 14 advanced ceramics. Left: GN10, GN10 composite,

NCX34, SN252 and NC132 silicon nitrides, NC203 silicon carbide, and 96 wt % alumina; Right: A2Y6, AS440,

NT154, “Kryptonite” composite, AS800 and SNW1000 silicon nitrides, and α-silicon carbide.
123

Research and Technology

slow crack growth takes place at the temperature. Specifically, the elevatedtemperature inert strength is close to its room-temperature counterpart

and can be obtained via a series of “ultra”-fast stress rates, including

33 000 MPa/sec in many cases. The strength determined at 33 000 MPa/

sec must be used as an inert strength to determine the required life prediction parameters of the material.

Bibliography

Choi, S.R.; and Gyekenyesi, J.P.: Elevated-Temperature, ‘Ultra’-Fast Fracture Strength

of Advanced Ceramics: An Approach to Elevated-Temperature Inert Strength. ASME

Paper 98–GT–479, 1998.

Choi, S.R.; and Gyekenyesi, J.P.: Elevated Temperature ‘Ultra’-fast Fracture Strength

Behavior of Advanced Structural Ceramics. Ceramic Materials and Components for

Engines, K. Niihara, et al., eds., Technoplaza Co., Japan, 1998, pp. 689–694.

Choi, S.R.; and Salem, J.A.: ‘Ultra’-Fast Fracture Strength of Advanced Ceramics

at Elevated Temperature. Mater. Sci. Eng., A242, 1998, pp. 129–136.

Lewis contacts:

Dr. Sung R. Choi, (216) 433–8366,

Sung.R.Choi@grc.nasa.gov; and

Dr. John P. Gyekenyesi, (216) 433–

3210, John.P.Gyekenyesi@grc.nasa.gov

Authors:

Dr. Sung R. Choi and

Dr. John P. Gyekenyesi

Headquarters program office: OAT

Programs/Projects: HITEMP, P&PM

Development of accurate three-dimensional (multiaxial) inelastic stressstrain models is critical in utilizing advanced ceramics for challenging 21st

century high-temperature structural applications. The current state of the

art uses elastic stress fields as a basis for both subcritical crack growth and

creep life prediction efforts aimed at predicting the time-dependent reliability response of ceramic components subjected to elevated service

temperatures. However, to successfully design components that will meet

tomorrow’s challenging requirements, design engineers must recognize

that elastic predictions are inaccurate for these materials when subjected

to high-temperature service conditions such as those encountered in

advanced heat engine components. Analytical life prediction methodologies developed for advanced ceramics and other brittle materials must

employ accurate constitutive models that capture the inelastic response

exhibited by these materials at elevated service temperatures.

A constitutive model recently developed at the NASA Lewis Research

Center helps address this issue by accounting for the time-dependent

(inelastic) material deformation phenomena (e.g., creep, rate sensitivity,

and stress relaxation) exhibited by monolithic ceramics exposed to hightemperature service conditions. In addition, the proposed formulation is

based on a threshold function that is sensitive to hydrostatic stress and

allows different behavior in tension and compression, reflecting experimental observations obtained for these material systems.

The objective of this effort was to demonstrate the capabilities and inherent features of the mathematical formulation of the constitutive theory. In

this regard, the viscoplastic constitutive equations formulated for the flow

law (i.e., the strain rate) and the evolutionary law were incorporated into

computer algorithms for predicting the multiaxial inelastic (creep) response

of a given homogeneous state of stress. For the solution of a full multiaxial

creep problem, 12 coupled differential equations had to be integrated.

These represent the constitutive

law: that is, six equations from the

flow law (i.e., strain rate) and six

from the evolutionary law.

Examples have been simulated

numerically to illustrate the model’s

ability to qualitatively capture the

time-dependent phenomena

suggested here. For an imposed

service (load) history, the computer

algorithm can generate creep

curves and viscoplastic flow surfaces that demonstrate its ability

to capture the inelastic creep

deformation response. No attempt

was made to assess the accuracy

of the model in comparison to the

experiment. A quantitative assessment has been reserved for a later

date—after the material constants

have been suitably characterized

for a specific ceramic material.

Creep rupture is not considered in

the model, although incorporating

damage mechanics concepts into

the present theory could yield a

workable creep rupture model. This

task also has been reserved for a

future enhancement.

Viscoplastic Constitutive Theory Demonstrated for

Monolithic Ceramic Materials
124

Bibliography

Janosik, L.A.; and Duffy, S.F.: A Viscoplastic Constitutive Theory for Monolithic

Ceramics—I. J. Engrg. Gas Turb. Power (ASME Paper 96–GT–368) vol. 120, no. 1,

Jan. 1998, pp. 155–161.

Lewis contact:

Lesley A. Janosik, (216) 433–5160, Lesley.A.Janosik@grc.nasa.gov

Author: Lesley A. Janosik

Headquarters program office: OAT

Programs/Projects:

Propulsion Systems R&T, P&PM

A team consisting of GE Aircraft Engines, Precision Cast Parts, Oremet, and

Chromalloy were awarded a NASA-sponsored Aerospace Industry Technology Program (AITP) to develop a design and manufacturing capability that

will lead to the engine test demonstration and eventual implementation of

a γ -Ti-47Al-2Nb-2Cr (at. %) titanium aluminide (TiAl) low-pressure turbine

blade into commercial service. One of the main technical risks of implementing TiAl low-pressure turbine blades is the poor impact resistance

of TiAl in comparison to the currently used nickel-based superalloy. The

impact resistance of TiAl is being investigated at the NASA Lewis Research

Center as part of the Aerospace Industry Technology Program and the

Advanced High Temperature Engine Materials Program (HITEMP).

The overall objective of this work is to determine the influence of impact

damage on the high cycle fatigue life of TiAl-simulated low-pressure

turbine blades. To this end, impact specimens were cast to size in a dogbone configuration and given a typical processing sequence followed by

an exposure to 650 °C for 20 hours to simulate embrittlement at service

conditions. Then, the specimens were impacted at 260 °C under a 69-MPa

load. Steel projectiles with diameters 1.6 and 3.2 mm were used to impact

the specimens at 90° to the leading edge. Two different impact energies

(0.74 and 1.5 joules) were used to simulate fairly severe domestic object

damage on a low-pressure turbine blade.

Fatigue tests were performed at 650 °C and at a frequency

of 100 Hz. In addition, three different loading ratios, R, were

used to assess the effect of mean stresses. As expected, the

specimens impacted at the higher energy levels failed at

lower fatigue stresses because of the larger “defect” associated with the impact. Both energy levels resulted in fatigue

strengths that were significantly lower, yet predictable,

than for the smooth bars (i.e., nonimpacted specimens).

In addition, a Goodman mean stress approach could be

used to accurately model the fatigue data for all impacted

specimens.

The industry-NASA team is using the results of this study to

minimize the technical risks associated with impact issues.

The actual damage tolerable for a low-pressure turbine

blade application will be determined by a combination of

fatigue testing and consideration of actual engine conditions. The current evaluations indicate that Ti-47Al-2Nb-2Cr

Resistance of Titanium Aluminide to Domestic Object

Damage Assessed

possesses the level of damage

tolerance required for implementation into service.

Bibliography

Draper, S.L.: Pereira, J.M.; and Nathal,

M.V.: Impact Resistance of γ-Ti-48Al2Nb-2Cr. HITEMP Review 1997, NASA

CP–10192, Vol. II, 1997, paper no. 25,

pp. 1–13. (Available from NASA Lewis‘

Subsonic Systems Office.)

Lewis contact:

Dr. Bradley A. Lerch, (216) 433–5522,

Bradley.A.Lerch@grc.nasa.gov

Authors:

Dr. Bradley A. Lerch, Susan L. Draper,

Dr. J. Michael Pereira, Dr. Michael V.

Nathal, and Curt Austin

Headquarters program office: OAT

Programs and projects:

Aeronautics Base R&T, HITEMP, AITP

Goodman diagram for γ -TiAl specimens tested at 650 °C.

0

10

20

30

40

50

0 1 02 03 04 05 0

Mean stress, ksi

Alternating stress, ksi

0.0

0.2

0.5

High energy 

Low 

energy

Smooth 

bar 

Loading

ratio,

R

0.0

0.2

0.5

Loading ratio,

R

Solid symbols denote 

 high-energy impact

Open symbols denote 

 low-energy impact
125

Research and Technology

0

–1000 –500 0 500 1000

Axial stress, s11, MPa

Shear stress, 

√

3 s12, MPa

Prestrain point

Subsequent locus 

Initial von Mises criterion 

1000

500

–1000

–500

Aeropropulsion components, such as disks, blades, and shafts, are commonly subjected to multiaxial stress states at elevated temperatures.

Experimental results from loadings as complex as those experienced in

service are needed to help guide the development of accurate viscoplastic,

multiaxial deformation models that can be used to improve the design of

these components. Typically, past studies investigated model materials,

concentrating, for experimental simplicity, on room temperature. This study

provides first-of-a-kind data by conducting axial-torsional tests at servicerelated temperatures on a popular aeropropulsion material, namely

INCONEL 718 (IN–718, Inco Alloys International, Inc., Huntington, West

Virginia). The data from this type of study can be used to develop the

evolution equations necessary for this and similar engineering materials.

Yield loci were determined for aged IN–718 at 650 °C. This was done

using experimental procedures developed over the past several years in a

cooperative program between Penn State University and the NASA Lewis

Research Center. This program represents an ongoing effort to develop test

methods necessary for the validation of viscoplastic models. The von Mises

(J2) yield criterion was found to not fit the data well because of a strength

differential in tension and compression. That is, the initial yield strength

in compression was greater than it was in tension. The strength differential was present for the entire range of loading (up to at least strains of

±1 percent), but it decreased as plastic flow increased. In addition, the

strength differential increased as the temperature increased in the range

of 20 to 650 °C.

Yield surfaces also were obtained after preloading the material to various

levels. After applying an axial-shear prestrain to an equivalent inelastic

strain of 500 microstrain, the yield

surface translated in the direction

of the prestrain (see the figure). In

addition, the yield surface elongated in the direction of prestraining and flattened on the

back side. Thus, there is a component of distortional hardening that

needs to be accounted for in the

theoretical evolution equations.

We are presently developing

techniques to transform the yield

surface data to flow surfaces,

which are similar to yield surfaces except that they are ratedependent, and thus more meaningful for viscoplastic models. Since

the von Mises yield criterion was

inadequate for IN–718, experiments and associated modeling

work are planned to determine the

dependence of initial yielding on

the stress invariants (I

1, J2, and J3)

and to establish the importance of

these invariants. In addition, we

plan to examine load path dependence for the evolution of state

and compare these results with

predictions from viscoplastic

models.

Bibliography

Gil, C.M.; Lissenden, C.J.; and Lerch,

B.A.: Determination of Yield in Inconel

718 for Axial-Torsional Loading at

Temperatures up to 649 °C. NASA

TM—1998-208658, 1998.

Gil, C.M.; Lissenden, C.J.; and Lerch,

B.A.: Investigation of Anomalous

Behavior in Metallic-Based Materials

Under Compressive Loading. NASA

TM—1998-206640, 1998.

Lissenden, C.J., et al.: Robinson:

Experimental Determination of Yield

and Flow Surfaces Under AxialTorsional Loading. Multiaxial Fatigue

and Deformation Testing Techniques.

ASTM STP–1280, S. Kalluri and P.J.

Bonacuse, eds., 1997, pp. 92–112.

Multiaxial Experiments Conducted to Aid in the

Development of Viscoplastic Models

Yield locus subsequent to an axial-shear inelastic prestrain of 500-microstrain for

IN–718 aged at 650 °C. Offset for individual probes, εoff, 30 microstrain.
126

Lewis contact: Dr. Bradley A. Lerch, (216) 433–5522, Bradley.A.Lerch@grc.nasa.gov

Authors: Dr. Bradley A. Lerch, Prof. Clifford J. Lissenden, and Christopher Gil

Headquarters program office: OAT

Programs/Projects: HITEMP

High-Speed Civil Transport (HSCT) engine combustor liners are subjected

to complex thermal environments and have to endure these for thousands

of hours with assured reliability. In the past, several deterministic analyses

have been performed, including detailed heat transfer analyses to obtain

thermal profiles and deterministic stress analyses to identify critical locations of high stresses. Actual rig tests also have been performed for segments by simulating these loading situations as closely as possible.

However, it is well known that many uncertainties exist in loading (primarily thermal loads due to heat transfer), boundary conditions (end fixity

unknowns), and material properties (moduli, thermal-expansion coefficients, and conductivities). The present in-house effort at the NASA Lewis

Research Center is directed toward accounting for these in a formal way

to assess the performance of liner

components under complex and

uncertain loading conditions as

well as subject to other geometryand material-related uncertainties.

Under the sponsorship of the

Enabling Propulsion Materials

(EPM) project, Lewis’ Structures

and Acoustics Division recently

developed a computational capability, NESTEM (see the figures),

Multidisciplinary Probabilistic Heat Transfer/Structural

Analysis Code Developed—NESTEM

Sample output from NESTEM for a conical ceramic-matrix-composite liner segment.

Cumulative stress distribution

due to variations in

properties

Sensitivity analyses

0.0

0.2

0.4

0.6

0.8

1.0

Stress, ksi

Cumulative probability

0.0

0.2

0.4

0.6

0.8

Random variables

Sensitivities

Thermal expansion 

coefficient

Radiation 

temperature

Hoop 

modulus

6.0 12.0 14.0 8.0 10.0

Conical liner sector (combustor liner)

Liner material for 

low-NOx combustor
127

Research and Technology

Distribution 

type

x1 x2 x3

Sensitivity factors Response cumulative

distribution function

(CDF)

CDF

z

Random 

variable

statistics Fast probability

integration (FPI)

analysis engine

CEMCAN performance

function z = f(x1, x2, x3) Output 

options

through a contract to Modern Technologies Corporation. NESTEM is a computer code that combines the heat transfer analysis capability of the EPM

backbone computer code CSTEM (Coupled Structural, Thermal and ElectroMagnetic Tailoring) with Lewis’ in-house probabilistic structural analysis

code NESSUS (Numerical Evaluation of Stochastic Structures Under Stress).

The code can now analyze and assess the complex combustor thermal

environment with uncertainties, as well as its effects on the overall ceramic

matrix composite (CMC) liner response. It enables us to formally assess the

uncertainties in loads, material property variations, and geometric imperfections on the overall structural behavior (such as stability, frequencies,

and stresses). Typical output of the code is probabilistic stress distributions

at the hot spots, probabilistic frequencies, and buckling loads. This information allows designers to make more informed judgments regarding the

preliminary design of the liner components without resorting to overly conservative deterministic approaches with ad hoc knockdown factors. The

information also permits more accurate calculation of the reliability and life

of such components. In addition, the code provides information on the

sensitivities of the various uncertainties and ranks them in the order of

importance as a byproduct. The accompanying figures show a sample output for a conical ceramic-matrix-composite liner segment.

NESTEM’s features include

• Automatic geometry and finite element mesh generation

• Finite element property card generation for arbitrary composite layups

• Uncertainties in heat-transfer-related variables, combustor temperatures,

geometry, material properties, and boundary conditions

• Heat transfer and structural

analysis in the same run

• Cumulative distribution functions of response, including

temperatures, stresses, natural

frequencies, buckling loads, and

displacements

• Sensitivities of variables with

uncertainties at various probability levels

Lewis contacts:

Dr. Pappu L.N. Murthy, (216) 433–

3332, Pappu.L.Murthy@grc.nasa.gov;

and Dr. Shantaram S. Pai, (216) 433–

3255, Shantaram.S.Pai@grc.nasa.gov

Authors:

Dr. Pappu L.N. Murthy and

Dr. Shantaram S. Pai

Headquarters program office: OAT

Programs/Projects: HSR, EPM

Special recognition:

Team Achievement Award

Design Tool Developed for Probabilistic Modeling of

Ceramic Matrix Composite Strength

Fast probability integration input and output.

Ceramic matrix composites are

being evaluated as candidate materials for many high-temperature

applications such as engine combustor liners for the High-Speed

Civil Transport (HSCT). They are

required to have an assured life of

several thousand hours. Estimating

the reliability of these components

is quite a complex process and

requires knowledge of the uncertainties that occur at various

scales. The properties of ceramic

matrix composites (CMC) are

known to display a considerable

amount of scatter due to variations

in fiber/matrix properties, interphase/coating properties, bonding,

amount of matrix voids, and many

geometry- and fabrication-related

parameters such as ply thickness

and ply orientations. The objective
128

0.0

0.2

0.4

0.6

0.8

1.0

100 150 200 250 300

Cumulative probability 

PCEMCAN

Experimental

First matrix fracture strength, MPa

of this research effort is to account for these uncertainties in a formal

way by probabilistically analyzing both the stiffness- and strength-related

properties of CMC’s. In current deterministic approaches, uncertainties are

usually accounted for by safety factors. This approach often yields overly

conservative designs, thereby reducing the potential of many advanced

composite materials.

Work is underway at the NASA Lewis Research Center to incorporate the

probabilistic distribution of material-behavior and fabrication-related

parameters into the micromechanics and macromechanics of CMC’s. The

primary objective of this work was to develop an efficient computational

design tool that could account for all the uncertainties in a more rigorous

and formal manner, providing overall composite properties and their

probabilistic distributions.

Therefore, we combined the CMC analysis embedded in the CEMCAN

(Ceramic Matrix Composite Analyzer) computer code with the fast probability integration (FPI) techniques available in the NESSUS (Numerical

Evaluation of Stochastic Structures Under Stress) code. CEMCAN provides

functional relationships that tie the constituent properties and other

geometry- and fabrication-related parameters to the overall composite

properties. Fast probability integration is used to perform probabilistic

analyses by utilizing the properties generated by CEMCAN. The results are

cumulative distribution functions (CDF’s) and probability density functions

(PDF’s) for the ply and laminate properties of the CMC. This technique is

more efficient than a standard Monte-Carlo technique, where a large

number of simulations are needed to generate a cumulative distribution

Comparison of the first matrix cracking strength

cumulative-distribution-function simulation (PCEMCAN,

Probabilistic Ceramic Matrix Composite Analyzer) with

the experimental data for a [0]8 SCS—6/RBSN composite

laminate.

Sensitivity of the first matrix cracking strength to the

primitive random variables of a [0]8 SCS—6/RBSN

composite laminate. Matrix tensile strength, SmT ; fiber

coefficient of thermal expansion, αf11; fiber longitudinal

modulus, Ef11; fiber volume ratio, FVR.

function. The probabilistic sensitivities of the output variables with

respect to inherent scatter in basic

variables are obtained as a byproduct of the fast probability integration analyses. This provides very

useful information to design and

test engineers in evaluating the

importance of variables that control

the scatter in an overall property.

The accompanying figures show

sample results from the analyses.

Lewis contact:

Dr. Pappu L.N. Murthy, (216) 433–

3332, Pappu.L.Murthy@grc.nasa.gov

University of Toledo contact:

Subodh K. Mital, (216) 433–3261,

Subodh.K.Mital@grc.nasa.gov

Authors:

Dr. Pappu L.N. Murthy, Subodh K. Mital,

and Ashwin R. Shah

Headquarters program office: OAT

Programs/Projects: HITEMP

–0.8 –0.6 –0.4 –0.2 0.0 0.2 0.4

Sensitivity

0.999 Probability

0.001 Probability

E

af11

f11

am

SmT

FVR

Em
129

Research and Technology

103

102

100 101 102 103 104 105 106 107 108 109 1010

 0.8

 0.5

 0.0

–1.0

R ratio

Maximum failure, MPa

Time to failure, sec

The CARES/Life computer program predicts the probability of a monolithic

ceramic component’s failure as a function of time in service. The program

has many features and options for materials evaluation and component

design. It couples commercial finite element programs—which resolve a

component’s temperature and stress distribution—to reliability evaluation

and fracture mechanics routines for modeling strength-limiting defects. The

capability, flexibility, and uniqueness of CARES/Life have attracted many

users representing a broad range of interests and has resulted in numerous

awards for technological achievements and technology transfer.

Recent work with CARES/Life was directed at enhancing the program’s

capabilities with regards to cyclic fatigue. Only in the last few years have

ceramics been recognized to be susceptible to enhanced degradation from

cyclic loading. To account for cyclic loads, researchers at the NASA Lewis

Research Center developed a crack growth model that combines the

Power Law (time-dependent) and the Walker Law (cycle-dependent) crack

growth models. This combined model has the characteristics of Power Law

behavior (decreased damage) at high R ratios (minimum load/maximum

load) and of Walker law behavior

(increased damage) at low R ratios.

In addition, a parameter estimation methodology for constantamplitude, steady-state cyclic

fatigue experiments was developed

using nonlinear least squares and

a modified Levenberg-Marquardt

algorithm. This methodology is

used to give best estimates of

parameter values from cyclic

fatigue specimen rupture data

(usually tensile or flexure bar

specimens) for a relatively small

number of specimens. Methodology to account for runout data

(unfailed specimens over the

duration of the experiment) was

also included.

The graph shows an example of

this regression technique for tensile

specimen data for partially stabilized zirconia. Data for various R

ratios and their corresponding

median regression lines are shown

for the combined fatigue law

model. Note that at lower R ratios

(0 and –1) the combined law predicted increased material strength

degradation, whereas at higher R

ratios (0.5 and 0.8) this trend was

reversed.

Find out more on the World Wide

Web: http://www.grc.nasa.gov/WWW/

LPB/cares/life/software.html

Bibliography

Nemeth, N.N., et al.: Designing Ceramic

Components for Durability. Advanced

Ceramic Matrix Composite—Design

Approaches, Testing and Life Prediction Methods, E.R. Generazio, ed.,

Technomic Publishing Company,

Lancaster, PA, 1996, pp. 3–16.

Nemeth, N.N., et al.: Durability Evaluation of Ceramic Components Using

CARES/Life. ASME J. Engrg. Gas Turb.

Power, vol. 118, Jan. 1996, pp. 150–

158.

CARES/Life Ceramics Durability Evaluation Software

Enhanced for Cyclic Fatigue

Median regression lines for partially stabilized zirconia tensile specimens with

various R-ratios (combined law model).
130

Rahman, S.; Nemeth, N.N.; and Gyekenyesi, J. P.: Life Prediction and Reliability

Analysis of Ceramic Structures Under Combined Static and Cyclic Fatigue. ASME

Paper 98–GT–569, 1998.

Lewis contacts:

Noel N. Nemeth, (216) 433–3215, Noel.N.Nemeth@grc.nasa.gov;

Lynn M. Powers, (216) 433–8374, Lynn.M.Powers@grc.nasa.gov; and

Lesley A. Janosik, (216) 433–5160, Lesley.A.Janosik@grc.nasa.gov

Authors: Noel N. Nemeth, Lynn M. Powers, and Lesley A. Janosik

Headquarters program office: OAT

Programs/Projects: P&PM, broad

range of propulsion technology

Special recognition:

Numerous awards including 1998 EDI

Innovation Award (Enterprise Development Inc.) given to Lewis for the

development of CARES

At the NASA Lewis Research Center, a technique for the ultrasonic characterization of plates has been extended to tubes and to curved structures

in general. In this technique, one performs measurements that yield a

thickness-independent value of the local through-the-thickness speed of

sound in a specimen. From such measurements at numerous locations

across the specimen, one can construct a map of velocity as a function of

location. The gradients of velocity indicated by such a map indicate local

through-the-thickness-averaged microstructural parameters that affect the

speed of sound. Such parameters include the pore volume fraction, mass

density, fiber volume fraction (in the case of a composite material), and

chemical composition. Apparatus was designed to apply the technique to

tubular and other curved specimens.

The specimen was mounted on

a horizontal turntable in a water

tank, with its axis vertical and

coincident with the turntable axis.

A machined metal reflector plate

narrow enough to fit within the

inner diameter of the specimen

was suspended vertically from

above and positioned inside the

specimen about 1 cm from the

inner tube wall. A horizontally

oriented ultrasonic transducer was

positioned outside the specimen,

facing the reflector plate. Then,

pulse/echo measurements were

taken in basically the same manner

as for the plate specimens. The

transducer was translated vertically

to obtain measurements at various

axial positions (e.g., increments

of 1 mm), and the turntable was

rotated to obtain measurements at

various azimuthal positions (e.g.,

increments of 1°).

The technique has been demonstrated in experiments on tubular

specimens of mullite (silica/

alumina), a polymer-matrix composite, a composite of SiC fibers

in an SiC matrix, and a hightemperature structural grade of

silicon nitride. Although the turntable, specimen, reflector plate,

and transducer should be aligned

as nearly perfectly as possible and

Novel Method Used to Inspect Curved and Tubular

Structural Materials

Apparent and thickness-independent velocity images of silicon nitride tube as

decaled onto tubular models. Baseline orientation. Left: Apparent velocity

image. Right: Thickness-independent velocity image.
131

Research and Technology

Metallic

interface

ring

Epoxy

overwrap

Silica 

phenolic

liner

the specimen should approximate a perfect round tube, it was observed

that, in general, some misalignment and out of roundness can be tolerated. This is an advantage over peak-amplitude-based ultrasonic techniques in which measurements are altered drastically by refractive effects

associated with out of roundness. The present technique made it possible

to eliminate most of the effects of variations in tube-wall thicknesses upon

velocity maps (through-the-thickness velocities as functions of axial and

azimuthal positions). However, edge effects associated with discontinuous

changes in thickness were not eliminated completely. In the case of the

silicon nitride tube, differences between velocities at different locations

were found to be correlated with differences between densities and pore

volume fractions revealed by x-radiography and destructive metallographic

analysis at those locations. The illustration shows the apparent (without

thickness variation subtracted) and thickness-independent velocity images

of the silicon nitride tube as decaled onto tubular models.

For more information about this research, visit us on the World Wide Web:

http://www.grc.nasa.gov/WWW/LPB/tiui/

Lewis contact:

Dr. Don J. Roth, (216) 433–6017,

don.j.roth@grc.nasa.gov

Authors:

Dr. Don J. Roth, George Y. Baaklini,

Dorothy V. Carney, James R. Bodis, and

Richard W. Rauser

Headquarters program office: OAT

Programs/Projects: HITEMP, P&PM

Special recognition: Part of the effort

that received an R&D100 award and

the Federal Laboratory Consortium

Excellence in Technology Transfer

Award; to be featured on the cover of

the International Journal of Materials

Evaluation in 1998

The Low Cost Booster Technology Program is an initiative to minimize the

cost of future liquid engines by using advanced materials and innovative

designs, and by reducing engine complexity. NASA Marshall Space Flight

Center’s 60K FASTRAC Engine is one example where these design philosophies have been put into practice. This engine burns a liquid kerosene/

oxygen mixture. It uses a one-piece, polymer composite thrust chamber/

nozzle that is constructed of a tape-wrapped silica phenolic liner, a metallic

injector interface ring, and a filament-wound epoxy overwrap (figure on

the left). This integral chamber/nozzle design minimizes engine operations

costs because it simplifies engine refurbishment procedures.

A cooperative effort between NASA Lewis Research Center’s Structures

Division and Marshall is underway to perform a finite element analysis of

the FASTRAC chamber/nozzle under all the loading and environmental

conditions that it will experience during its lifetime. The chamber/nozzle

is a complex composite structure.

Of its three different materials, the

two composite components have

distinctly different fiber architectures and, consequently, require

separate material model descriptions. Since the liner is tape

wrapped, it is orthotropic in the

nozzle global coordinates; and

since the overwrap is filament

wound, it is treated as a monoclinic

material. Furthermore, the wind

angle on the overwrap (figure on

the right) varies continuously along

the length of the chamber/nozzle.

Composite Nozzle/Thrust Chambers Analyzed for Low-Cost

Boosters

Thrust chamber/nozzle of 60K FASTRAC Engine. Thrust chamber/nozzle showing

filament wind angles.

c, wind 

 angle

Z'

u
132

Tangential thermal expansion coefficient versus axial position.

The angle is very shallow in the throat region and becomes steep toward

the ends.

During early fabrication attempts, cracking of the liner posed a significant

problem. The cracking was the result of residual stresses that developed

during processing because of the large differences between the thermal

expansion coefficients of the silica phenolic and the epoxy overwrap. The

final figure shows the tangential thermal expansion coefficient as a function of the axial position for both the liner and the overwrap. Although the

liner tangential thermal expansion coefficient is constant with position, the

overwrap thermal coefficient varies considerably along the length because

of the varying wind angle.

A finite element analysis of the chamber/nozzle was performed under

processing conditions. The results were instrumental in resolving the

residual-stress cracking problem and helped to establish a representative

analog using straight cylinders. The analyses verified that cylindrical analogs

0.0

0.5

1.0

1.5

2.0x10–5

0 10 20 30 40 50

Axial position, in.

Thermal expansion coefficient,

 in./in. °F

Epoxy overwrap

Silica phenolic liner

will duplicate the highest stress

states in the nozzle, which occur at

the nozzle throat. Furthermore,

they helped to identify large discrepancies between the material

strengths measured using the

traditional dogbone configuration

and the strengths measured in a

cylindrical configuration.

Future analyses of the chamber/

nozzle are planned to determine

the residual stress levels for a variety of possible material systems,

nozzle designs, and fiber architectures. The objective is to choose

the optimal material system to

minimize residual stresses. Furthermore, we plan to perform finite

element analyses for all the loading

and temperature conditions that

the chamber/nozzle will experience

during 60K FASTRAC Engine

operation.

Lewis contact:

Roy M. Sullivan, (216) 433–3249,

Roy.M.Sullivan@grc.nasa.gov

Author: Roy M. Sullivan

Headquarters program office: OAT

Programs/Projects: ASTP, LCBT

NASA’s Advanced Subsonic Technology (AST) Program seeks to develop

new technologies to increase the fuel efficiency of commercial aircraft

engines, improve the safety of engine operation, and reduce emissions and

engine noise. For new designs of ducted fans, compressors, and turbines

to achieve these goals, a basic aeroelastic requirement is that there should

be no flutter or high resonant blade stresses in the operating regime. For

verifying the aeroelastic soundness of the design, an accurate prediction/

analysis code is required. Such a three-dimensional viscous propulsion

aeroelastic code, named TURBO–AE, is being developed at the NASA Lewis

Research Center. The development and verification of the flutter version of

the TURBO–AE code (version 4) has been completed. Validation of the

code is partially complete.

The TURBO–AE aeroelastic code is

based on a three-dimensional

unsteady aerodynamic Euler/

Navier-Stokes turbomachinery code

TURBO, developed previously

under a grant from Lewis. This code

can model viscous flow effects,

which play an important role in

certain aeroelastic problems such

as flutter with flow separation,

flutter at high loading conditions

near the stall line (stall flutter), and

Flutter Version of Propulsion Aeroelasticity Code Completed
133

Research and Technology

Phase angle, deg

Unsteady moment

0 90 180 270 360

–2

–1

0

 1

 2 TURBO–AE v. 4, real

TURBO–AE v. 4, imaginary

Linear theory, real

Linear theory, imaginary

flutter in the presence of shock and boundary-layer interaction. In the

TURBO–AE code, the structural dynamics representation of the blade is

based on a normal mode representation. Any finite-element analysis code

can be used to calculate in-vacuum vibration modes and the associated

natural frequency. As an alternative, experimental measurements of these

vibration characteristics can be used.

A work-per-cycle approach is used to determine aeroelastic stability

(flutter). With this approach, the motion of the blade is prescribed to be a

harmonic vibration in a specified in-vacuum normal mode. The work done

by aerodynamic forces on the vibrating blade during a cycle of vibration

is calculated. If this work is positive, the blade is dynamically unstable,

because it will extract energy from the flow, leading to an increase in the

amplitude of the blade’s oscillation.

As part of the verification of the TURBO–AE code, calculations were performed for a helical fan test configuration. This configuration was used to

compare the code to two-dimensional linear potential (classical linear/flat

plate) theory. Results from the midspan of the blade undergoing torsional

vibrations were compared with linear theory results for the pitching motion

Unsteady moment for the torsional blade vibrations of a helical fan.

of a flat plate cascade. Excellent

agreement was observed for this

test case, providing a fundamental

verification of the TURBO–AE code.

TURBO–AE code will provide a

useful aeroelastic prediction/

analysis capability for engine

manufacturers. It will reduce design

cycle times by allowing new blade

designs to be verified for aeroelastic soundness before blades

are built and tested. Using this prediction capability, it will be possible

to build thinner, lighter, and faster

rotating blades without encountering aeroelastic problems like stall

flutter and high-cycle fatigue due

to forced vibrations.

Bibliography

Bakhle, M.A., et al.: Aeroelastic Calculations Based on Three-Dimensional

Euler Analysis. AIAA Paper 98–3295,

July 1998.

Lewis contacts:

Milind A. Bakhle, (216) 433–6037,

Milind.A.Bakhle@grc.nasa.gov;

Rakesh Srivastava, (216) 433–6045,

Rakesh.Srivastava@grc.nasa.gov; and

George L. Stefko, (216) 433–3920,

George.L.Stefko@grc.nasa.gov

Author: Milind A. Bakhle

Headquarters program office: OAT

Program/Project: AST, TURBO–AE

Applications of strongly gyroscopic rotors are becoming important, including flywheels for terrestrial and space energy storage and various attitude

control devices for spacecraft. Some of these applications, especially the

higher speed ones for energy storage, will have actively controlled magnetic bearings. These bearings will be required where speeds are too high

for conventional bearings, where adequate lubrication is undesirable or

impossible, or where bearing losses must be minimized for efficient energy

storage.

Flywheel rotors are highly gyroscopic, and above some speed that depends

on the bandwidth of the feedback system, they always become unstable in

an actively controlled magnetic bearing system. To assess ways to prevent

Stability of the Tilt Modes of an Actively Controlled

Flywheel Analyzed

instability until speeds well above

the desired operating range,

researchers at the NASA Lewis

Research Center used a commercial controls code to calculate the

eigenvalues of the tilt modes of a

rigid gyroscopic rotor supported by

active magnetic bearings. The real

part of the eigenvalue is the negative of the damping of the mode,

and the imaginary part is approximately equal to the mode’s

frequency.
134

A modal controller was presumed in which the pure translation and pure

tilt modes were separated. We addressed only the tilt modes. The controller included simple proportional-derivative (PD) gains for each tilt angle

and proportional and derivative cross-coupling gains. Bandwidths were

imposed in the control loop to represent the various phase lags of a

magnetic bearing system. The effects of bandwidths, cross-axis proportional gain, and cross-axis derivative gain (gyroscopic canceling) were

considered individually and in combination to show the qualitative and

quantitative effects of each. Both cross-axis gains were “scheduled” by

being multiplied by the rotor spin speed Ω since they were needed only to

counteract the effects of gyroscopic torques.

The tilt mode of a nonspinning rotor becomes two modes under rotation:

one mode (called the forward whirl mode) goes up in frequency with

the rotor speed; the other (called the backward whirl mode) goes down.

–40 –30 –20

Real part of eigenvalue

–10 0

Frequency, Hz

100

101

102

103 crstc

0

Rotor spin

speed,

rpm

Rotor spin

speed,

rpm

4 8 12 16 20 30x103

20

10

5

0

crstc

20 16 12 8 4 0

30

20

10

5x103

Frequency, Hz 

–160 –140 –120 –100 –80

Real part of eigenvalue

–60 –40 –20 0 20

0

200

400

600

800

1000

1200

1400

1600

Rotor

spin

speed,

rpm

Gyro 

cancelling, 

percent

60x103

50

40

30

10 5 0 15

25 20

20

10

0

We found that cross-axis proportional gain, crstc, increases the

damping of the backward whirl

mode, which is otherwise a poorly

damped, low-frequency mode at

high rotor spin speed (see the top

figure, which was calculated with

high bandwidth). The point common to all curves is for zero rpm.

Damping increases toward the left.

This gain (crstc) decreases the

damping of the forward whirl

mode, which is otherwise better

damped, but which could be

subject to strong forcing by rotor

unbalance as rotor speed increases

(see the top figure). Cross-axis

derivative gain (also called gyroscopic cancelling) can improve the

forward whirl mode damping but

has little effect on the backward

whirl mode. An appropriate combination of these cross-axis gains can

result in a closed-loop system that

is stable over a wide speed range

without additional gain scheduling,

as shown in the bottom figure,

which was calculated with more

realistic bandwidths. Again, the

point common to all curves is for

zero rpm. Modes to the right of the

vertical axis are unstable; modes

to the left are progressively more

stable as damping increases. Note

that with 10-percent gyroscopic

canceling, for example, the flywheel would be stable to nearly

60 000 rpm.

Lewis contacts:

Dr. Gerald V. Brown, (216) 433–6047,

Gerald.V.Brown@grc.nasa.gov; and

Albert F. Kascak, (216) 433–6024,

Albert.F.Kascak@grc.nasa.gov

Authors:

Dr. Gerald V. Brown and

Albert F. Kascak

Headquarters program office: OAT

Program/Project: Magnetic bearings

for flywheels for energy storage and

space attitude control, ACESE

Effect of position cross coupling with Ω factor; five corners at 50 kHz.

Effect of small-percent gyro canceling with position cross coupling; five

corners at 2 kHz; crstc, 12.
135

Research and Technology

A dual-probe, integrated fiber-optic laser system for measuring blade tip

clearance in rotating turbomachinery was developed cooperatively with

Integrated Fiber Optic Systems, Inc., and tested at the NASA Lewis Research

Center. Because the probes are nearly flush with the casing inner lining,

there is minimal flow disturbance. The two probes are closely spaced in a

circumferential plane and are slanted at an angle relative to each other so

that the time it takes the blade tip to traverse the space between the two

beams varies with the tip radius, making it possible to determine the tip

clearance at rotor operating conditions. The tip clearance can be obtained

for all the blades in a rotor with a single system, provided there are no

synchronous vibrations present at a particular operating condition.

The two probes for the laser system were installed in two holders. For one,

the included angle between the probes was 20°, and for the other, it was

40°. The two configurations were calibrated in a vacuum spin-rig facility

that can reproduce realistic blade-tip speeds. A specially designed, nondeflecting rotor was used to calibrate the probes. This rotor consisted of a

tapered titanium bar with three teeth on each end, the thickness of which

was representative of a typical blade-tip thickness. The bar and measuring

teeth were in a horizontal plane in the spin rig, and the axis of rotation was

vertical. To set the clearances, we used a remotely controlled optical stage

that could be traversed vertically.

The 40° probe system also was

used to measure the tip clearance

during an engine fan prototype test

in a wind tunnel. Using the spinrig calibrations, we estimated the

accuracy of the tip-clearance measurement in the test to be 0.005 in.

This program is ongoing.

Bibliography

Dhadwal, H.S.; and Kurkov, A.P.: DualLaser Probe Measurements of Blade-Tip

Clearance. ASME Paper 98–GT–183,

1998.

Lewis contact:

Dr. Anatole P. Kurkov, (216) 433–5695,

Anatole.P.Kurkov@grc.nasa.gov

Authors:

Dr. Harbans S. Dhadwal and

Dr. Anatole P. Kurkov

Headquarters program office: OAT

Program/Project: AST

Dual-Laser Probe for Measuring Blade-Tip Clearance Tested

Under the Advanced Subsonic Technology (AST) Program, an aeroelastic

analysis code (TURBO–AE) based on Navier-Stokes equations is currently

under development at NASA Lewis Research Center’s Machine Dynamics

Branch. For a blade row, aeroelastic instability can occur in any of the possible interblade phase angles (IBPA’s). Analyzing small IBPA’s is very computationally expensive because a large number of blade passages must be

simulated. To reduce the computational cost of these analyses, we used

time-shifted, or phase-lagged, boundary conditions in the TURBO–AE code.

These conditions can be used to

reduce the computational domain

to a single blade passage by

requiring the boundary conditions

across the passage to be lagged

depending on the IBPA being

analyzed. The time-shifted

boundary conditions currently

Time-Shifted Boundary Conditions Used for Navier-Stokes

Aeroelastic Solver

x/c

Unsteady pressure difference

Real

0.00 0.25 0.50 0.75 1.00

–10

10

–5

0

5

Imaginary

Phase-lagged

Two-passage

Analysis

x/c

Unsteady pressure difference

Real

0.00 0.25 0.50 0.75 1.00

–10

10

–5

0

5

Imaginary

Phase-lagged

Two-passage

Analysis

Comparison of the unsteady pressure difference variation at midspan for 180o IBPA oscillations; distance along chord from

leading edge, x; chord, c. Left: Pitching. Right: Plunging.
136

Vibration cycle

Work per cycle

0 2 4 6 8

–1.8

–1.6

–1.4

–1.2

–1.0

–0.8x10–4

Blade 1

Blade 2

Blade 3

Blade 4

Time-shifted

Analysis

Vibration cycle

Work per cycle

0 2 4 6 8

–2.0

–1.5

–1.0

–0.5

–0.0x10–4

45

45

90

90

180

IBPA,

deg

implemented are based on the direct-store method. This method requires

large amounts of data to be stored over a period of the oscillation cycle. On

CRAY computers this is not a major problem because solid-state devices

can be used for fast input and output to read and write the data onto a

disk instead of storing it in core memory.

In aeroelastic analyses using TURBO–AE, the unsteady aerodynamic loads

are obtained by solving the Reynolds-averaged Navier-Stokes equations.

The aerodynamic equations are solved via a finite volume scheme. Flux

vector splitting is used to evaluate the flux Jacobians on the left side, and

a higher order Total Variation Diminishing (TVD) scheme based on Roe’s

flux vector splitting is used to discretize the fluxes on the right side. Newton subiterations are used at each time step to maintain higher accuracy.

Then, symmetric Gauss-Sidel iterations are applied to the discretized equations, and a dynamic grid deformation technique is used to simulate the

blade motions. TURBO–AE updates the grid at each time step by recalculating it using linear interpolation and by assuming the far field to be fixed.

It determines the aeroelastic characteristics of the blade row by calculating

the energy exchange between the vibrating blade and its surrounding fluid

for all possible frequencies and IBPA’s of interest. Positive work on the

blade indicates instability.

To verify and validate the time-shifted analysis, we used a flat-plate, helicalfan geometry. The helical fan consisted of 24 flat-plate blades with zero

thickness. The inflow conditions and stagger angle were set to induce an

inflow with a relative Mach number of 0.7 at an incidence of zero at the

midspan. The results were obtained for several IBPA’s and compared with

those of the analysis using multiple blade passages. Only a few typical

results are shown here; for more detail please refer to the reference. The

graphs on the preceeding page show the unsteady pressure difference for

the 180o IBPA pitching and plunging cases. The time-shifted results were

compared with results from a two-passage analysis. A very good comparison

was obtained in both cases, indicating that the time-shifted boundary condition was properly implemented and reproduced the results of the multiplepassage analysis.

In the graph on the left, the rate of convergence for the time-shifted

analysis for –90° IBPA pitching is compared with that of the multiplepassage analysis. From this figure,

it can be seen that the work-percycle for the four-passage analysis

converged within four to five oscillation cycles, whereas roughly

seven to eight cycles were required

for the time-shifted analysis. Even

though approximately 40-percent

more cycles were required for

convergence, the computational

time was reduced by almost

60 percent. This is because only

one blade passage was needed in

the time-shifted analysis. We also

found that the convergence rate

for the time-shifted analysis was

independent of the IBPA analyzed

(see the graph on the right).

Bibliography

Srivastava, R., et al.: Application of

Time-Shifted Boundary Conditions to

a 3D Euler/Navier-Stokes Aeroelastic

Code. ASME Paper 98–GT–42, 1998.

Lewis contacts:

Rakesh Srivastava, (216) 433–6045,

Rakesh.Srivastava@grc.nasa.gov;

Milind A. Bakhle, (216) 433–6037,

Milind.A.Bakhle@grc.nasa.gov; and

George L. Stefko (216) 433–3920,

George.L.Stefko@grc.nasa.gov

Author: Rakesh Srivastava

Headquarters program office: OAT

Program/Project: AST, TURBO–AE

Comparison of the work-per-cycle convergence for a timeshifted analysis with multiple passages.Comparison of the work-per-cycle convergence history

for various interblade-phase-angle pitching oscillations.
137

Research and Technology

Case 1 Baseline

Case 3 ~30° Sweep

No lean

Case 2 tip spacing

Case 4 ~30° Sweep

~30° Lean (with fan rotation)

Case 2 with lean

Vane tip

Leaned van

Hub

Shroud

Direction

of rotor

rotation

Axial view looking upstream

Xtip

Case 2 No sweep or lean

Increased axial spacing

Xtip

Straight 

vane

C-98-3638 C-98-3638

Low-Noise Potential of Advanced Fan Stage Stator Vane

Designs Verified in NASA Lewis Wind Tunnel Test

NASA/Allison Engine Company Low Noise Fan test in Lewis’ 9- by 15-Foot Low-Speed

Wind Tunnel.

NASA/Allison Low Noise Fan test; stator vane sweep and lean.

With the advent of new, more

stringent noise regulations in the

next century, aircraft engine manufacturers are investigating new

technologies to make the current

generation of aircraft engines as

well as the next generation of

advanced engines quieter without

sacrificing operating performance.

A current NASA initiative called the

Advanced Subsonic Technology

(AST) Program has set as a goal

a 6–EPNdB (effective perceived

noise) reduction in aircraft engine

noise relative to 1992 technology

levels by the year 2000. As part of

this noise program, and in cooperation with the Allison Engine

Company, an advanced, low-noise,

high-bypass-ratio fan stage design

and several advanced technology

stator vane designs were recently

tested in NASA Lewis Research

Center’s 9- by 15-Foot Low-Speed

Wind Tunnel (an anechoic facility).

The project was called the NASA/

Allison Low Noise Fan (see the

figures).

A bypass fan stage consists of a fan

and a stator vane assembly. The

advanced technology bypass fan

stage for this project was 22 in. in

diameter and consisted of a lowpressure ratio, 18-bladed fan and

an integral, 42-bladed stator vane

assembly. The testing used an air

turbine propulsion simulator to

power the model fan stage across a

wide range of operating conditions

from idle to full takeoff power at a

wind tunnel velocity of Mach 0.10,

or about 67 knots. The main objective of the project was to evaluate

the noise reduction potential of

a series of advanced stator vane

designs. Variations in stator vane

geometry that were introduced

include increasing the spacing

between the fan and the stator

vanes, incorporating sweep into
138

the stator vanes (similar to aircraft wing sweep), and finally leaning the

swept stator vanes over in the fan direction of rotation. During testing, the

acoustic characteristics of each fan and stator vane assembly combination

were measured using fixed and traversing microphones inside the wind

tunnel. The aerodynamic performance of each fan/stator vane combination

was also measured using pressure/temperature rakes and force balances

to determine what effect the new low-noise stator vanes would have on

the fan stage efficiency.

The research results generated as part of this project allowed a major

interim milestone, to demonstrate a 3–EPNdB reduction in noise by 1997,

to be reached for the Advanced Subsonic Technology Program. Test results

will be used to create a data base that engineers can draw upon for comparison with computer predictions of the effect of fan stage geometry on

aircraft engine noise. In addition, aircraft manufacturers will be able to use

these results to influence the design of the next generation of advanced

technology aircraft engines or the next growth version of current technology engines.

Lewis contacts: Christopher E. Hughes,

(216) 433–3924,

Christopher.E.Hughes@grc.nasa.gov;

and Richard P. Woodward,

(216) 433–3923,

Richard.P.Woodward@grc.nasa.gov

Author: Christopher E. Hughes

Headquarters program office: OAT

Programs/Projects:

AST (Noise Reduction Element)

The need to improve space mechanism reliability is underscored by a long

history of flight failures and anomalies caused by malfunctioning mechanisms on spacecraft and launch vehicles. Some examples of these failures

are listed in the table. Mechanism anomalies continue to occur and to be a

cause of catastrophic mission failures. Several factors cause problems for

space system mechanisms. The space environment produces wide temperature ranges, thermal gradients, and rapid changes in temperature,

which can bind the moving parts of mechanisms. Ultraviolet radiation and

vacuum cause the properties of many materials to degrade to unacceptable levels or to behave differently in space than on Earth, making it

difficult to simulate operation during ground tests. The lack of gravity in

space causes mechanisms to operate differently than on the ground.

Sometimes the effects of zero gravity can be simulated to some degree in

ground testing, such as by offloading the weight of a deployable appendage. Other effects, such as lubricant migration, cannot be simulated and

must be considered in the design. Finally, the launch environment imposes

severe dynamic loads on mechanisms and can cause structural damage,

loosen fasteners, and damage delicate surfaces.

Given these complexities, it is not surprising that it is not always possible

to uncover and correct all the hidden problems with mechanisms prior

to launch. Fortunately, there are ways to reduce the number of failures

involving mechanisms and/or mitigate the effects of a failure of a component. In many cases, failures were caused by design problems that have

caused similar failures in the past, and thus could have been avoided had

the designers been aware of the past mistakes. Because much experience

has been gained over the years, many specialized design practices have

evolved and many unsatisfactory design approaches have been identified.

In many cases, however, this knowledge has remained with the individual

NASA Space Mechanisms Handbook—Lessons Learned

Documented

mechanism designer and has not

been widely shared.

To alleviate this situation, NASA and

the NASA Lewis Research Center

conducted a Lessons Learned

Study (refs. 1 and 2) and wrote a

handbook to document what has

been learned in the past. The

primary goals of the handbook

were to identify desirable and

undesirable design practices for

space mechanisms and to reduce

the number of failures caused by

the repetition of past design errors.

Another goal was to identify a

variety of design approaches for

specific applications and to provide

the associated considerations and

caveats for each approach in an

effort to help designers choose the

approach most suitable for each

application. The handbook also

provides some design principles.

These principles, which can be

applied to any mechanism to avoid

common failure modes, can be

particularly useful for the esoteric

mechanism configurations that
139

Research and Technology

dwell on topics that are not unique to space applications, it does cite

references, where appropriate, for additional information or more indepth

discussion of specific topics.

The handbook is divided into six parts. Part I, Introduction to Space Mechanisms, starts with an overview of various types of spacecraft mechanisms. It

then discusses the requirements that are typically imposed on space mechanisms, their implications, and what steps can be taken to ensure that the

requirements are met. The discussion concludes with a description of a

typical mechanism design process and addresses how the design evolves

from concept to fabrication. Part II, Design Considerations for Space Mechanisms, provides guidelines for recommended design practices for most

spacecraft mechanisms. It also contains subsequent chapters that are

devoted to guidelines applicable to specific types of mechanisms. Part III,

Space Mechanism Components, proceeds to the next level of detail and

discusses design considerations for

mechanisms. This part is divided

into general design guidelines that

are applicable to the various components of spacecraft mechanisms.

Part IV delves into two areas of

testing, environmental testing and

tribological testing of space mechanisms. Part V lists expert areas

and the names and addresses of

individuals who are experts in

those areas of testing. Finally,

Part VI lists testing laboratories

and the individuals involved in the

testing programs.

SUMMARY OF SPACECRAFT MECHANISM FAILURES

Program Date Problem Cause

Program 461 1964 Solar array failed to deploy fully Mishandling during stowage

STP 67–2 (OV2–5) 1968 Solar array booms failed to deploy fully Field modification problem

777 1970 Omni antenna latch broke during spin-up Attitude control instability

Program A 1971 Antenna failed to deploy fully Wire harness binding

Program B 1971 Solar array deployed late Silicon rubber sticking

STP 71–5 1972 Boom failed to deploy Dynamic clearance problem

Skylab 1973 Solar array failed to deploy Interference with cabling or thermal blankets

Transit 1975 Solar array failed to deploy fully; cable hung up Anomalous flat trajectory caused high heating rates

Viking 1975 Sampling arm failed to deploy Debris in gear train

STP 74–1 (Solrad) 1976 Solar panel failed to deploy Release mechanism binding

DMSP F–1 1976 Solar array failed to deploy fully Excessive wire harness stiffness

DMSP F–2 1977 Solar array delayed release Friction welding

Science boom failed to deploy fully Microswitch failed

Voyager 1977 Science boom failed to deploy fully Microswitch failed

Scan platform gearbox seized Lubricant failed

Magnetometer boom misaligned Unknown

Seasat 1978 Spacecraft power failed Slip ring debris between power and ground rings

Apple 1981 Solar array failed to deploy Failure of deployment device

DE 1981 Sensing antenna failed to deploy Unknown

Insat 1 1982 Solar sail failed to deploy Unknown

ERBS 1982 Solar array failed to deploy Unknown

GLOMR 1985 Spacecraft failed to separate from orbiter Canister door did not open fully

VUE 1988 Telescope failed to rotate about azimuth Inadequate torque margin on azimuth caging arm

Galileo 1989 High-gain antenna failed to deploy Cold welding in ball and socket joint

Instrument cover jettisoned late Thermal binding

Magellan 1989 Solar array failed to latch at end of travel Microswitch misadjusted

Macsat 1990 Gravity gradient boom failed to deploy Inadequate force margin

CRRES 1990 Magnetometer boom failed to orient fully Interference between thermal blanket Velcro and

wiring harness

Ulysses 1990 Spin-stabilized spacecraft wobbled Antenna boom thermal distortion caused spacecraft

center-of-gravity offset

Hubble 1990 Solar array booms jittered as telescope went

between sun and shade

Thermal gradient across boom diameter

ANIK E2 1991 C-band antenna failed to fully deploy Thermal blanket interference

Unknown Sampling arm failed to deploy Screw backed out and wedged against housing

Tether Satellite System 1993 Reel-out mechanism jammed Screw added for structural margin interfered with

reel-out mechanism

GOES 10 1997 Solar Array Drive malfunctioned Under investigation
140

We anticipate that this handbook will be useful to a variety of readers. By

studying the numerous guidelines presented in this handbook, entry-level

design engineers will be able to quickly gather practical information on

how to avoid common pitfalls. Experienced mechanical design engineers

who are new to space mechanism applications will benefit from learning

the unique requirements created by the space and launch environments.

Also, users who need to evaluate their suppliers’ products, but have little

personal experience in the design of mechanisms, can find useful information on identifying key performance, risk, and cost drivers for most

space mechanisms and components. The Space Mechanisms Handbook

is available from Lewis’ Mechanical Components Branch.

References

1. Shapiro, W., et al.: Space Mechanisms Lessons Learned Study, Volume I—

Summary. NASA TM–107046, 1995.

2. Shapiro, W., et al.: Space Mechanisms Lessons Learned Study,

Volume II—Literature Review. NASA

TM–107047, 1995.

Lewis contact:

Robert L. Fusaro, (216) 433–6080,

Robert.L.Fusaro@grc.nasa.gov

Author: Robert L. Fusaro

Headquarters program office:

OA (Chief Engineer)

Programs/Projects:

Safety and Mission Assurance

Gears, bearings, and similar mechanical elements transmit loads through

the surfaces that are in contact with one another. Thus, the fatigue lives of

gears can be improved by providing smoother tooth surfaces. At the NASA

Lewis Research Center, we completed a metrology evaluation of one

method for making gears with a highly polished, mirrorlike surface (see the

following photo). The polished gears were measured carefully. The measurement data showed that the polishing process did, indeed, reduce the

surface roughness but did not change the overall tooth shape in any

harmful way.

This work was done as a

partnership of NASA, the

U.S. Army, and the University of Wales. NASA provided conventionally

ground gear specimens and

has begun testing to determine the fatigue lives of the

superfinished gears. Under

contract, the University of

Wales superfinished the

gears and inspected them

before and after the superfinishing operation. The U.S.

Army European Research

Office provided the funds

and procured the contract

with the University of Wales.

For gears, the rate of fatigue is

greatly affected by the ratio of the

oil film thickness to the roughness

of the contacting surfaces. In this

work we are seeking to improve

fatigue lives by reducing the surface roughness. Conventionally

ground, aerospace quality gears

were manufactured, and their

geometry was inspected. Next,

the gears were superfinished by

placing them in a vibrating bath

consisting of water, detergent,

abrasive powder, and small pieces

of zinc. Upon removal from the

bath, the surfaces were highly

polished (see the preceding

photo). The gears were then again

inspected, and the measurements

of the gears before and after the

superfinishing operation were

compared. Typical inspection data

are provided in the plots on the

following page. The grinding marks

are clearly evident in the figure on

the left. Superfinishing removed

the peaks of the grinding marks

and left a much smoother surface

Metrology Evaluation of Superfinished Gears Completed

Highly polished surface of superfinished gears.
141

Research and Technology

100 μm

1 μm

100 μm

100 μm

1 μm

100 μm

Profile of tooth surfaces. Left: Ground. Right: Superfinished.

as is evident in the figure on the right. Profile and spacing checks proved

that the overall gear tooth shape was not affected in any harmful way.

Superfinishing uniformly removed approximately 2.5 μm from each surface.

See reference 1 for a complete report.

Reference

1. Snidle, R.W.; Evans, H.P.; and Alanou, M.P.: The Effect of Superfinishing on Gear

Tooth Profile. Report AD–A327916, June 1997. (Available from the Defense

Technical Information Center (DTIC) (http://www.dtic.mil/) or the National

Technical Information Service (NTIS) (http://www.ntis.gov/), or the Center for

AeroSpace Information (http://www.sti.nasa.gov/rselect/FAQ.html#docorder).

Lewis contact:

Timothy L. Krantz, (216) 433–3580,

Timothy.L.Krantz@grc.nasa.gov

Author: Timothy L. Krantz

Headquarters program office: OAT

Programs/Projects:

Rotorcraft Base Program

Gears used in current helicopters and turboprops are designed for light

weight, high margins of safety, and high reliability. However, unexpected

gear failures may occur even with adequate tooth design. To design an

extremely safe system, the designer must ask and address the question,

“What happens when a failure occurs?” With gear-tooth bending fatigue,

tooth or rim fractures may occur. A crack that propagates through a rim

will be catastrophic, leading to disengagement of the rotor or propeller,

loss of an aircraft, and possible fatalities. This failure mode should be

avoided. A crack that propagates through a tooth may or may not be

catastrophic, depending on the design and operating conditions. Also,

early warning of this failure mode may be possible because of advances

in modern diagnostic systems.

One concept proposed to address bending fatigue fracture from a safety

aspect is a split-tooth gear design (ref. 1). The prime objective of this

design would be to control crack propagation in a desired direction such

that at least half of the tooth would remain operational should a bending

failure occur. A study at the NASA Lewis Research Center analytically validated the crack-propagation failsafe characteristics of a split-tooth gear

(ref. 2). It used a specially developed three-dimensional crack analysis

program that was based on boundary element modeling and principles

of linear elastic fracture mechanics.

Crack shapes as well as the crackpropagation life were predicted on

the basis of the calculated stress

intensity factors, mixed-mode

crack-propagation trajectory theories, and fatigue crack-growth

theories.

The following figures show the

effect of the location of initial

cracks on crack propagation. Initial

cracks in the fillet of the teeth produced stress intensity factors of

greater magnitude (and thus,

greater crack growth rates) than

those in the root or groove areas

of the teeth. Crack growth was

simulated in a case study to evaluate crack-propagation paths (see

the final figure). Tooth fracture was

Three-Dimensional Gear Crack Propagation Studied
142

Effect of initial crack location on mode I stress intensity factors. Left: Tooth fillet locations. Center: Tooth foot locations.

Right: Tooth groove locations.

Defense and Space Group. The

crack-propagation package was

developed by the Cornell Fracture

Group at Cornell University. The

reported results, which are the

initial findings of Lewis’ gear-crackpropagation research for the

Rotorcraft Base Program, will be

further investigated to develop

generalized gear design guidelines.

References

1. Drago, R.J.; Sane, A.D.; and Brown,

F.W.: UltraSafe Gear Systems for

Critical Applications—Initial Development. AGMA TP–97FTM10, 1997.

2. Lewicki, D.G., et al.: ThreeDimensional Gear Crack Propagation

Studies. NASA TM—1999-208827,

1999.

Lewis contact:

Dr. David G. Lewicki, (216) 433–3970,

David.G.Lewicki@grc.nasa.gov

Author: Dr. David G. Lewicki

Headquarters program office: OAT

Programs/Projects:

Rotorcraft Base Program

Crack

surface

Exploded gear tooth view of predicted crack growth after 15 steps.

0.0 0.2 0.4 0.6 0.8 1.0

15

20

25

30

+x

Normalized position along crack front

Mode I stress intensity factor, MPa √m

Initial crack 5

Initial crack 6

0.0 0.2 0.4 0.6 0.8 1.0

0

5

10

15

+z

Normalized position along crack front

Mode I stress intensity factor, MPa √m

Initial crack 7

Initial crack 8

Initial crack 9

Normalized position along crack front

0.0 0.2 0.4 0.6 0.8 1.0

Mode I stress intensity factor, MPa √m

20

25

30

35

40 +x

Left

tooth

Right

tooth

Initial crack 4

Initial crack 4

Initial crack 3

Initial crack 2

Initial crack 1

Rear

Front

predicted from the crack-growth simulation for an initial crack in the tooth

fillet region. This was the desired failure mode for an ultrasafe design.

Lastly, tooth loads on the uncracked mesh of the split-tooth design were

up to five times greater than those on the cracked mesh if equal deflections of the cracked and uncracked teeth were considered. This effect

needs to be considered in the design of a split-tooth configuration.

This work was done in-house at Lewis in support of the National Rotorcraft Technology Center project, Ultra-Safe Gear Design, with the Boeing
143

Research and Technology

CD-98-77670

Throat 

assembly Bearing 

assembly

Forward exit

 cone assembly

Nozzle inlet

assembly Cowl housing

assembly

Aft exit cone

assembly

Exhaust 

flow

Vent port

Secondary

O-ring

Leak

check port

1

5

4

2

3

Rocket centerline

A

A

Section A - A

Thermal barrier

Primary

O-ring

Nozzle-tocase joint

Assembly joints of modern solid-rocket motor cases are usually sealed with

conventional O-ring seals. The 5500 °F combustion gases produced by

rocket motors are kept a safe distance away from the seals by thick layers

of insulation and by special compounds that fill assembly splitlines in the

insulation. On limited occasions, NASA has observed charring of the primary O-rings of the space shuttle solid-rocket nozzle-assembly joints

due to parasitic leakage paths opening up in the gap-fill compounds

during rocket operation. Thus, solid-rocket motor manufacturer Thiokol

approached the NASA Lewis Research Center about the possibility of applying Lewis’ braided-fiber preform seal as a thermal barrier to protect the

O-ring seals. This thermal barrier would be placed upstream of the primary

O-rings in the nozzle-to-case joints to prevent hot gases from impinging on

the O-ring seals (see the following illustration). The illustration also shows

joints 1 through 5, which are potential sites where the thermal barrier

could be used.

Burn tests at temperatures representative of the rocket thermal

environment were used to evaluate

the thermal resistance of braided

rope thermal barriers made of

different materials. Thermal barriers

were placed in the hottest part of

the flame of an oxyacetylene torch

at 5500 °F, and the amount of

time to completely cut through

them was measured. A 0.125-in.-

diameter stainless steel rod was

cut through in 5 sec, whereas

0.125-in.-diameter thermal barriers

of braided ceramic and superalloy

materials lasted less than 15 sec.

In contrast, 0.125-in.-diameter thermal barriers of braided carbon

fibers lasted over 2 min. Carbonfiber thermal barriers with diameters of 0.200 and 0.260 in. lasted

over 6.5 and 8.5 min., respectively,

before they were cut through (see

the photo on the following page).

As a point of reference, the solidrocket motors of the space shuttle

only burn for 2 and 4 sec, much

shorter than the burn-through

times of the carbon-fiber thermal

barriers. On the basis of these

results, the decision was made

to use carbon fibers to braid the

thermal barriers.

Flow tests performed on the thermal barriers showed that they

blocked hot gas flow but were

permeable enough to allow leak

checks of the primary/secondary

O-ring system. The barriers also

were resilient enough to accommodate flange movements in the

solid-rocket motor. To simulate a

rocket environment, Thiokol

performed subscale rocket “char”

motor tests in which the thermal

barrier was subjected to hot gases

that flowed through an intentional

0.060-in. circumferential gap defect

both upstream and downstream of

NASA Lewis Thermal Barrier Feasibility Investigated for Use

in Space Shuttle Solid-Rocket Motor Nozzle-to-Case Joints

Potential shuttle solid-rocket motor joint locations (circled numbers) for candidate

thermal barrier. Top: Enlarged view of nozzle-to-case joint showing primary and

secondary pressure O-rings, leak-check part, and proposed thermal barrier

location. Bottom: Overall nozzle cross section (half view).
144

Candidate thermal barrier (0.2-in. diameter) for shuttle solid-rocket motor in 5500 °F

oxyacetylene burn test. Time for burn through, 6.5 min.

Preliminary subscale (70-lbf thrust) “char” motor tests examining thermal barrier effectiveness. Left: Test

configuration: thermal barrier (0.125-in. diameter) filling an intentional joint defect. Right: Temperature data:

upstream (Thot ) and downstream (Tcold ) sides of thermal barrier. (Copyright Thiokol Corp.; used with permission.)

CD-98-77561

the thermal barrier. During the 11-sec rocket firing, temperatures over

3200 °F were measured on the hot side of the 0.260-in.-diameter thermal

barrier, whereas temperatures on the cold side were about 950 °F, for a

temperature drop of over 2200 °F across the thermal barrier (see the

graphs).

On the basis of these results, additional mechanical and thermal testing is

planned at Lewis to help further characterize the thermal barrier. Thiokol

is planning to test the thermal barrier in a subscale solid-rocket motor,

where it would be installed first in its undamaged state and then with an

intentional defect (Spring and Fall

1999). If all planned tests show

success, the Lewis-developed

thermal barrier would be prepared

for Full Scale RSRM static tests in

November 2000 (no-joint defect)

and May 2002 (with joint defect);

it would be subsequently qualified

for flight.

11-sec

burn time Time, sec

"Char" motor nozzle

test article

Rocket flow

(5000 °F

estimate)

Tcold

Thot

Tcold

Thot

Thermal barrier seal

3000

4000

2000

1000

0

Temperature, °F

Intentional

defect

0 10 20 30
145

Research and Technology

Find out more on the World Wide Web:

http://www.grc.nasa.gov/WWW/TU/InventYr/1996Inv_Yr.htm

Bibliography

Steinetz, B.M.; and Dunlap, P.H.: Feasibility Assessment of Thermal Barrier Seals for

Extreme Transient Temperatures. AIAA Paper 98-3288 (NASA TM—1998-208484)

1998.

Lewis contact:

Dr. Bruce M. Steinetz, (216) 433–3302, Bruce.M.Steinetz@grc.nasa.gov

Modern Technologies Corp. contact:

Patrick H. Dunlap, Jr., (216) 433–6374, Patrick.H.Dunlap@grc.nasa.gov

Authors:

Dr. Bruce M. Steinetz and

Patrick H. Dunlap, Jr.

Headquarters program office: OAT

Programs/Projects: HITEMP,

High Temperature Seal Development

Special recognition: 1996 NASA

Invention of the Year awarded to the

fiber preform seal, precursor to the

thermal barrier seal
147

Space
148

The Advanced Communications Technology Satellite (ACTS) system has

been modified to support operation in an inclined orbit that is virtually

transparent to users, and plans are to continue this final phase of its

operation through September 2000. The next 2 years of ACTS will provide

a new opportunity for using the technologies that this system brought

online over 5 years ago and that are still being used to resolve the technical issues that face NASA and the satellite industry in the area of seamless

networking and interoperability with terrestrial systems.

New goals for ACTS have been defined that align the program with recent

changes in NASA and industry. ACTS will be used as a testbed to

(1) Show how NASA and other Government agencies can use commercial

systems for future support of their operations

(2) Test, characterize, and resolve technical issues in using advanced

communications protocols such as asynchronous transfer mode (ATM) and

transmission control protocol/Internet protocol (TCP/IP) over long latency

links as found when interoperating satellites with terrestrial systems

(3) Evaluate narrow-spot-beam Ka-band satellite operation in an inclined

orbit

(4) Verify Ka-band satellite technologies since no other Ka-band system is

yet available in the United States

On July 9, 1998, the final north/south stationkeeping was performed on

ACTS. Inclined-orbit operations technically began when the spacecraft left

its original north/south limits of +0.05° on August 8, 1998. The spacecraft

will be maintained at 100° W longitude throughout the remaining life of

the program. At the program’s completion, a final sequence of burns will

superorbit the satellite into a higher orbit where the payload and spacecraft will be shut down.

Preparation for inclined-orbit

operations involved reviewing all

aspects of the communications

link that could be affected by the

spacecraft no longer being stationary. Changes were made to

onboard software and operations

procedures to maintain the pointing accuracy of the narrow 0.3°

spot beams so that the coverage

pattern would not change. With

the spacecraft increasing in north/

south drift by 0.78° per year,

inclined-orbit operations will be

limited to about 2.5 years, or

30 months.

The ground segment was modified in-house at the NASA Lewis

Research Center to track the satellite. The four major components

modified with tracking hardware

were the T1VSAT’s,a Gigabit Earth

Stations, USAT’s,b and the Link

Evaluation Terminal. The Master

Ground Station at Lewis was originally designed with a tracking

system for its 5.5-m antenna, so no

further modifications were needed.

Inclined-orbit operations offer the

ACTS program a final opportunity

to affect the design of the next

generation of communications

satellites. To encourage new experiments and announce the ongoing

opportunity to use ACTS, Lewis

Advanced Communications Technology Satellite Now

Operating in an Inclined Orbit

Tracking modifications to a 1.2-m T1VSAT.

Space Communications

_______________

aT1 Very Small Aperture Terminal—an Earth

station that operates at the telephony

standard data rate of T1 or 1.544 Mbps

and uses a 1.2-m-diameter reflector.

bUltra Small Aperture Terminal—a small

Earth station that uses a 0.6-m-diameter

reflector and supports data rates of up to

T1 for transmissions while receiving at up

to 40 Mbps.
149

Space

released an experiment opportunity guide (http://kronos.grc.nasa.gov/

acts/eoa/guide.html) in March 1998.

A set of in-house technical experiments is planned as the core set of

experiments for operations. Other experiments hosted by Lewis’ Space

Communications Program include “High Data Rate Interoperability,” “Internet Protocol Performance and Coding Effects,” “Hybrid Networking Using

Digital Imaging and Communications in Medicine (DICOM)/Telemammography,” and “Advanced Air Transportation Technology.” Experiments are

currently planned by the Air Force Research Laboratory (Rome Labs),

Carnegie Mellon University, Lockheed Martin Western Development Laboratory, Savannah State University, and Intelsat. Over 20 organizations have

expressed interest in using ACTS during inclined-orbit operations, including

the Boeing Company, Southwest Research Institute, Texas Medical Center,

and the Naval Research Laboratory. The next 2 years hold promise for new

experiment plans as ACTS continues to be an important resource

for resolving issues in satellite

communications.

Find out more about this research on

the World Wide Web:

http://acts.grc.nasa.gov

Lewis contact:

Robert A. Bauer, (216) 433–3431,

Robert.A.Bauer@grc.nasa.gov

Author: Robert A. Bauer

Headquarters program office: OSF

Programs/Projects:

ACTS, SOMO, NII/GII, CSOC

The recently completed DARTFire sounding rocket microgravity combustion

experiment launched a new era in the imaging of flames in microgravity.

DARTFire stands for “Diffusive and Radiative Transport in Fires,” which

perfectly describes the two primary variables—diffusive flow and radiation

effects—that were studied in the four launches of this program (June 1996

to September 1997). During each

launch, two experiments, which

were conducted simultaneously

during the 6 min of microgravity,

obtained results as the rocket

briefly exited the Earth’s atmosphere.

Within the rocket, mirror-image

flow tunnels supported low-speed

oxidizer flow through each tunnel. An infrared laser diode was

mounted in the ceiling of one of

the two tunnels to provide external

radiant heating onto the black

polymethyl-methacrylate (PMMA,

an acrylic plastic) fuel surface.

Controlled external radiant heating

compensated for inherent heat

losses such as those from surface

radiation. The walls of the tunnels

had windows for ultraviolet-visible

and infrared imaging of the flame.

DARTFire was the first combustion

experiment to image flames in

these wavelengths, producing

unique images of microgravity

flames as they have never been

seen before.

DARTFire Sees Microgravity Fires in a New Light—Large

Data Base of Images Obtained

Filtered infrared multispectral images of a microgravity flame were obtained

during a DARTFire sounding rocket experiment. This figure is shown in color in

the online version of this article (http://www.grc.nasa.gov/www/RT1998/6000/

6711olson.htm)

CO, 4.8 μm

CO2, 4.3 μm

Fuel vapor, 3.4 μm Soot, 3.8 μm

CO2, 2.7 μm H2O, 1.8 μm

Microgravity Science
150

Red Green Blue

RGB composite CH, 430 nm OH, 310 nm

Ultraviolet and visible light (intensified) multispectral images of a microgravity

flame obtained during a DARTFire sounding rocket experiment. This figure is

shown in color in the online version of this article (http://www.grc.nasa.gov/WWW/

RT1998/6000/6711olson.htm).

Infrared emissions from the gaseous combustion products were monitored

in real time during the DARTFire experiment. An infrared camera was modified to include an internal rotating filter wheel that holds six different filters

so that it can look at radiant emissions from six different infrared bands as

the filter wheel rotates in front of the camera. The images in the preceding figure were obtained simultaneously with six different infrared filters,

including a fuel vapor image, a carbon monoxide (CO) image, a soot

image, two different wavelengths of carbon dioxide (CO2), and a water

vapor (H2O) image. This infrared imaging provided immense quantities

of data regarding the species fields in the experiment. A blackbody calibration of each filter provided effective blackbody temperature distributions

for each image, and numerical predictions of the temperature and species

fields were used to calculate the radiation distribution. The radiation distribution could then be compared directly with these calibrated images to

test our understanding of the experiment.

A multispectral intensified array video camera was also used to image the

flame in the ultraviolet and visible wavelength ranges. As with the infrared

camera, six filters were used to obtain intensified images of the dim flames

in different wavelengths, as shown in the following figure. Color images

were obtained by combining red, green, and blue filtered images to obtain

a 24-bit RGB color composite image of the flame. So that the combustion

intermediate (free radical) species could be imaged directly, filters were

used to capture the ultraviolet chemiluminescence of the methyl (CH*)

and hydroxyl (OH*) excited-state free radicals in the reaction zone of the

flames. The color images are what astronauts would see with their eyes.

The size, shape, and color of the

visible flame can be compared with

the large normal-gravity flame data

base. The chemilumiscent images

allow an extended view into the

heart of the reaction zone, where

detailed chemistry models of the

combustion process are sorely lacking experimental data.

The DARTFire experiment provided

the first comprehensive spectral

imaging of microgravity combustion flames. The effects of the flow

and external radiant heating on the

flame are well captured by this nonintrusive imaging. This large data

base of information is a wealth of

information for researchers interested in the spectral emissions of

microgravity flames. The DARTFire

hardware was built by the NASA

Lewis Research Center, and the

image analysis was supported by

Lewis’ Graphics Visualization

Laboratory.

Find out more on the World Wide

Web: http://zeta.grc.nasa.gov/expr3/

combust.htm

Bibliography

Olson, S.L., et al.: Diffusive and Radiative Transport in Fires Experiment:

DARTFire. Proceedings of the Fourth

International Microgravity Combustion

Workshop. NASA CP–10194, 1997.

Lewis contact:

Sandra L. Olson, (216) 433–2859,

Sandra.L.Olson@grc.nasa.gov

Authors:

Sandra.L. Olson, Uday Hegde, Robert A.

Altenkirch, and Subrata Bhatacharjee

Headquarters program office: OLMSA

Programs/Projects: Microgravity

Combustion Science, DARTFire

Special recognition:

Special Act and Service Awards

presented to the DARTFire team
151

Space

The spread of a flame over solid fuel is not only a fundamental textbook

combustion phenomenon, but also the central element of destructive fires

that cause tragic loss of life and property each year. Throughout history,

practical measures to prevent and fight fires have been developed, but

these have often been based on lessons learned in a costly fire. Since the

1960’s, scientists and engineers have employed powerful tools of scientific

research to understand the details of flame spread and how a material can

be rendered nonflammable. High-speed computers have enabled complex

flame simulations, and lasers have provided measurements of the chemical

composition, temperature, and air velocities inside flames. The microgravity

environment has emerged as the third great tool for these studies.

Spreading flames are complex combinations of chemical reactions and

several physical processes including the transport of oxygen and fuel vapor

to the flame and the transfer of heat from the flame to fresh fuel and to

the surroundings. Depending on its speed, air motion in the vicinity of the

flame can affect the flame in substantially different ways. For example,

consider the difference between blowing on a campfire and blowing out

a match. On Earth, gravity induces air motion because of buoyancy (the

familiar rising hot gases); this process cannot be controlled experimentally.

For theoreticians, buoyant air motion complicates the modeling of flame

spread beyond the capacity of modern computers to simulate. The microgravity environment provides experimental control of air motion near

spreading flames, with results that can be compared with detailed theory.

The Solid Surface Combustion

Experiment (SSCE) was designed

to obtain benchmark flame spreading data in quiescent test atmospheres—the limiting case of

flames spreading. Professor Robert

Altenkirch, Vice President for

Research at Mississippi State University, proposed the experiment

concept, and the NASA Lewis

Research Center designed, built,

and tested the SSCE hardware. It

was the first microgravity science

experiment built by Lewis for the

space shuttle and the first combustion science experiment flown in

space.

Hardware for SSCE consisted of a

sealed chamber module containing

the test samples, ashless filter

paper or polymethylmethacrylate

(PMMA, an acrylic plastic), and

a test atmosphere of 35- to 70-

percent oxygen mixed with nitrogen. The samples were instrumented with thermocouples. A

second module contained cameras,

a computer, and a battery.

SSCE has flown 10 times on the

space shuttle, with the final flight,

STS–91, occurring during 1998.

In each experiment, the samples

ignited successfully. The resulting

flames were recorded on film and

video; thermocouple temperatures

were recorded digitally.

Several scientifically significant

observations have emerged from

the SSCE tests. Steady flame spread

occurs over thin samples, even in

quiescent atmospheres. Thicker

PMMA samples burn unsteadily

and finally self-extinguish, but

they require as long as 9 min (e.g.,

the STS–85 test) to quench.

Solid Surface Combustion Experiment Yields Significant

Observations

Solid Surface Combustion Experiment hardware, including the chamber module

and camera module.
152

Comparisons of flight data with Professor Altenkirch’s numerical simulations indicate that (unlike in normal gravity) radiative heat transfer is a

dominating influence in flame spread in quiescent atmospheres. In summary, these experiments demonstrate that the controlling mechanisms for

material flammability and flame spread are quite different in microgravity

and normal Earth gravity. Results of the SSCE experiment will help spacecraft designers and Earth-bound fire safety technologists to improve

fire safety, through enhanced fire prevention, detection, and mitigation

techniques.

Lewis contact: Kurt. R. Sacksteder, (216) 433–2857,

Kurt.R.Sacksteder@grc.nasa.gov

Authors: Kurt R. Sacksteder, John M. Koudelka, and Franklin Vergilii

Headquarters program office: OLMSA

Programs/Projects: Microgravity Science, SSCE

Astronaut Tamara Jernigan performing the Solid Surface Combustion

Experiment in the Space Shuttle Columbia during STS–40.

Side-view image of unsteady flame

spread over a flat sample of PMMA

in a test atmosphere of 50-percent

oxygen/50-percent nitrogen at 1 atm

of pressure. The flame advanced

nearly 2 cm from the ignitor, then

regressed and finally extinguished,

requiring nearly 9 min to quench.
153

Space

C-97-4013

In practical combustion systems, the flame is often anchored at the inlet

where the fuel is injected into an air duct. This type of system is found

in powerplant combustors, gas turbine combustors, and the jet engine

afterburner. Despite its successful use, this configuration is vulnerable to

adverse flow conditions that can cause the flame to literally lift off from

the inlet or even blowout. Poor flame stability is, of course, unwanted,

especially where safety has a high priority. Our understanding of the mechanisms that control flame stability is incomplete in part because the interaction of buoyant (i.e., gravity-induced) convection makes it difficult to

Stability of Enclosed Laminar Flames Studied

in Microgravity

interpret normal-gravity results.

However, a comparison of normalgravity and microgravity results can

provide a clear indication of the

influence of forced and buoyant

flows on flame stability. Therefore,

a joint microgravity study on the

stability of Enclosed Laminar

Flames (ELF) was carried out by

researchers at The University of

Iowa and the NASA Lewis Research

Center. The microgravity tests were

conducted in the Microgravity

Glovebox (MGBX), during the

STS–87 space shuttle mission

in late 1997, using hardware

designed and produced at Lewis

and shown in the photo to the left.

The primary objective of the ELF

investigation was to determine

the mechanisms controlling the

stability of round, laminar, gas-jet

diffusion flames in a coflow air

duct. The study specifically focused

on the effect of buoyancy on the

flame characteristics and velocities

at the lift-off, reattachment, and

blowout of the flame. When the

ELF flight hardware assembled for testing.

Numerical predictions of flame lifting and shape change with increasing air velocity.

Heat release rate and mixture fraction for attached (left), lifted (center), and near

blowout (right) condition. Fuel-jet velocity, 0.2 m/sec.
154

fuel or air velocity is increased to a critical value, the flame base abruptly

jumps downstream, and the flame is said to have reached its lift-off condition. Flow conditions are such that the flame cannot be maintained at the

burner rim despite the presence of both fuel and oxygen. When the velocity is further increased, the flame eventually extinguishes at its blowout

condition. In contrast, if the velocity is reduced, the flame base eventually

returns to anchor at the burner rim, at a velocity lower than that of lift-off,

indicating a hysteresis effect.

During the STS–87 shuttle mission, approximately 50 tests were conducted

using a 50/50 mixture (by volume) of methane and nitrogen as the fuel.

Stable lifted flames were observed in microgravity, except at high fuel flows

where the microgravity flames blew out immediately after lift-off. The

experimental results verify the hypothesis that substantially greater velocities are required to destabilize a flame in microgravity because of the

absence of buoyant acceleration in the flow. Preliminary results reveal that

the increase in air velocity required to induce lift-off in microgravity (in

comparison to normal gravity) was nearly equal to the increase required

to induce blowout. Furthermore, the air velocity increase was relatively

independent of the fuel flow, except at low fuel flows. Preliminary numerical predictions are in qualitative agreement with the experimental results;

an example is shown in the bottom figure on the preceding page.

Further analysis is underway, and

on the basis of ELF’s success on

STS–87, a ground-based proposal

has been approved, offering ELF

researchers the opportunity to

compete for a reflight on the space

shuttle.

Enclosed Laminar Flames (ELF) is

described on the World Wide Web:

http://zeta.grc.nasa.gov/expr/elf.htm

Lewis contact:

Dennis P. Stocker, (216) 433–2166,

Dennis.P.Stocker@grc.nasa.gov

Author: Dennis P. Stocker

Headquarters program office: OLMSA

Programs/Projects: HEDS, Microgravity

Science, STS–87, USMP–4, MGBX, ELF

Most combustion processes in industrial applications (e.g., furnaces and

engines) and in nature (e.g., forest fires) are turbulent. A better understanding of turbulent combustion could lead to improved combustor

design, with enhanced efficiency and reduced emissions. Despite its

importance, turbulent combustion is poorly understood because of

its complexity. The rapidly changing and random behavior of such

flames currently prevents detailed analysis, whether experimentally or

computationally. However, it is possible to learn about the fundamental

behavior of turbulent flames by exploring the controlled interaction of

steady laminar flames and artificially induced flow vortices. These interactions are an inherent part of turbulent flames, and understanding them is

essential to the characterization of turbulent combustion. Well-controlled

and defined experiments of vortex interaction with laminar flames are not

possible in normal gravity because of the interference of buoyancy- (i.e.,

gravity) induced vortices. Therefore, a joint microgravity study was established by researchers from the Science and Technology Development Corp.

and the NASA Lewis Research Center. The experimental study culminated

in the conduct of the Turbulent Gas-Jet Diffusion Flames (TGDF) Experiment on the STS–87 space shuttle mission in November 1997. The fully

automated hardware, shown in photo, was designed and built at Lewis.

During the mission, the experiment was housed in a Get Away Special

(GAS) canister in the cargo bay.

Unlocking the Keys to Vortex/Flame Interactions in

Turbulent Gas-Jet Diffusion Flames—Dynamic Behavior

Explored on the Space Shuttle

In the TGDF experiment, toroidal

flow vortices were generated by

sinusoidally varying the open

diameter of an iris around the base

of a propane flame, thereby causing large-scale fluctuations in the

entrained airflow. The iris frequency,

which was varied to be 1.5, 3.0, and

5.0 Hz, was the experiment’s key

variable. TGDF’s primary objective

was to gain an understanding of the

dynamic behavior of the vortex/

flame interaction, including possible

vortex breakup or merging, and the

interaction’s effect on the timeaveraged values of flame temperature, radiation, and shape. In addition, the experiment has provided

insights into the development and

extinction of microgravity flames.

The experimental results were

compared with a comprehensive
155

Space

C-98-655

Interior of the TGDF combustion chamber, with the innovative vortex generation

mechanism visible in the center.

numerical model of the pulsed flame to validate the accuracy of the model

and verify our understanding of the flame behavior.

Analysis of the results revealed that the microgravity flame possesses three

axially spaced zones of vortex interaction. In the lower portion of the flame,

there is a significant transfer of energy from the mean field to the oscillations. In the numerical simulation, the vortex dissipation rate and the

kinetic energy of the oscillations both peak in the lower part of the flame.

In the central region, the energy

transfer is from the primary oscillation to its harmonic and/or the

mean field. In the upper portion of

the flame, the oscillations decay,

and there is no significant transfer

of energy between the oscillations

and the mean flow, or between

oscillations. In contrast, the oscillations grow in the upper portion

of buoyant normal-gravity flames.

Radiation measurements also

indicate that the core of the upper

zone receives energy from the

flame and the mean flow at the

flame base. Thus, the temperature

(and combustion) in the upper

zone is enhanced by the vortex

interaction. This has been confirmed by numerical computation

and is in agreement with observations of pulsation-induced flametip closure in supporting groundbased studies of microgravity

flames. Further analysis is underway, and a TGDF reflight proposal,

based on the successful STS–87

results, will be submitted.

The TGDF Experiment is described on

the World Wide Web: http://

zeta.grc.nasa.gov/expr2/gjdfe.htm

Lewis contact:

Dennis P. Stocker, (216) 433–2166,

Dennis.P.Stocker@grc.nasa.gov

Author: Dennis P. Stocker

Headquarters program office: OLMSA

Programs/Projects: HEDS, Microgravity

Science, STS–87, GAS, TGDF
156

0

5

10

15

20

0

1

2

3

4x10–2

0

2

4 6x10–3

Ampoule diameter, 

y, mm

Length solidified 

along ampoule, 

x, mm

Concentration of tin, at.%

2000 sec

5000 sec

8000 sec

The shuttle flight experiment “In Situ Monitoring of Crystal Growth Using

MEPHISTO” was accomplished during STS–87 as part of the fourth flight of

the United States Microgravity Payload (USMP–4), which was flown from

November 19 to December 5, 1997. The data returned from that flight are

just now beginning to yield quantitative results. This project is an international collaboration: the furnace system known as MEPHISTO was built in

France by CNES (French National Space Agency) and CEA (French Atomic

Energy Commission); the principal investigator, Prof. Reza Abbaschian, is

from the University of Florida at Gainesville; and numerical and analytical

modeling support includes collaborators from the University of New South

Wales, Australia, the University of Wisconsin at Milwaukee, the National

Institute of Standards and Technology, and the NASA Lewis Research

Center. MEPHISTO is a French acronym that translates into English as Materials for the Study of Interesting Phenomena of Solidification on Earth and

in Orbit. Since this was the fourth flight of the MEPHISTO furnace, the

experiment is referred to as MEPHISTO–4.

MEPHISTO–4 was a directional solidification experiment that studied the

liquid-to-solid transformation of bismuth alloyed with tin. Directional solidification is a freezing technique common to the processing of the electronic

materials used in integrated circuits and detectors, such as silicon and

germanium. When liquids are frozen on Earth, they must be cooled. The

cooling causes stirring because

of density variations in the liquid

(ref. 1). This stirring, known as

natural convection, influences the

quality of the resulting solid. During

freezing, regions of high and low

concentrations of tin are created.

This introduces another important

phenomenon: diffusion, or the

movement by molecular action

of matter from regions of high

concentration to regions of lower

concentration. In MEPHISTO–4, it

is tin that diffuses from the highconcentration region in front of

the solid-liquid interface to more

distant low-concentration regions.

Finally, the concentration gradients

of tin also cause density variations,

and hence are another source of

stirring of the liquid on Earth. The

interplay among temperature, concentration, convection, diffusion,

and solidification becomes extremely complex, and a complete

understanding of solidification is

beyond the present state of the art.

In MEPHISTO-4 the microgravity

environment is being used to

greatly reduce convection effects

in the solidification process.

In this way, preliminary theories

describing solidification in terms

of diffusion-only can be examined,

and the effects of diffusion can be

more readily assessed than is

possible on Earth, where natural

convection swamps the more

delicate diffusion phenomena.

When theories that include only

diffusion have been tested using

the flight data, more advanced

theories that include both diffusion

and convection can be developed

and tested using both flight and

Earth-based data.

All pertinent aspects of the

MEPHISTO–4 experiment were

monitored and controlled remotely

Crystal Growth Using MEPHISTO

Length solidified along the ampoule, x, versus the ampoule diameter from

bottom to top, y, and concentration of tin. No radial segregation is shown

after 5000 sec of growth, indicating diffusion dominance; radial segregation

is shown after 8000 sec because of solutal convection.
157

Space

through NASA’s Payload Operations Control Center located at the NASA

Marshall Space Flight Center. To monitor all the parameters needed for the

study of solidification, three samples were processed identically and simultaneously in the MEPHISTO furnace: one used the Seebeck technique (refs.

2 and 3) to measure the temperature at the interface between the solid

and liquid; the second employed a pulse of electricity to mark the shape

of the solid/liquid interface; and the third measured the rate of solidification by monitoring the electrical resistance of the sample. This third sample

was also quenched at the end of the experiment to freeze the concentration gradients built up in the liquid near the solid/liquid interface. As part

of the analysis of the flight sample, this quenched concentration profile will

be measured and used to calculate a diffusion coefficient for tin in liquid

bismuth. The diffusion coefficient is an important parameter in the analyses and theories of solidification. In the MEPHISTO furnace, several experiments can be done through the repeated melting, solidifying, and

remelting of the samples. Due in part to the results of the numerical

modeling teams (refs. 4 and 5), MEPHISTO–4 operations were extended

and continued through nearly the entire 16-day mission. The flexible and

responsive performance of the furnace, the optimization through numerical

modeling of the experiment, and the dedication of the principal investigator’s teams enabled about twice the number of experiments to be

accomplished during flight than were originally planned.

An example of the numerical modeling results is shown in the figure

(ref. 6). The modeling enabled the team to estimate the level and effects

of convection during flight. It also determined how long we needed to wait

between solidification runs for the high concentrations of tin at the solid/

liquid interface to be reduced to low levels by diffusion and convection,

thereby allowing our mission time to be optimized.

After the USMP–4 mission, the MEPHISTO hardware was shipped to the

French support contractor SEP (Société Européenne de Propulsion) for

checkout and removal of the samples. The samples and flight data are

now being analyzed by the University of Florida team. MEPHISTO–4 is

scheduled to be completed, with its final report given to NASA, by April

1999. Mission and flight data results were recently summarized (ref. 7)

and include measurement of the average diffusion coefficient of tin in

liquid bismuth of 3.5310–5 cm2/sec.

Find out more on the World Wide Web:

http://zeta.grc.nasa.gov/expr/insitu.htm

http://cml.grc.nasa.gov/degroh/

http://cfd.mech.unsw.edu.au/

http://www.uwm.edu/Dept/ECL/garimella.htm

References

1. de Groh, H.C., III; and Lindstrom, T.:

Interface Shape and Convection

During Solidification and Melting of

Succinonitrile. NASA TM–106487,

1994.

2. Abbaschian, R., et al.: A Study of

Directional Solidification of Faceted

Bi-Sn Alloys in Microgravity. AIAA

Paper 95–0608, 1995.

3. Alvarez, J.; Peteves, S.D.; and

Abbaschian, G.J.: Thermoelectric

Effects Across Solid-Liquid Gallium

Interface. Thin Films and Interfaces

II. J.E.E. Baglin, D.R. Campbell, and

W.K. Chu, eds. MRS Symposium

Proceedings, vol. 25, 1984,

pp. 345–350.

4. Chen, P., et al.: Rehomogenization—

Diffusion and Correction in Microgravity. AIAA Paper 98–0740, 1998.

5. Yao, M.; de Groh, H.C., III; and

Abbaschian, R.: Numerical Modeling

of Solidification in Space with

MEPHISTO–4 (Part I). AIAA Paper

97–0449, 1997.

6. Timchenko, V., et al.: Directional

Solidification in Microgravity. Heat

Transfer 1998, Proceedings of 11th

IHTC, Kyongju, Korea, Taylor &

Francis, vol. 7, pp. 515–520, 1998.

7. Cambon, G., et al.: MEPHISTO/

USMP–4/STS–87, In-Situ Monitoring

of Crystal Growth Using MEPHISTO,

Preliminary Flight Results. IAF/IAA–

98–J.3.04, 49th International

Astronautical Congress, 1998.

Lewis contact:

Henry C. de Groh III, (216) 433–5025,

Henry.C.deGroh@grc.nasa.gov

Author: Henry C. de Groh III

Headquarters program office: OLMSA

Programs/Projects: Microgravity

Materials Science Program, USMP–4

Special recognition: Best poster in

session, 11th International Heat

Transfer Conference, Korea, 1998
158

The microgravity environment offers the potential to measure the binary

diffusion coefficients in liquids without the masking effects introduced by

buoyancy-induced flows due to Earth’s gravity. However, the background

g-jitter (vibrations from the shuttle, onboard machinery, and crew) normally encountered in many shuttle experiments may alter the benefits of

the microgravity environment and introduce vibrations that could offset

its intrinsic advantages. An experiment during STS–85 (August 1997) used

the Microgravity Vibration Isolation Mount (MIM) to isolate and introduce

controlled vibrations to two miscible liquids inside a cavity to study the

effects of g-jitter on liquid diffusion.

Diffusion in a nonhomogeneous liquid system is caused by a nonequilibrium condition that results in the transport of mass (dispersion of the

different kinds of liquid molecules) to approach equilibrium. The dynamic state of the system tends toward equilibrium such that the system

becomes homogeneous. An everyday example is the mixing of cream and

coffee (a nonhomogeneous system) via stirring. The cream diffuses into

the coffee, thus forming a homogeneous system. At equilibrium the system

is said to be mixed. However, during stirring, simple observations show

complex flow field dynamics—stretching and folding of material interfaces, thinning of striation thickness, self-similar patterns, and so on. This

example illustrates that, even though mixing occurs via mass diffusion,

stirring to enhance transport plays a major role. Stirring can be induced

either by mechanical means (spoon or plastic stirrer) or via buoyancyinduced forces caused by Earth’s gravity. Accurate measurements of binary

diffusion coefficients are often inhibited by buoyancy-induced flows. The

microgravity environment minimizes the effect of buoyancy-induced

flows and allows the true diffusion limit to be achieved. One goal of this

experiment was to show that the microgravity environment suppresses

buoyancy-induced convection, thereby mass diffusion becomes the dominant mechanism for transport. Since g-jitter transmitted by the shuttle to

the experiment can potentially

excite buoyancy-induced flows,

we also studied the effects of

controlled vibrations on the system.

The physics of the interface

behavior of two miscible liquids

subjected to steady and oscillatory

body forces was predicted, and

the results were documented in a

series of publications (refs. 1 to 4).

The experiment was conceived and

designed by W.M.B. Duval at the

NASA Lewis Research Center. However, because an important finding

of the theoretical prediction was

the presence of standing waves

due to oscillatory body forces (refs.

1 and 3), a special experimental

apparatus was needed. Showing

the existence of standing waves

required the use of a platform that

could both isolate the experiment

from vibrations and input known

vibration sources to it. These

requirements were fulfilled by the

use of the Microgravity Vibration

Isolation Mount (MIM), designed

by B.V. Tryggvason through an international collaboration with the

Canadian Space Agency. The experimental setup flown on STS–85 is

shown in the photo to the left.

Hardware consisted of the mixing

cell, a light source, a motor to

remove the shim, and a black &

white camera that was mounted

on the MIM platform. Experiments

on the MIM were performed in two

different modes: (1) an isolation

mode to damp out background

g-jitter and (2) an active mode to

input known vibration sources to

the experiment.

On the next page, the top and

bottom photos, respectively, show

ground-based and microgravity

results. Because of the large size

of the mixing cell, 5-cm high by

5-cm wide by 1-cm deep, it was

extremely challenging to damp

Effects of g-Jitter on Diffusion in Binary Liquids

Experiment mounted on the MIM platform on STS–85
159

Space

Ground-based experiment. Left: Initial condition of two miscible liquids at time zero. Fluid on the left side of the

photo is 2.2-vol % deuterium oxide with deionized water and blue dye; fluid on the right side of the photo is

100-vol % deionized water with red dye—a specific gravity difference of 0.0026. Right: Configuration of the

interface after removal of the shim showing an internal breakwave after approximately 7 sec.

Microgravity experiments. Left: Isolation mode 17 sec after removal of the shim, showing a stationary interface

(fluid on the left side of the photo contains 20-vol % ethylene-glycol with water and blue dye; fluid on the right

side of the photo contains 100-vol % deionized water with red dye—a specific gravity difference of 0.028).

Right: Forcing mode at a frequency of 1 Hz and an amplitude of 0.02g, showing a four-mode standing wave.

out buoyancy-induced forces, which can stretch and fold the interface. The

top photos capture a typical scenario showing the effect of a steady body

force on the ground. The initial condition at time zero consists of two

miscible liquids separated by a thin shim, as in the top left photo. Once

the shim is pulled, the hydrostatic pressure imbalance between the two

liquids causes an overturning motion that stretches and folds the interface.

The top right photo shows the evolution of an internal breakwave after

approximately 7 sec.

In contrast, the microgravity result in the bottom left photo shows a stationary interface after 17 sec. The MIM was operating in isolation mode;

the background acceleration was nearly steady on the order of a micro-g

(10–6g). This result illustrates the effectiveness of the microgravity environment in damping buoyancy-induced flows, thus permitting liquids to mix

via mass diffusion. Note that the presence of any residual buoyancyinduced force would cause the interface to deform. Mass diffusion over

a long time scale would cause the

two fluids to mix while the interface remained stationary in the

vertical configuration. Given that

the interface can remain stationary

over a time interval, the MIM was

operated in the excitation mode

with a frequency of 1 Hz and an

amplitude of 0.02g with a sinusoidal input. The bottom right photo

shows that the interface becomes

unstable against Kelvin-Helmholtz

instability and produces a fourmode standing wave. In this case,

the instability of the interface

enhances local mass transport.
160

The analysis of the vibration data input by the MIM on the experiment is

ongoing. Recent computational results in fiscal year 1998 agreed with the

experimental findings of the four-mode standing wave; however, subtleties of the experimental findings are being addressed. These results are

expected to shed light on the use of isolation systems for the International Space Station.

Visit Lewis on the World Wide Web: http://www.grc.nasa.gov

References

1. Duval, W.M.B.; and Jacqmin, D.A.: Interfacial Dynamics of Two Liquids Under an

Oscillating Gravitational Field. AIAA J., vol. 28, Nov. 1990, pp. 1933–1941.

2. Duval, W.M.B.: Numerical Study of Mixing of Two Fluids Under Low Gravity. NASA

TM–105865, 1992.

3. Duval, W.M.B.: The Kinematics of Buoyancy Induced Mixing. Proceedings of

the 8th European Symposium on Materials and Fluid Sciences in Microgravity,

ESA–SP–333, Vol. 2, 1992, pp. 855–861.

4. Duval, W.M.B.: Flow Field Topology of Buoyancy Induced Mixing. AIAA Paper

96–0253, 1996.

Lewis contact:

Walter M.B. Duval, (216) 433–5023,

Walter.M.Duval@grc.nasa.gov

Author: Walter M.B. Duval

Headquarters program office:

OLMSA (UG)

Programs/Projects:

Microgravity Science

Special recognition:

Superior Accomplishment Award given

in 1998 for space flight experiment

conducted on STS–85 in August 1997

Why does water flow faster than honey? The short answer, that honey has

a greater viscosity, merely rephrases the question. The fundamental answer

is that viscosity originates in the interactions between a fluid’s molecules.

These interactions are so complicated that, except for low-density gases,

the viscosity of a fluid cannot be accurately predicted. Progress in understanding viscosity has been made by studying moderately dense gases and,

more recently, fluids near the critical point. Modern theories predict a universal behavior for all pure fluids near the liquid-vapor critical point, and

they relate the increase in viscosity to spontaneous fluctuations in density

near this point. The Critical Viscosity of Xenon (CVX) experiment tested

these theories with unprecedented precision when it flew aboard the

Space Shuttle Discovery (STS–85) in August 1997.

Near the critical point, xenon is a billion times more compressible than

water, yet it has about the same density. Because the fluid is so “soft,”

it collapses under its own weight when exposed to the force of Earth’s

gravity—much like a very soft spring. Because the CVX experiment is

conducted in microgravity, it achieves a very uniform fluid density even

very close to the critical point. The accompanying graph reveals the

dramatic difference between Earth-based measurements and the data

obtained in microgravity.

At the heart of the CVX experiment is a novel viscometer built around a

small nickel screen. An oscillating electric field forces the screen to oscillate

between pairs of electrodes. Viscosity, which dampens the oscillations, can

be calculated by measuring the screen motion and the force applied to the

screen. So that the fluid’s delicate state near the critical point will not be

disrupted, the screen oscillations are set to be both slow and small.

CVX is the culmination of a series of Earth-based viscosity measurements

near critical points. Robert F. Berg and Michael R. Moldover of the National

Institute of Standards and Technology are the principal investigators,

and Gregory Zimmerli of Lewis and

the National Center for Microgravity

Research is the project scientist.

NASA funded both the Earth-based

and flight experiments through

Lewis’ Microgravity Science Division. NASA Lewis science contractors, NYMA, ADF, and Sverdrup,

developed and built the flight

instrument.

Since the flight, the principal investigators have analyzed the flight

data in great detail. The data show

that the viscosity increases as the

critical point is approached and

that it agrees with the theoretical

prediction. CVX also measured lowfrequency viscoelasticity in a pure

fluid for the first time. Viscoelasticity is a partly elastic response to

shear stress. It is common in

polymers, foams, and other complex fluids, but has never been

observed in a simple fluid such as

xenon. A journal publication

describing the scientific findings

has been published (ref. 1).

Viscosity of Xenon Examined in Microgravity
161

Space

10–4 10–3 10–2 10–1 100 101 102

Viscosity, 

μPa-sec

50

60

70

CVX calibration

 at Tc + 1 K

CVX

microgravity

CVX

ground

High-Q oscillator

ground

Asymptotic

power law

(T – Tc) / K

Ground-based and microgravity measurements of the viscosity of xenon near the

critical point. In microgravity, the increase of viscosity is measured 100 times closer

to the critical temperature, Tc , and much more accurately than on Earth. The

asymptotic power law characterizes the universal behavior of viscosity in all pure

fluids near the critical point. Temperature of sample, T; quality factor, Q. (High-Q

indicates weak damping.)

As an extension to this very successful flight experiment, the principal

investigators are preparing a second flight experiment, CVX–2. This experiment, which will measure the shear thinning of viscosity near the critical

point of xenon, will be the first experiment of its kind to measure shear

thinning in a simple fluid. A quantitative understanding of shear

thinning near the critical point will

lead to a better understanding of

shear thinning in more complex

fluids, such as polymers and

emulsions, and thus may have

broad applications in science and

industry.

Reference

1. Berg, R.F.; Moldover, M.R.; and

Zimmerli, G.A.: Viscoelasticity of

Xenon near the critical point. Phys.

Rev. Lett., vol. 82, Feb. 1, 1999,

pp. 920–923.

Lewis contacts:

Richard W. Lauver, (216) 433–2860,

Richard.W.Lauver@grc.nasa.gov;

Gregory A. Zimmerli, (216) 433–6577,

Gregory.A.Zimmerli@grc.nasa.gov;

Dr. Bhim S. Singh, (216) 433–5396,

Bhim.S.Singh@grc.nasa.gov; and

Susan M. Motil, (216) 433–8589,

Susan.M.Motil@grc.nasa.gov

Authors: Gregory A. Zimmerli, Robert F.

Berg, and Michael R. Moldover

Headquarters program office:

OLMSA (UG)

Programs/Projects:

Microgravity Science

Laser light scattering is the technique of choice for noninvasively sizing

particles in a fluid. The members of the Advanced Technology Development (ATD) project in laser light scattering at the NASA Lewis Research

Center have invented, tested, and recently enhanced a simple and elegant

way to extend the concentration range of this standard laboratory particlesizing technique by several orders of magnitude. With this technique,

particles from 3 nm to 3 μm can be measured in a solution. But, until

recently, it was difficult at best to look into the depths of fluids that are not

as clear as water. With higher concentrations of particles, the large increase

in multiple scattering would corrupt the data.

Recently, laser light scattering evolved to successfully size particles in both

clear solutions and concentrated milky-white solutions. The enhanced

technique uses the property of light that causes it to form tall interference

patterns at right angles to the scattering plane (perpendicular to the laser

Laser Light Scattering With Multiple Scattering Suppression

Used to Measure Particle Sizes

beam) when it is scattered from a

narrow laser beam. Such multiplescattered light forms a broad fuzzy

halo around the focused beam,

which, in turn, forms short interference patterns. By placing two fiber

optics on top of each other and

perpendicular to the laser beam

(see the drawing), and then crosscorrelating the signals they produce, only the tall interference

patterns formed by singly scattered

light are detected. To restate this,

unless the two fiber optics see the

same interference pattern, the
162

scattered light is not incorporated into the signal. With this technique, only

singly scattered light is seen (multiple-scattered light is rejected) because

only singly scattered light has an interference pattern tall enough to span

both of the fiber-optic pickups. This technique is simple to use, easy to

align, and works at any angle. Placing a vertical slit in front of the signal

collection fibers enhanced this approach. The slit serves as an optical

mask, and it significantly shortens the time needed to collect good data

by selectively masking out much of the unwanted light before crosscorrelation is applied.

Different scattering angles and polarizations can be used to extract additional information from the samples—information that was previously

unavailable and not theoretically interpreted. Both of these concerns are

being addressed, and we believe that this will have a significant impact in

the particle-sizing community.

In addition, the Advanced Technology Development light-scattering group

at Lewis is extending the capabilities of laser light scattering by using fiber

optics to overcome the problems of stray light (which until now could

corrupt the interpretation of otherwise good data). Launching the laser

beam used to probe the particles both into the solution and into a fiber

optic that is coupled to the signal-receiving fiber will introduce a local

oscillator at the detector. This overcomes the problems introduced by a

small amount of stray light. It also provides significant (homodyne) gain

for the signal and eliminates many of the cross-terms that would otherwise

make data analysis more complicated. Although bulk-optic homodyne

instruments have been used in the past, alignment problems made these

Diagram of multiangle laser-light-scattering instrument with multiplescattering suppression using single-color cross-correlation. Both fibers

are stacked perpendicular to the scattering plane.

Selectable

angle Sample

cell

Laser beam

Beam

stop

Fiber-optic

pair to

detector pair

Fiber-optic

pair to

Lens detector pair

difficult, onerous beasts (especially

when multiple-angle measurements were being made). With

fiber optics, homodyning is much

easier. Mixing the scattered light

signal and the laser beam in

polarization maintains monomode

fiber optics and is 100-percent

efficient because of the wavefront

matching imposed by the fiberoptic couplers.

Lewis’ light-scattering group is

continuing to provide new laserlight-scattering technologies and

techniques that help answer

fundamental science questions—

questions that can be answered by

studying colloid particle interactions and fluid-crystal-glass phase

transitions in space (where gravity

does not destroy the study with

conventional effects). These studies

serve as a good model for atomic

interactions and crystal formation

and have led to a number of

pleasant surprises and advances.

The commercial biotechnology

community is also asking Lewis’

light-scattering group about using

light scattering to study proteins.

Further interest is surfacing in the

computer industry, where polishing

compounds grind themselves up

when they are used to make computer chips. Here, constant monitoring of the particle grit size is

needed to increase silicon wafer

production yield. Other opportunities are also being addressed.

Lewis contact:

William V. Meyer, (216) 433–5011,

William.V.Meyer@grc.nasa.gov

Authors:

William V. Meyer, Padetha Tin, James A.

Lock, David S. Cannell, Anthony E.

Smart, and Thomas W. Taylor

Headquarters program office:

OLMSA (UG)

Programs/Projects:

Microgravity Science, NCMR, ATD
163

Space

Phase-shifted, laser feedback interferometry is a new diagnostic tool developed at the NASA Lewis Research Center under the Advanced Technology

Development (ATD) Program directed by NASA Headquarters’ Microgravity

Research Division. It combines the principles of phase-shifting interferometry (PSI) and laser-feedback interferometry (LFI) to produce an instrument

that can quantify both optical path length changes and sample reflectivity

variations. In a homogenous medium, the optical path length between two

points is the product of the index of refraction and the geometric distance

between the two points. LFI differs from other forms of interferometry by

using the laser as both the source and the phase detector. In LFI, coherent

feedback of the incident light either reflected directly from a surface or

reflected after transmission through a region of interest will modulate the

output intensity of the laser. The combination of PSI and LFI has produced

a robust instrument, based on a low-power helium-neon (HeNe) gas laser,

with a high dynamic range that can be used to measure either static or

oscillatory changes of the optical path length. Small changes in optical path

length are limited by the fraction of a fringe that can be measured; we can

measure nonoscillatory changes with a root mean square (rms) error of the

wavelength/1000 without averaging.

The photograph on the left shows the phase-shifted, laser feedback interferometer installed in a replica of the “glovebox” that was used aboard

the space shuttle. The laser can be observed just behind the center panel

(which was intentionally left open). To the left of the laser is a photodetector, and to the right is the electro-optic modulator (gold box) used to introduce the phase shifts. To achieve higher spatial resolution, we coupled the

interferometer with a high numerical aperture objective and placed the

sample to be scanned at the focus

of the objective. The glovebox was

suspended from elastic cords to

isolate the apparatus from vibrations. The photograph on the right

shows an expanded view of the

electro-optic modulator, the corner

cube used to bend the laser beam,

and the microscope objective. A

demonstration of this apparatus

was recently provided at the Third

Microgravity Fluid Physics Conference, and changes in the optical

path length as small as 5 nanometers (nm) were readily achieved

even during times of heavy pedestrian traffic.

We also combined the phaseshifted, laser feedback interferometer with a reflecting light

microscope and obtained images

of contours of static fluid drops on

Phase-Shifted Laser Feedback Interferometry

Phase-shifted, laser feedback interferometer installed in a replica of

a “glovebox” used aboard the space shuttle. The light from a

helium-neon laser (black cylinder) illuminates an object, which is

scanned under the laser beam. The gold box between the laser and

the object is a phase modulator that is used to introduce controlled

phase shifts. A photodetector (black box on the far left side) is used

to monitor the laser’s intensity.

Scanning of an object in the

interferometer. After the light leaves

the phase modulator (gold box), it

crosses a polarizer and is bent

downward by a corner cube (glass)

and focused by a microscope

objective. An object is scanned across

the focused laser beam to determine

its topology.
164

Optical path length through a silicone oil drop on a coated silicon

substrate. A spherical drop has been correctly reconstructed.

coated silicone substrates. The preceding figure shows the optical path

length through a silicone oil drop. This image, which was obtained by

scanning the drop beneath a fixed laser beam, clearly shows the edge of

the drop. We are currently pursuing the measurement of dynamic fluidsolid interfaces.

Bibliography

Ovryn, B.; and Andrews, J.H.: PhaseShifted Laser Feedback Interferometry.

Opt. Lett., vol. 23, no. 14, July 15, 1998,

pp. 1078–1080.

Lewis contact:

Dr. Ben Ovryn, (216) 433–8335,

Ben.Ovryn@grc.nasa.gov

Author: Dr. Ben Ovryn

Headquarters program office:

OLMSA (UG)

Programs/Projects:

Microgravity Science, ATD

Special recognition: Named Innovator

of the Year 1997 by NYMA

Microscopic spheres suspended in liquid become highly ordered under

the proper conditions. Such collections of particles, called colloidal suspensions or colloids, are the subject of a series of ongoing microgravity experiments at the NASA Lewis Research Center. By studying the way these

colloidal suspensions order themselves, scientists can better understand

how atoms of a liquid become ordered to form a solid. In addition, highly

ordered colloids have special properties that may make them useful in

future high-tech applications. Work is underway at Lewis to develop an

optical microscope to view these colloidal suspensions sphere by sphere

in microgravity.

Previous Lewis experiments have studied the average properties and

macroscopic features of colloidal suspensions using laser light scattering

techniques and macrophotography. However, different techniques are

needed to observe the behavior of individual spheres at specific locations

within a suspension. Using a commercially available optical microscope,

one can see local features and phenomena, as in the photo at the top of

the next page.

The initial thrust in developing a

microscope for microgravity

research was to build up Lewis’

expertise in applying optical

microscopy to colloidal suspensions and to determine and

demonstrate the core capabilities

required for microgravity experiments. Many of the colloidal

suspensions of interest contain

50 to 60 vol % spheres. To image

beyond the surface of these highvolume-fraction colloids, one must

closely match the optical index of

refraction of the liquid and the

spheres, rendering the spheres

invisible under normal illumination. However, recently at Lewis

Microgravity Experiments Being Developed for Microscopic

Study of Colloidal Spheres

μm

μm

μm
165

Space

Grain boundaries Phase boundary

10 μm

Edge dislocation CD-98-77749

50 μm 50 μm 10 μm

Photomicrograph illustrating various local features in a suspension of

1.0-μm-diameter spheres.

Left: Darkfield illumination. Colloidal crystal grains are delineated by color variations produced by Bragg

scattering from different crystallographic planes. Inset: Oblique illumination of same region. Right: Differential

Interference Contrast technique image of dark-bordered region of image on left. Note grain boundary running

from lower left to upper right. The left image is shown in color in the online version of this article (http://

www.grc.nasa.gov/WWW/RT1998/6000/6712rogers.htm).

several well-established optical contrast techniques from the biological and

materials science fields were combined with real-time image-processing

hardware to successfully image high-volume-fraction suspensions of

spheres ranging from 0.5 to 1.2 μm in diameter. Both Nomarski Differential

Interference Contrast (DIC) and phase contrast techniques have produced

images of colloidal crystals well beyond the walls of experiment cells. In

addition, darkfield and oblique illumination techniques have been used

to distinguish between colloidal crystal grains, bridging the gap between

microscopic and macroscopic imaging as shown in the following figure.

NASA Lewis and Dynacs Engineering Company, Inc., engineers have begun

developing concepts for the flight instrument, currently planned to be the

first piece of experiment hardware

in the Fluids Integrated Rack

aboard the International Space

Station. This microscope will

support the Physics of Hard

Spheres Experiment-2 (PhaSE–2;

principle investigator, Paul Chaikin

of Princeton University) and the

Physics of Colloids in Space-2

experiments (PCS–2; principal

investigator, David Weitz of the

University of Pennsylvania).

Current plans call for the contrast

techniques described here, lightscattering optics, confocal imaging,

on-orbit sample homogenization,

and the capability for programmed

and ground-controlled operations.

In addition to the microgravity

activities just described, hardware

is nearly complete for a separate

series of ground-based experiments that actually depend on

gravity to control the formation of

colloidal crystals. These experiments, supported by the Lewis

Director’s Discretionary Fund

(DDF), will use a fixture that allows

the microscope to rotate to various

angles with respect to gravity. The

first use of this apparatus will be to

study the sedimentation-induced

growth of long, thin columns of
166

colloidal crystals that bear similarities to columnar growth commonly seen

in commercial metal castings. The thrust of this DDF-sponsored research is

to begin applying the results of colloidal sphere experiments to a wide

range of materials science topics.

Lewis contacts:

Richard B. Rogers, (216) 433–6512, Richard.B.Rogers@grc.nasa.gov; and

Michael P. Doherty, (216) 433–6641, Michael.P.Doherty@grc.nasa.gov

Author: Richard B. Rogers

Headquarters program office: OLMSA

Programs/Projects: Microgravity Science, PHASE–2, PCS–2

Under certain conditions, such as in thin liquid films or microgravity, surface tension variations along a free surface can induce convection. Convection onset due to surface tension variation is important to many terrestrial

technological processes in addition to microgravity materials processing

applications. Examples include coating, drying crystallization, solidification,

liquid surface contamination, and containerless processing. In doublediffusive and multicomponent systems, the spatial variations of surface

tension are associated with two or more stratifying agencies, respectively.

For example, both temperature and species (concentration) gradients are

associated with convection in the solidification of binary alloys or salt

ponds. The direction of the two (or more) gradients has a profound effect

on the nature of the flow at or slightly beyond the onset of convection, as

illustrated in the preceding photos. In the photo on the left, salt fingers

have formed at the interface between a 10-percent sugar-water solution

over a 12-percent salt-water solution. In the photo on the right, a salt-water

solution, the temperature has increased from bottom to top while the salt

concentration increased from top to bottom. In this case, oscillatory instability occurred via a “Hopf bifurcation,” resulting in the diffusive convection

pattern observed. (We are indebted to Prof. C.F. Chen for providing these

photographs.) The fluid properties and operating conditions that lead to a

specific convection onset behavior can be predicted from neutral stability

curves and maps, like those of the remaining figures (which are shown in

color in the online version of this article: http://www.grc.nasa.gov/WWW/

RT1998/6000/6712skarda.htm).

Our recent work at the NASA

Lewis Research Center focused

on characterizing surface-tensioninduced onset of convection, often

referred to as Marangoni-Benard

convection. Exact solutions for the

stationary neutral stability of multicomponent fluid layers with interfacial deformation were derived.

These solutions also permit the

computation of a boundary curve

that separates the long and finite

wavelength instabilities, which is

shown in the following graph.

Computing points along this

boundary using the exact solution (when possible) is more efficient than the typical numerical

approaches, such as finite difference or spectral methods. Above

the curve, a long wavelength instability was predicted, suggesting that

convection would occur principally

through one large flow cell in the

layer; whereas below the curve,

finite wavelength instabilities occur,

which suggest multiple finite-sized

circulation cells. For many common

liquids with layer depths greater

than 100 μm, finite wave instability is predicted under terrestrial

conditions; however, with little

exception, long wavelength instability is predicted in microgravity

for the identical fluid systems.

Onset of Convection Due to Surface Tension Variations in

Multicomponent and Binary Fluid Layers

Left: Salt fingering formation. Right: Onset of double-diffusive instability (ref. 1).

(Copyright Professor C.F. Chen, Univ. of Arizona; used with permission.)
167

Space

Frontier boundary

Extremum boundary for Ma Σ(α = 0)

10–1

10–1 100

10–2

10–2

10–2

10–3

10–3

10–4

10–4

10–1 100 10–2 10–3 10–4

10–5

10–5

10–6

10–6

10–7

10–7

10–8

10–9

10–9

Bo

Bo

Finite critical

wavenumber modes

Infinitesimal critical wavenumber modes

(long wavelength modes)

D Cr

D Cr

D

Dk

MaΣ

Mak

= k=1

Σ

N

Ma Σ(α = 0) is 

local minimum

Ma Σ(α = 0) is 

local maximum

Effect of diffusivity ratio, τ , on stability maps in (Smc , Mac ) space, where Smc

and Mac are the critical surface tension Soret number and Marangoni number,

respectively. The concentration difference is induced by an applied temperature

difference. Oscillatory instability occurs above the oscillatory boundaries in the

upper right quadrant.

The governing equations for double-diffusive systems that include crossdiffusion, such as the Soret and Dufour effects, were rescaled for limitzero buoyancy, and exact solutions were derived for the stationary neutral

stability of these systems. An extensive investigation of stationary and

oscillatory instabilities was performed. It spanned parameter values associated with binary systems, such as lead-tin alloys, to those associated with

water-alcohol mixtures. Stability maps similar to the one shown in the

following graph were developed to characterize the fluid behavior. Exact

solutions for the location of the

asymptotes, also shown in this

graph, are suitable for validating

the computational schemes that

will ultimately be used to analyze

the behavior of more complex

systems.

Researchers at Lewis are also

examining the response of a fluid

layer to time periodic accelerations or gravity modulation. For

Marangoni-Benard type convection,

gravity modulation can stabilize an

unstable (or destabilize a stable)

unmodulated system. Results such

as those in the preceding stability

plots can be used to assess the

sensitivity of these fluid systems

to residual accelerations. The

response of a double-diffusive

fluid system to modulation can be

more complex, exhibiting regions

of quasi-periodic, subharmonic,

and synchronous behavior along

the neutral stability curve, as

shown in the final graph. For more

information and definitions of

symbols used in the figures, see

references 2 to 4.

400

200

0

–200

–400

–0.50 –0.25 0.00 0.25 0.50

Unstable

Stable Unstable

StationaryOscillatoryDiffusivity

ratio,

τ

Mac

Smc

τ= 0.0001

τ= .01

τ= .10

τ= 1.00

Frontier point boundaries in (Bo, D Cr) space. The dashed line in the insert is the extremum boundary for MaΣ (α = 0). Above the dashed line, MaΣ (α = 0) is a local minimum; below this line, it is a

local maximum.

~

~
168

0.00

14

12

10

8

6

4

2

0

g1

0.04 0.08 0.12 0.16

1/Ω

Synchronous instability

Subharmonic instability

300

200

100

0

–100

0123456

α

Ma

Modulated neutral stability curve

Unmodulated neutral stability curve

400

500

300

200

100

0

012345

α

Ma

Quasi-periodic

Subharmonic

Bifurcation points

Unmodulated

Left: Stability boundaries in (1/Ω, g1) space; α = 2, Ra = 1000, Ma = 118.77,

Pr = 1, go = 0. Right: Neutral stability curve; Pr = 1, go = 0, g1 = 5, Ra = 1000.

Neutral stability curves for gravitymodulated double diffusion;

Ms = –700, Pr = 10, D22 = 0.1,

Ra = 1000, g0 = 0, g1 = 1, Ω = 5.

References

1. Tanny, J.; Chen, C.C.; and Chen, C.F.: Effects of Interaction Between Marangoni

and Double-Diffusive Instabilities. J. Fluid Mech., vol. 303, 1995, pp. 1–21.

2. Skarda, J.R.L.; Jacqmin, D.; and McCaughan, F.E.: Exact and Approximate Solutions to the Double-Diffusive Marangoni-Benard Problem With Cross-Diffusive

Terms. J. Fluid Mech., vol. 366, 1998, pp. 109–133.

3. Skarda, J.R.L.; and McCaughan, F.E.: Exact Solution to Stationary Onset of

Convection Due to Surface Tension Variation in a Multicomponent Fluid Layer

With Interfacial Deformation. Int. J. Heat Mass Transfer, vol. 42, no. 13, 1999,

pp. 2387–2398.

4. Skarda, J.R.L.: Convective Instability of a Gravity Modulated Fluid Layer With

Surface Tension Variation. AIAA Paper 98–2599 (NASA TM—1998-207941), 1998.

For more information, visit us on the

World Wide Web:

http://zeta.grc.nasa.gov/6712/

home6712.htm

http://zeta.grc.nasa.gov/6712/people/

skarda.htm

Lewis contact: Dr. J. Raymond Lee

Skarda, (216) 433–8728,

J.R.Skarda@grc.nasa.gov

Author: Dr. J. Raymond Lee Skarda

Headquarters program office: OLMSA

Programs/Projects:

Microgravity Fluid Physics Science

Experimenters from the fluids, combustion, materials, and life science

disciplines all use the microgravity environment of space to enhance their

understanding of fundamental physical phenomena caused by disturbances from events such as spacecraft maneuvers, equipment operations,

atmospheric drag, and (for manned flights) crew movement. Space conditions reduce gravity but do not eliminate it. To quantify the level of these

disturbances, NASA developed the Space Acceleration Measurement

System (SAMS) series to collect data characterizing the acceleration environment on the space shuttles. This information is provided to investigators

so that they can evaluate how the microgravity environment affects their

experiments. Knowledge of the microgravity environment also helps investigators to plan future experiments.

The original SAMS system flew 20 missions on the shuttle as well as on

the Russian space station Mir. Presently, Lewis is developing SAMS–II for

the International Space Station; it will be a distributed system using digital

output sensor heads. The latest operational version of SAMS, SAMS–FF,

was originally designed for free

flyer spacecraft and unmanned

areas. SAMS–FF is a flexible, modular system, housed in a lightweight

package, and it uses advances in

technology to improve performance. The hardware package

consists of a control and data

acquisition module, three different types of sensors, data storage

devices, and ground support equipment interfaces.

Three different types of sensors

are incorporated to measure both

high- and low-frequency accelerations and the roll rate velocity.

Space Acceleration Measurement System for Free Flyers
169

Space

SAMS–FF–1: Implementation of TSH with laptop computer and battery.

SAMS–FF–2: TSH in hermetic enclosure.

Small, low-power triaxial sensor heads (TSH’s) offer high resolution and

selectable bandwidth, and a special low-frequency accelerometer is available for high-resolution, low-frequency applications. A state-of-the-art,

triaxial fiber-optic gyroscope that measures extremely low roll rates is

housed in a compact package.

The versatility of the SAMS–FF system is shown in the three different types

of missions SAMS–FF has supported. The first mission was on a sounding

rocket supporting NASA Lewis Research Center’s DARTFire combustion

experiment on September 10, 1997. The results indicated that a sounding

rocket is a very good platform to conduct short-duration research under

high-quality microgravity conditions.

A SAMS–FF TSH was used to support experiments flown on the NASA

KC–135, flying parabolic arcs for low-gravity conditions. This TSH, which

was utilized standalone connected to a laptop computer, served as a compact acceleration measurement system. Its small size and low power

requirements allowed it to be mounted close to the area of interest.

SAMS–FF was requested by the NASA Goddard Space Flight Center to

support the Hubble Space Telescope program. On STS–95, a SAMS–FF

system consisting of a control and data acquisition unit and two TSH’s

was mounted on the Hubble Space Telescope Orbital System Test (HOST)

experiment that was launched October 29, 1998. HOST was a shuttle test

flight of upgraded systems for the Hubble Space Telescope, including a

cryocooler for the Near Infrared Camera and Multi-Object Spectrometer

(NICMOS). SAMS–FF measured the vibrations produced by the cryocooler

and the ability of the vibration isolation mounts for the cooler to conduct

these vibrations to the telescope. The Hubble Space Telescope has very

strict requirements for a low-vibration environment. SAMS–FF was chosen

to perform this task since it has small volume and weight, but high measurement resolution. The acceleration data produced by SAMS–FF were

analyzed and documented by the SAMS–FF team, and the Hubble team

used the report in quantifying the cryocooler vibration signature. The

SAMS–FF system, which was designed and developed by Lewis and its

contractors, is now manifested

for support of microgravity measurement on several different

platforms, including the KC–135,

sounding rockets, the space

shuttles, and the International

Space Station (ISS).

Find out more on the World Wide

Web: http://www.grc.nasa.gov/WWW/

MMAP/SAMSFF

Lewis contacts:

Ronald J. Sicker, (216) 433–6498,

Ronald.J.Sicker@grc.nasa.gov; and

Thomas J. Kacpura, (216) 977–1057,

Thomas.J.Kacpura@grc.nasa.gov

SAMS–FF–3: Hermetically enclosed

SAMS–FF system supporting HOST

payload on STS–95.

SAMS–FF–4: SAMS–FF control and

data acquisition unit in hermetic

enclosure.
170

Author: Thomas J. Kacpura

Headquarters program office: OLMSA (UG)

Programs/Projects: Microgravity Science, HST, SAMS–FF

Colloids are tiny (submicron) particles suspended in fluid. Paint, ink, and

milk are examples of colloids found in everyday life. The Binary Colloidal

Alloy Test (BCAT) is part of an extensive series of experiments planned to

investigate the fundamental properties of colloids so that scientists can

make colloids more useful for technological applications. Some of the

colloids studied in BCAT are made of two different sized particles (binary

colloidal alloys) that are very tiny, uniform plastic spheres. Under the

proper conditions, these colloids can arrange themselves in a pattern to

form crystals, which may have many unique properties. For example,

someday these colloidal crystals may form the basis of new classes of light

switches, displays, and optical devices that can fuel the evolution of the

next generation of computer and communication technologies. Windows

made of liquid crystals are already in the marketplace. These windows

change their appearance from transparent to opaque when a weak electric

current is applied. In the future, if the colloidal crystals can be made to

control the passage of light through them, such products could be made

much more cheaply. These experiments require the microgravity environment of space because good quality crystals are difficult to produce on

Earth because of sedimentation and convection in the fluid.

The BCAT experiment hardware included two separate modules for two

different experiments. The “Slow Growth” hardware consisted of a 35-mm

camera with a 250-exposure photo film cartridge. The camera was aimed

toward the sample module, which contained 10 separate colloid samples.

A rack of small lights provided backlighting for the photographs.

The BCAT hardware was launched on

the shuttle and was operated aboard

the Russian space station Mir by American astronauts John Blaha and David

Wolf (launched September 1996 and

returned January 1997; reflown September 1997 and returned January 1998).

To begin the experiment, one of these

astronauts would mix the samples to

disperse the colloidal particles and

break up any crystals that might have

already formed. Once the samples were

mixed and the experiment was powered

on, the hardware operated autonomously, taking photos of the colloidal

samples over a 90-day period.

The “Rapid Growth” hardware was

set up in the Microgravity Glovebox. This experiment used a video

camera with a low-magnification

lens to record five separate samples. The astronauts operated the

experiment by first mixing the

samples and then recording each

sample over a 24-hr period to

monitor the colloidal behavior.

Analysis of the BCAT data is in

progress and is almost complete.

Results indicate that it is indeed

possible to form new materials in

microgravity. Some samples in the

BCAT Slow Growth experiment

were observed to form ordered

crystalline structures in space even

though they do not do so in Earth’s

gravity. During the BCAT Rapid

Growth experiment, the colloidal

structure persisted through the

whole flight. The structure did not

collapse as it does on Earth.

BCAT results are providing important new information about the

behavior of several different colloidal systems. The investigation

proved that gravity plays a central

role in the formation and stability

of these types of colloidal crystal

structures. The investigation also

helped identify the optimum

conditions for the formation of

colloidal crystals, which will be

used for optimizing future microgravity experiments in the study

of colloidal physics. Dr. David Weitz

of the University of Pennsylvania

and Dr. Peter Pusey of the UniverBinary Colloidal Alloy Test Conducted on Mir

Colloidal crystal photographed by

astronaut David Wolf aboard Mir,

approximately 90 days after the

start of the BCAT experiment.
171

Space

BCAT Rapid Growth hardware. BCAT Slow Growth hardware.

sity of Edinburgh, United Kingdom, conceived the experiments. The hardware and software were designed, built, and tested by a contractor team

from NYMA, Inc., and Aerospace Design & Fabrication (ADF) at the NASA

Lewis Research Center in Cleveland, Ohio.

Find out more about BCAT on the World Wide Web:

http://zeta.grc.nasa.gov/expr4/bcat.htm

Find information about related research at the following web sites:

Colloidal Gelation (CGEL): http://zeta.grc.nasa.gov/expr3/cgelpage.htm

Physics of Hard Spheres Experiment (PhaSE): http://zeta.grc.nasa.gov/phase/

Physics of Colloids in Space (PCS): http://dept.physics.upenn.edu/~weitzlab/

research/nasaproj.html

Lewis contacts: Rafat R. Ansari

(Project Scientist), (216) 433–5008,

Rafat.R.Ansari@grc.nasa.gov; and

Monica I. Hoffmann (Project

Manager), (216) 433–6765,

Monica.I.Hoffmann@grc.nasa.gov

Authors: Monica I. Hoffmann and

Rafat R. Ansari

Headquarters program office:

OLMSA (UG)

Programs/Projects:

Microgravity Science, BCAT

Ostwald ripening, or coarsening, is a process in which large particles in

a two-phase mixture grow at the expense of small particles. It is a ubiquitous natural phenomena occurring in the late stages of virtually all phase

separation processes. In addition, a large number of commercially important alloys undergo coarsening because they are composed of particles

embedded in a matrix. Many of them, such as high-temperature superalloys used for turbine blade materials and low-temperature aluminum

alloys, coarsen in the solid state. In addition, many alloys, such as the

tungsten-heavy metal systems, coarsen in the solid-liquid state during

liquid phase sintering. Numerous theories have been proposed that

predict the rate at which the coarsening process occurs and the shape

of the particle size distribution. Unfortunately, these theories have never

been tested using a system that satisfies all the assumptions of the theory.

In an effort to test these theories, NASA studied the coarsening process in

a solid-liquid mixture composed of solid tin particles in a liquid lead-tin

matrix. On Earth, the solid tin particles float to the surface of the sample,

like ice in water. In contrast, in a

microgravity environment this does

not occur. The microstructures in

the ground- and space-processed

samples (see the photos) show

clearly the effects of gravity on the

coarsening process. The STS–83-

processed sample (right image)

shows nearly spherical uniformly

dispersed solid tin particles. In

contrast, the identically processed,

ground-based sample (left image)

shows significant density-driven,

nonspherical particles, and because

of the higher effective solid volume

fraction, a larger particle size after

the same coarsening time.

Coarsening in Solid-Liquid Mixtures Studied on the

Space Shuttle
172

Photomicrographs of 10 vol % solid tin phase taken after samples were

held 10 hr at 185 °C. Left: Ground experiment results show significant

density-driven sedimentation and nonspherical particles. Because of the

compacting of the particles (effective higher solid volume fraction) the

coarsening rate was artificially increased and the result was a larger

particle size. Right: STS–83 microgravity results show spherical, uniformly

dispersed solid tin particles in a lead tin eutectic.

The “Coarsening in Solid-Liquid Mixtures” (CSLM) experiment was conducted in the Middeck Glovebox facility (MGBX) flown aboard the shuttle

in the Microgravity Science Laboratory (MSL–1/1R) on STS–83/94. The

primary objective of CSLM is to measure the temporal evolution of the

solid particles during coarsening. The particles were coarsened four times

from 0 sec to 24 hours after a 9-min heat-up to 185 °C.

Quantitative analysis of the samples with a volume fraction of 10 percent

solid is yielding data that are as spectacular as the microstructures. The

data are so good that the kinetics and particle size distribution can be

determined with unprecedented accuracy. CSLM has provided the first

careful test of theory of Ostwald

ripening by conducting the study

at low solid volume fractions with

the solids uniformly distributed in

the sample.

CSLM hardware was developed

and built jointly by the NASA

Lewis Research Center and NYMA,

Inc. CSLM has been selected for

additional flight experiments as

well as continued ground studies.

Several new evaluation techniques have been developed, and

numerous research papers have

been published from this study

by the principal investigator, Prof.

Peter W. Voorhees, and his team.

For more information, visit us on

the World Wide Web: http://

zeta.grc.nasa.gov/cslm/cslmrslt.htm

Lewis contact:

John J. Caruso, (216) 433–3324,

John.J.Caruso@grc.nasa.gov

Author: John J. Caruso

Headquarters program office:

OLMSA (UG)

Programs/Projects:

HEDS, Microgravity Science, STS–83,

STS–94, MSL–1, MSL–1R, CSLM

Dendrites describe the treelike crystal morphology commonly assumed in

metals and alloys that freeze from supercooled or supersaturated melts.

There remains a high level of engineering interest in dendritic solidification

because the size, shape, and orientation of the dendrites determine the

final microstructure of a material. It is the microstructure that then determines the physical properties of cast or welded products.

Although it is well known that dendritic growth is controlled by the transport of latent heat from the moving solid-liquid interface, an accurate and

predictive model has not yet been developed. The effects of gravityinduced convection on the transfer of heat from the interface have prevented adequate testing, under terrestrial conditions, of solidification

models.

The Isothermal Dendritic Growth

Experiment (IDGE) constituted a

series of three microgravity experiments flown aboard the Space

Shuttle Columbia. The apparatus

was used to grow and record dendrite solidification in the absence

of gravity-induced convective heat

transfer, thereby producing a

wealth of benchmark-quality data

for testing solidification models

and theories. The principal investigator is Prof. Martin E. Glicksman

Third and Final Shuttle Mission of the Isothermal Dendritic

Growth Experiment Conducted—Highest Supercooling

Ever Recorded Achieved
173

Space

0 sec

60 sec

120 sec

30 sec

90 sec

Each color band

represents an 

advance of 30 sec

Temperature change, 

0.390 °K

~1 mm

of Rensselaer Polytechnic Institute in Troy, New York. The project is managed by the Microgravity Sciences Division at the NASA Lewis Research

Center and is supported by a contractor team of engineers and technicians

employed by Aerospace Design & Fabrication (ADF), Inc., under subcontract

to the NYMA, Inc.

The first two flights of IDGE (STS–62 and STS–75) aboard the Second and

Third United States Microgravity Payload (USMP–2 and –3) used ultrapure

succinonitrile (SCN) as the test material. SCN is an organic crystal that

mimics the solidification of body-centered-cubic metals, such as ferrous

metals. The third and final IDGE flight (USMP–4), which was launched on

STS–87 in November 1997, employed a different test material. This flight

used pivalic acid (PVA)—a face-centered-cubic organic crystal that solidifies

like many nonferrous metals. PVA, like SCN, has convenient properties,

such as its transparent nature and its low melting point, that make it ideal

for conducting benchmark experiments. Because a different sample was

used, the universality of solidification theories could be tested.

The third and final flight of IDGE

was a resounding success. Dendritic growths were obtained across

the full range of supercooling

required. In fact, during the mission, a growth was achieved at the

high end of the supercooling range,

1.25 K—something never before

accomplished even under terrestrial conditions. The data and subsequent analysis from the final flight

experiment are currently at a preliminary stage. Thus far, the results

indicate that dendritic growth in

PVA is (as in SCN) diffusion-limited,

with little, if any, kinetic response.

This observation conflicts with the

conclusion reached by other

investigators that there are large

interfacial kinetic effects in PVA.

In addition to the investigation of

dendritic solidification kinetics and

morphology, the IDGE has been

part of the development of remote,

university-based teleoperations.

These teleoperation tests point the

way to the future of microgravity

science operations on the International Space Station (ISS). NASA

Headquarters and the Telescience

Support Center at Lewis set a goal

for developing the experience and

expertise to set up remote, nonNASA locations from which to control space station experiments.

Recent IDGE space shuttle flights

provide proof-of-concept and

tests of remote space flight

teleoperations.

The IDGE flight series is now complete. The principal investigator and

his team are currently completing

analyses and moving toward final

data archiving. The IDGE published

results and archived data sets are

being used actively by other scientists and engineers. In fact, they

have been referenced in the classroom by a graduate-level instructor

at Case Western Reserve University.

In addition, the techniques and

IDGE hardware systems that were

Time-elapsed photographic sequence of a PVA dendrite grown during the

STS–87 mission. This figure is shown in color in the online version of this article

(http://www.grc.nasa.gov/WWW/RT1998/6000/6728malarik.htm).
174

Frame

grabber

Laserdisk

S-VHS tapedeck

Hi8 tapedeck

S-Video

S-Video

S-Video

Digital

1280 1024

Mekel film transport

Color 

wheel

Kodak 

Megaplus 1.4

Smart

serial

card

P6-200

3

4

5

6

Ethernet

developed are being employed for future dendritic growth flight

experiments.

Find out more about this research on the World Wide Web:

http://www.rpi.edu/locker/56/000756

Lewis contacts: Diane C. Malarik (Project Manager), (216) 433–3203,

Diane.C.Malarik@grc.nasa.gov; Terri D. Rodgers (Deputy Project Manager),

(216) 433–8740, Terri.D.Rodgers@grc.nasa.gov; and Dr. Francis D. Chiaramonte

(Project Scientist), (216) 433–8040, Francis.D.Chiaramonte@grc.nasa.gov

Author: Diane C. Malarik

Headquarters program office: OLMSA

Programs/Projects: Microgravity Science

Tracker is an object-tracking and image-processing program designed and

developed at the NASA Lewis Research Center to help with the analysis

of images generated by microgravity combustion and fluid physics experiments. Experiments are often recorded on film or videotape for analysis

later. Tracker automates the process of examining each frame of the

recorded experiment, performing image-processing operations to bring

out the desired detail, and recording the positions of the objects of interest.

It can load sequences of images from disk files or acquire images (via a

frame grabber) from film transports, videotape, laser disks, or a live camera. Tracker controls the image source to automatically advance to the

next frame. It can employ a large array of image-processing operations

to enhance the detail of the acquired images and can analyze an arbitrarily large number of objects simultaneously. Several different tracking

algorithms are available, including conventional threshold and

correlation-based techniques, and

more esoteric procedures such as

“snake” tracking and automated

recognition of character data in the

image. The Tracker software was

written to be operated by researchers, thus every attempt was made

to make the software as user

friendly and self-explanatory as

possible.

Tracker is used by most of the

microgravity combustion and fluid

physics experiments performed by

Lewis, and by visiting researchers.

This includes experiments performed on the space shuttles,

Mir, sounding rockets, zero-g

research airplanes, drop towers,

and ground-based laboratories.

This software automates the

analysis of the flame or liquid’s

physical parameters such as position, velocity, acceleration, size,

shape, intensity characteristics,

color, and centroid, as well as a

number of other measurements.

It can perform these operations

on multiple objects simultaneously.

Another key feature of Tracker is

that it performs optical character

recognition (OCR). This feature is

useful in extracting numerical

instrumentation data that are

Tracker—Image-Processing and Object-Tracking System

Developed

Object-tracking system hardware layout.
175

Space

embedded in images. All the results are saved in files for further data

reduction and graphing.

There are currently three Tracking Systems (workstations) operating

near the laboratories and offices of Lewis’ Microgravity Science Division

researchers. These systems are used independently by students, scientists, and university-based principal investigators. The researchers bring

their tapes or films to the workstation and perform the tracking analysis.

The resultant data files generated by the tracking process can then be analyzed on the spot, although most of the time researchers prefer to transfer

them via the network to their offices for further analysis or plotting. In addition, many researchers have installed Tracker on computers in their office

for desktop analysis of digital image sequences, which can be digitized by

the Tracking System or some other means. Tracker has not only provided

a capability to efficiently and automatically analyze large volumes of data,

saving many hours of tedious work, but has also provided new capabilities

to extract valuable information and phenomena that was heretofore

undetected and unexploited.

Find out about Tracker on the World Wide Web:

http://zeta.grc.nasa.gov/Tracker/Tracker.html

Lewis contacts: Robert B. Klimek, (216) 433–2837,

Robert.B.Klimek@grc.nasa.gov; and Dr. Theodore (Ted) W. Wright, (216) 433–5341,

Theodore.W.Wright@grc.nasa.gov

Authors: Robert B. Klimek and Dr. Theodore W. Wright

Headquarters program office: OLMSA (UG)

Programs/Projects:

Microgravity Science, Investigations of

the behavior of flames and fluids:

On the shuttles and Mir—SSCE, CFM,

WIF, ICE, FSDC, FFFT, RITSI, DCE, CHT,

LSP, ELF, TGDF, ALB, OFFS, Behavior of

Rapidly Sheared Bubbly Suspensions,

TIGER–3D, SIBAL

On sounding rockets—SAL, DARTFire

In aircraft and ground-based

facilities—Measurement of Flame

Temperatures in Reduced-Gravity and

other combustion experiments, Bubble

Formation and Detachment in Cross

Liquid Flow Under Reduced and

Normal Gravity and other fluid behavior

experiments, General Motors R&D

Center, Engine Research Department

Special recognition:

NASA Tech Brief; runnerup for NASA

Software of the Year award
177

Engineering and Technical Services
178

There are nearly 10 000 tracked objects orbiting the Earth. These objects

encompass manned objects, active and decommissioned satellites, spent

rocket bodies, and debris. They range from a few centimeters across to

the size of the MIR space station. Anytime a new satellite is launched, the

launch vehicle with its payload attached passes through an area of space in

which these objects orbit. Although the population density of these objects

is low, there always is a small but finite probability of collision between the

launch vehicle and one or more of these space objects. Even though the

probability of collision is very low, for some payloads even this small risk is

unacceptable.

To mitigate the small risk of collision associated with launching at an

arbitrary time within the daily launch window, NASA performs a prelaunch

mission assurance Collision Avoidance Analysis (or COLA). For the COLA

of the Cassini spacecraft, the NASA Lewis Research Center conducted an

in-house development and analysis of a model for launch collision probability. The model allows a minimum clearance criteria to be used with the

COLA analysis to ensure an acceptably low probability of collision. If, for

any given liftoff time, the nominal launch vehicle trajectory would pass a

space object with less than the minimum required clearance (see the

following figure), launch would not be attempted at that time.

The model assumes that the

nominal positions of the orbiting

objects and of the launch vehicle

can be predicted as a function

of time, and therefore, that any

tracked object that comes within

close proximity of the launch

vehicle can be identified. For any

such pair, these nominal positions

can be used to calculate a nominal

miss distance. The actual miss

distances may differ substantially

from the nominal miss distance,

due, in part, to the statistical

uncertainty of the knowledge of

the objects’ positions. The model

further assumes that these position

uncertainties can be described with

position covariance matrices. With

these, and some additional simplifying assumptions, a closed-form

solution for the probability of

collision is obtained. This solution

provides clear insights into how

each of the independent parameters affects the probability of

collision. It shows that for a given

maximum probability of collision

and prior knowledge of the objects’

position uncertainties and sizes,

only knowledge of the nominal

closest approach distance is

required to make the launch/no

launch decision.

The model was completed and

used for the mission assurance

COLA analysis for the Cassini

spacecraft, which was launched on

a Titan IV/Centaur rocket on

October 15, 1997. Although the

model was specifically developed

for the Cassini mission, it is clearly

applicable for other launches as

well.

z-axis

x-axis

y-axis

Trajectory of object 1

(e.q., launch vehicle)

Trajectory of object 2

(e.q., space object)

Direction of

relative motion

Minimum

approach

distance

NASA Lewis Launch Collision Probability Model Developed

and Analyzed

Launch vehicle and orbiting object nominal trajectories and miss distance.

Computer Services
179

Engineering and Technical Services

8:50 9:00 9:10 9:20 9:30 9:40 9:50 10:00 10:10 10:20 10:30 10:40 10:50 11:00

Window

close

11:03

Window

open

8:43

GMT:

This figure illustrates the effect of COLA closures on the launch window.

The bar represents the entire 140-min launch window on October 15,

1997; the blackened areas represent the loss of launch opportunities due

to the potential for a collision with an orbiting object.

Find out more on the World Wide Web:

Launch vehicle history: http://www.grc.nasa.gov/WWW/PAO/html/lvpo.htm

Cassini: http://www.jpl.nasa.gov/cassini

Bibliography

Bollenbacher, G.; and Guptill, J.D.: Launch Collision Probability. NASA TP—1999-

208852, 1999.

October 15 Cassini launch window (launch was not allowed in darkened areas because of the risk of collision).

GMT, Greenwich Mean Time.

Lewis contacts:

Gary Bollenbacher, (216) 433–2276,

Gary.Bollenbacher@grc.nasa.gov; and

James D. Guptill, (216) 433–5213,

James.D.Guptill@grc.nasa.gov

Authors:

Gary Bollenbacher and James D. Guptill

Headquarters program office: OSF

Programs/Projects: Cassini

A common problem occurs when refrigerant fluids wick inside the insulation of thermocouple wires through a compressor’s casing feedthrough

and then leak into the adjacent disconnect box outside the casing. Leaking

fluids create an unfavorable situation inside the disconnect box and may

contaminate the fluids. To address this problem, NASA Lewis Research

Center’s Manufacturing Engineering Division developed a customized

hermetic feedthrough for a bank of Worthington compressors.

Customized Hermetic Feedthrough Developed to

Isolate Fluids

Cutaway diagram of a customized hermetic feedthrough.

In these compressors, bearing

temperatures are measured by

internal thermocouples embedded

in bearings located inside the

compressor casings. The thermocouple wires need to be routed

outside the casing and read at

Wire size and type

determined by application

Number of wires 

determined by 

application

Shrink tubing 

strain relief (on 

both ends of tube)

Solid epoxy Staggered bare wire 

stripped, then encapsulated

with prep coat of epoxy

0.500-in Stainless

 steel tubing

Outer insulation

removed

Manufacturing Engineering
180

another location. These wires are short and are terminated to a disconnect strip inside the casing. The bearings operate at about 170 °F, but

because the casing is filled with R12 refrigerant oil, the casing has a

maximum temperature of about 100 °F.

The operating conditions of these compressors permit the use of an

epoxy that is compatible with the R12 fluid. The desired finished product

is a stainless steel tube that has been filled solid with epoxy after thermocouple wires bonded and sealed by epoxy have been inserted through

its length. Shrink tubing extends from both ends of the tube. The process

that was developed to isolate the thermocouple wires from the R12 fluid

follows.

For this application, use an 8-in.-long piece of 0.500-in. 304 stainless steel

tube with six pairs of 24-gauge stranded, PTFE-insulated (polytetrafluoroethylene) type “T” thermocouple wires for each feedthrough. Use shrink

tubing to strain relief the insulated wires at their exit from the stainless

steel tube.

Cut the wire to length and identify the location of the stainless steel tube

sleeve with masking tape. Then, remove the outer insulation from a 2-in.

section of wire that will be inside the tube, and carefully strip to bare

wire a 1-in. section in the middle of the section with the outer insulation

removed. For an effective seal, the epoxy must penetrate between the

strands when stranded conductors are used. Make the seal with epoxy

bond on the bare wire. The bare wire must be encapsulated with a thin

layer of the epoxy that leaves only a very low profile. These encapsulated

wires must cure before the assembly can be continued. Then, inspect the

cured wires for complete encapsulation before going to the next step.

Insert the wires in the stainless steel tube and orient them so that the

epoxied stripped sections are staggered within the tube (see the illustration); then, apply shrink tubing to

one end of the cleaned wires,

positioning it inside the edge of the

tube. The small gaps between the

wires on the other end will be used

to inject the epoxy into the tube.

Let the epoxy cure inside the tube,

free of any voids. Then, continue to

fill the tube until the entire 8-in.

length is nearly filled, allowing

room for the other strain-relieving

shrink tubing.

Since this first design, the process

has been adjusted to fit many

needs and situations. Customized feedthroughs have been

assembled from various wire types,

wire gauges, and/or stainless steel

tube passages. The fittings selected

to mount these feedthroughs allow

their use in other areas, such as

pressure or vacuum systems.

Lewis contact:

Roger D. Meredith, (216) 433–2103,

Roger.D.Meredith@grc.nasa.gov

Author: Roger D. Meredith

Headquarters program office: OAT

Programs/Projects:

Turbomachinery and combustion

research, wind tunnel research

Pyrotechnic shock, or pyroshock, is the transient response of a structure to

loading induced by the ignition of pyrotechnic (explosive or propellant

activated) devices. These devices are typically used to separate structural

systems (e.g., separate a spacecraft from a launch vehicle) and deploy

appendages (e.g., solar panels). Pyroshocks are characterized by high peak

acceleration, high-frequency content, and short duration. Because of their

high acceleration and high-frequency, pyroshocks can cause spaceflight

hardware to fail. Verifying by test that spaceflight hardware can withstand

the anticipated shock environment is considered essential to mission

success.

Pyroshock Environments Characterized for

Spacecraft Missions

The Earth Observing System (EOS)

AM–1 spacecraft for NASA’s Mission

to Planet Earth is scheduled to be

launched on an Atlas IIAS vehicle

in 1999, and the NASA Lewis

Research Center is the launch

vehicle integrator for this NASA

Goddard Space Flight Center

spacecraft. The EOS spacecraft was

subjected to numerous ground

Engineering Design and Analysis
181

Engineering and Technical Services

EOS spacecraft following shock separation testing at Lockheed Martin

Missiles & Space, Valley Forge, Pennsylvania.

shock tests to verify that its scientific instruments and avionics components

will withstand the shock-induced vibration produced when the spacecraft

separates from the launch vehicle. Shock test data from these tests represent the third largest available pyroshock data base in the United States.

Future spacecraft missions will directly benefit from the knowledge gained

from these tests.

The payload separation system used for EOS is a new system that operates

by firing six separation nuts. This system was tested to verify its functional

operation and to characterize the resulting shock levels. The launch vehicle contractor (Lockheed Martin Astronautics) and spacecraft contractor

(Lockheed Martin Missiles & Space) completed 16 separation test firings.

This resulted in an unusually large amount of pyroshock data. Typically,

only one or two pyroshock test firings are performed for a spacecraft

mission.

Because of the size of this separation system shock data base, engineers

were able to perform unique statistical analyses to characterize the distribution of the test data. For example, it was proven that the shock data

follow a lognormal distribution, a concept often assumed but rarely proven.

The test-to-test repeatability of the

shock source level was analyzed,

and the effects of various test

configurations and separation nut

production lots were examined and

quantified.

Engineers investigated the change

in shock level as the shock traveled

from the spacecraft separation

interface to the avionics components of the upper stage and

analyzed the effects of the structural fidelity (simulator versus real)

of the components and their

weight on vibrational response.

In addition, the shock attenuation

with distance and across joints was

quantified and compared with

concepts originally generated in

1970, and the effects of separation

nut preload and firing sequences

effects were examined.

Because of this EOS shock testing

and the analyses performed at

NASA Lewis, a significant amount

of new information on pyroshock

and its characteristics is now

available to the aerospace industry.

We hope that this information will

help future spacecraft test planners

to perform better and cheaper

spacecraft separation shock tests

and to better understand their test

data.

Find out more about EOS on the

World Wide Web:

http://eos-am.gsfc.nasa.gov/

spacecraft.html

http://eospso.gsfc.nasa.gov

http://www.earth.nasa.gov/

Bibliography

Hughes, W.O.; and McNelis, A.M.:

Statistical Analysis of a Large Sample

Size Pyroshock Test Data Set. NASA

TM—1998-206621, 1998.

Hughes, W.O.; and McNelis, A.M.:

Characterization of the Pyroshock

Environment for the EOS Mission,

Proceedings of the 44th IES Annual

Technical Meeting of the Institute of

Environmental Sciences and Technology,

Mount Prospect, IL, 1998, pp. 93–104.
182

Lewis contacts: William O. Hughes, (216) 433–2597,

William.O.Hughes@grc.nasa.gov; and Anne M. McNelis, (216) 433–8880,

Anne.M.McNelis@grc.nasa.gov

Authors: William O. Hughes and Anne M. McNelis

Headquarters program office: OES

Programs/Projects: MTPE, EOS

The Aero-Acoustic Propulsion Laboratory is an acoustically treated, 65-fttall dome located at the NASA Lewis Research Center. Inside this laboratory is the Nozzle Acoustic Test Rig (NATR), which is used in support of

Advanced Subsonics Technology (AST) and High Speed Research (HSR) to

test engine exhaust nozzles for thrust and acoustic performance under

simulated takeoff conditions. Acoustic measurements had been gathered

by a far-field array of microphones located along the dome wall and 10-ft

above the floor. Recently, it became desirable to collect acoustic data for

engine certifications (as specified by the Federal Aviation Administration

(FAA)) that would simulate the noise of an aircraft taking off as heard

from an offset ground location. Since nozzles for the High-Speed Civil

Transport have straight sides that cause their noise signature to vary

radially, an additional plane of acoustic measurement was required.

Desired was an arched array of 24 microphones, equally spaced from

the nozzle and each other, in a 25° off-vertical plane.

The various research requirements made this a challenging task. The

microphones needed to be aimed at the nozzle accurately and held firmly

in place during testing, but it was also essential that they be easily and

routinely lowered to the floor for calibration and servicing. Once serviced,

the microphones would have to be returned to their previous location

near the ceiling. In addition, there could be no structure between the

microphones and the nozzle, and any structure near the microphones

would have to be designed to minimize noise reflections. After many

concepts were considered, a single arched truss structure was selected that

would be permanently affixed to the dome ceiling and to one end of the

dome floor.

This structure, which was mostly installed in November 1997, forms a

track on which each microphone rides on its own “carriage.” The 150-ftlong, roller-coaster-like track consists of three 1-in.-diameter tubular rails,

a 4-in.-diameter structural tube, diagonal stiffeners, and web plates that

hold the components together. Twenty-four carriages constructed of light

aluminum tubing traverse in series along the track. Each carriage engages

all three rails and is shaped like a tripod, with wheels at each leg to

provide complete support to each microphone. This configuration also

Traversing Microphone Track Installed in NASA Lewis’

Aero-Acoustic Propulsion Laboratory Dome

allows each carriage to nest into

the next when the microphones

are brought to the ground for

calibration and servicing. As a tow

cable attached to the lead carriage

is pulled up by a winch on the

ground, the other carriages follow

in series and are properly spaced

along the cable. The tow cable is

guided along the truss by rollers on

each web. Most of the wheels and

rollers are roller-skate wheels—

a likely choice because of their

rugged urethane construction (like

that of the roller-coaster wheels

used at the Cedar Point amusement park in Sandusky, Ohio),

integral dual ball bearings, and

availability.

Each microphone extends from its

carriage on a “stinger” tube. These

stingers can be offset from the

carriage enough to place the microphones in different planes, up to

5° to either side of the 25° track

plane, extending the research

capability of the system. This new

traversing microphone system will

greatly increase the amount of

valuable research data that can be

collected during future nozzle tests

in this already highly productive

research facility.
183

Engineering and Technical Services

Find out more about the AAPL facility

on the World Wide Web:

http://www.grc.nasa.gov/WWW/Aapl/

traverse.htm

http://www.grc.nasa.gov/WWW/AFED/

facilities/aapl.html

http://www.grc.nasa.gov/WWW/Aapl/

traverse.htm

Lewis contacts:

Steven W. Bauman, (216) 433–3826,

Steven.W.Bauman@grc.nasa.gov, and

Gail P. Perusek, (216) 433–8729,

Gail.P.Perusek@grc.nasa.gov

Authors:

Steven W. Bauman and Gail P. Perusek

Headquarters program office: OAT

Programs/Projects: Propulsion

Systems R&T, HSR, HSCT, AST

Traversing microphone track assembly in Lewis’ Aero-Acoustic

Propulsion Laboratory.

The Virtual Interactive Classroom (VIC) allows Internet users, specifically

students, to remotely control and access data from scientific equipment.

This is a significant advantage to school systems that cannot afford experimental equipment, have Internet access, and are seeking to improve

science and math scores with current resources. The VIC concept was

initially identified as an application of Embedded Web Technology in June

of 1997. VIC uses the 1998 NASA Software of the Year Award winner,

Tempest, an embedded web server developed at the NASA Lewis Research

Center (ref. 1). VIC and Tempest were developed by the same civil servant

team at Lewis.

Virtual Interactive Classroom: A New Technology Developed

for Distance Learning

A VIC Development Lab was

established at Lewis to demonstrate that scientific equipment

can be controlled by remote users

over the Internet. Current projects

include a wind tunnel, a room

camera, a science table, and a

microscope:
184

· The wind tunnel project accesses the 6- by 6-in. laminar flow test section

of a research-grade wind tunnel. Users can read pressure transducers

located in the upper and lower surfaces of a wing as well as an airflow

transducer in the test section. Remote users can turn the airflow on and

off, set wind velocity, adjust the wing angle of attack, and read the pressure

values.

· The room camera project uses a charge-coupled discharge (CCD) device

mounted to a pan/tilt mechanism. Users can control the pan, tilt, zoom,

and focus. The image received by the user is refreshed at regular intervals.

· The science table project includes a horizontal surface with a two-axis

servomechanism that positions a color camera. The table has interchangeable surfaces. Several surfaces are available with fossils, Native American

artifacts, geology samples, and antique tools. The user, from any remote

location, can select a specific position for the camera and retrieve either

high-speed or high-quality images in real time.

· The microscope comprises a two-axis servomechanism that positions

a microscope camera and lens, and a microscope slide holder that can

contain up to 360 slides. Remote users can position the microscope over

any of the slides in the holder and download high-quality color images.

Metrics are being collected, as well as feedback from users, to determine

the desirability, popularity, and educational benefits of each project.

Modifications will be made in response to user feedback. After the projects

have been developed and tested, they will be duplicated at remote VIC

Operations Labs. These labs may be owned and/or operated by institutions

of learning. Students will be able to operate these remote labs to learn

about the Internet or study the subjects made available by VIC projects.

These opportunities are directly related to careers in science and engineering.

Virtual Interactive Classroom concept. Lewis’ Virtual Interactive Classroom leverages the vision

of “Internet in every classroom,” allowing students to remotely control scientific experiments.

Find out more about Lewis’ Virtual

Interactive Classroom on the World

Wide Web:

Learning Technologies Project:

http://www.grc.nasa.gov/WWW/K-12/

Embedded Web Technology:

http://vic.grc.nasa.gov/

Reference

1. Daniele, C.J.: Embedded Web

Technology: Internet Technology

Applied to Real-Time System Control.

Research & Technology 1997, NASA

TM—98-206312, 1998, pp. 165–166.

(Available online: http://

www.grc.nasa.gov/WWW/RT1997/

7000/7750daniele.htm)

Lewis contacts:

David W. York, (216) 433–3162,

David.W.York@grc.nasa.gov;

Maria Babula, (216) 433–5221,

Maria.Babula@grc.nasa.gov;

Lisa M. Lambert, (216) 433–3994,

Lisa.M.Lambert@grc.nasa.gov; and

Joseph G. Ponyik, (216) 433–8592,

Joseph.G.Ponyik@grc.nasa.gov

Authors:

David W. York and Maria Babula

Headquarters program office:

OLMSA, External Programs

Programs/Projects:

SEMAA, LTP, FCF, EWT

Special Recognition: Tempest received

the 1998 NASA Software of the Year

Award.
185

Commercial Technology
186

Through a Space Act Agreement, the NASA Lewis Research Center has

helped develop the base silicon carbide (SiC) epitaxial growth technology

for Westinghouse’s efforts to bring silicon carbide products to the marketplace. SiC is a high-temperature, high-voltage semiconductor that can

deliver greater than three times the power of conventional silicon devices.

The technology was initially disclosed in the 1994 R&T report (refs. 1 to 3).

NASA Lewis’ High Temperature Integrated Electronics and Sensors (HTIES)

team is developing SiC as a material for advanced semiconductor electronic device applications because SiC-based electronics and sensors can

operate in hostile environments where conventional silicon-based electronics cannot function.

SiC transmitters hold great promise for television stations because they can

convert broadcasts from analog to digital signals. A modular solid-state

design provides broadcasters with an option to gradually add modules,

increasing the power of their transmitters as they expand their highdefinition television (HDTV) coverage. Using these high-power transistors

will significantly reduce the space needed for high-power transmitters at

television stations and will offer a solid-state solution, reducing long-term

maintenance costs. Thus, transmitter manufacturers will be able to abandon their reliance on tube-based technology for high-power transmitters

and will be able to build smaller, high-power, solid-state transmitters.

Find out more about this research on the World Wide Web:

http://www.grc.nasa.gov/WWW/SiC/SiC.html

References

1. Neudeck, P.G.; and Powell, J.A.: Performance-Limiting Micropipe Defects

Identified in SiC Wafers. Research & Technology 1994, NASA TM–106764, 1995,

p. 8. (Available online: http://www.grc.nasa.gov/WWW/RT1994/Aeronautics/

02Instrumentation.html)

2. Neudeck, P.G.; and Larkin, D.J.: Silicon Carbide Junction Field-Effect Transistors

Developed. Research & Technology 1994, NASA TM–106764, 1995, pp. 10–11.

(Available online: http://www.grc.nasa.gov/WWW/RT1994/Aeronautics/

02Instrumentation.html)

3. Larkin, D.J., et al.: Site-Competition Epitaxy Controls Doping of Silicon Carbide.

Research & Technology 1994, NASA TM–106764, 1995, pp. 11–12. (Available

online: http://www.grc.nasa.gov/WWW/RT1994/Aeronautics/

02Instrumentation.html)

Silicon Carbide Being Developed for High-Definition

Television Transmitter Modules

SiC transmitter module.

Lewis contacts:

Donald E. Costello, (216) 433–6635,

fax (216) 433–2555,

Donald.E.Costello@grc.nasa.gov; and

Dr. Philip G. Neudeck, (216) 433–8902,

Philip.G.Neudeck@grc.nasa.gov

Author: Donald E. Costello

Headquarters program office: OAT

Programs/Projects: HTIES, HDTV
187

Commercial Technology

Engineers at the NASA Lewis Research Center will support the America

True design team led by America’s Cup innovator Phil Kaiko. The America’s

Cup is an international sailing competition, and its trophy is the oldest

in professional sports. America True is the San Francisco Yacht Club’s

challenger for the America’s Cup race in the year 2000. The joint effort

between NASA and America True is encouraged by Mission HOME, the

official public awareness campaign of the U.S. space community.

NASA Lewis and America True have entered into a Space Act Agreement

to focus on the interaction between the airfoil and the large deformation

of the pretensioned sails and rigs along with the dynamic motions related

to the boat motions. This work will require a coupled fluid and structural

simulation. Included in the simulation will be both a steady-state capability, to capture the quasi-state interactions between the air loads and sail

geometry and the lift and drag on the boat, and a transient capability, to

NASA Has Joined America True’s Design Mission for 2000

capture the sail/mast pumping

effects resulting from hull motions.

This agreement is segmented into

two phases of a fluid structure

interaction (FSI) simulation. FSI

simulations will enable validation

and improvements of the boat

design, with higher fidelity, more

reliable predictions of what will

happen during testing, and will

reduce the number of design iterations. Phase I is the generation of

an FSI simulation template. This

template will provide America True

with a baseline simulation that can

be used in generating subsequent

models. Phase II consists of production FSI simulations. In this

phase, the templates will be used

to simulate potential and final boat

designs.

Find out more on the World Wide

Web: http://www.americatrue.org

Lewis contacts:

Laurel J. Stauber, (216) 433–2820,

Laurel.J.Stauber@grc.nasa.gov;

Matthew E. Melis, (216) 433–3322,

Matthew.E.Melis@grc.nasa.gov

Author: Gynelle C. Steele

Headquarters program office: OAT

Programs/Projects:

Smart composites, hypersonics

The agreement between Lewis and America True calls for detailed analysis of

different sail and mast designs.

National Combustion Code, a Multidisciplinary

Combustor Design System, Will Be Transferred to the

Commercial Sector

The NASA Lewis Research Center and Flow Parametrics will enter into an

agreement to commercialize the National Combustion Code (NCC). This

multidisciplinary combustor design system utilizes computer-aided design

(CAD) tools for geometry creation, advanced mesh generators for creating

solid model representations, a common framework for fluid flow and structural analyses, modern postprocessing

tools, and parallel processing. This

integrated system can facilitate

and enhance various phases of

the design and analysis process.
188

Typical results for the National Combustion Code (NCC), a combustor design and analysis system.

can-combustor and a multishear

flow for a low-NOx fuel nozzle;

calculating a multiwalled production fuel nozzle and an IMFH/

Cyclone “1-cup sector” (which

contains one cyclone and an integrated set of mixing and flame

holder tubes); and providing computational support for tests of the

NASA LDI–MVS sector rig combustor (which uses lean direct injection and a multiple venture

swirler).

Lewis contacts:

Gynelle C. Steele, (216) 433–8258,

Gynelle.C.Steele@grc.nasa.gov; and

Dr. Nan-Suey Liu, (216) 433–8722,

Nan-Suey.Liu@grc.nasa.gov

Author: Gynelle C. Steele

Headquarters program office: OAT

Programs/Projects: Propulsion

Systems R&T, HPCCP, SGE, P&PM

Postprocessing

(CORPERF/PV3)

• Design parameters

• Pattern factor

• Radial temperature 

profile 

• Temperature rise

• Pressure loss

• NOX postprocessor

• PV3 visualization

Analysis

(CORSAIR)

• Improved parallel performance:

96-percent efficiency/

 24 processors

• Better communication

• Accelerated kinetics

• Spray model

• Turbulent combustion module

23 Speedup 1.53 Speedup

∅ ∅

∅ ∅ ∅

∅

∅

Z

6 holes 10 6 holes 20

DILUTION

RC=37

50 80 80

SWIRLER

FUEL

PRIMARY

28 15 40

14

120 CL

Design

(CAD)

• Geometry 

 modeling

• Commercial 

 CAD software

Preprocessing

(PATRAN/CFDGEOM)

• CAD file cleanup

• Unstructured mesh 

 generation

• Full combustion system

Overall speedup of 33

The National Combustion Code was developed under a NASA/Department

of Defense/Department of Energy/U.S. industry partnership. Recent efforts

have been focused on developing a computational combustion dynamics

capability that will meet combustor designer requirements for model

accuracy and analysis turnaround time, incorporating both short and longterm technology goals. As a first step, a baseline solver for turbulent combustion flows, CORSAIR–CCD, was developed under a joint modeling and

code development effort between the aeronautics industry and NASA

Lewis. CORSAIR–CCD is a Navier-Stokes flow solver based on an explicit

four-stage Runge-Kutta scheme that uses unstructured meshes and runs

on networked workstations. The solver can be linked to any computeraided design system via the Patran file system. Turbulence closure is

obtained via the standard k-ε model with a high Reynolds number wall

function. The following combustion models have been implemented into

the code: finite-rate chemical kinetics emulations for Jet–A and methane

fuels, turbulence-chemistry interactions via an assumed probability density

function for temperature fluctuations, and thermal emissions of nitrogen

oxides. CORSAIR–CCD can switch between a parallel virtual machine interface and a message-passing interface by using compiler flags. Its parallel

performance on several platforms has been analyzed; and on the basis of

the results, several improvements have been made. Applications of the

CORSAIR–CCD code to date include simulating swirling flow and simulating ignition-delay experiments; computing a generic swirling flow for a
189

Definitions of NASA Headquarters

Program Offices

OA Office of the Administrator

DDF Director’s Discretionary Fund

OAT Office of Aeronautics and Space Transportation Technology

OES Office of Earth Sciences

OLMSA Office of Life & Microgravity Sciences & Applications

LS Life Sciences

UG Microgravity Research

OSF Office of Space Flight

OSS Office of Space Science

ATMS Advanced Technology & Mission Studies
190

Definitions of Programs and Projects

ACESE Attitude Control and Energy Storage Experiment

ACTS Advanced Communications Technology Satellite

AITP Aerospace Industry Technology Program

ALB Angular Liquid Bridge Experiment

AOS Aviation Operation Systems

ASCR Advanced Subsonic Combustor Rig

AST Advanced Subsonic Technology

ASTP Advanced Space Transportation Program

ATD Advanced Technology Development

ATM Asynchronous Transfer Mode

AXAF Advanced X–Ray Astrophysics Facility

BCAT Binary Colloidal Alloy Test

CARES Ceramics Analysis and Reliability Evaluations of Structures

CETDP Cross Enterprise Technology Development Program

CFM Candle Flames in Microgravity Experiment

CHT Capillary-Driven Heat Transfer Investigation

CSLM Coarsening in Solid-Liquid Mixtures

CSOC Consolidated Space Operations Contract

D3 Direct Data Distribution

DARTFire Diffusive and Radiative Transport in Fires

DCE Droplet Combustion Experiment

EAPU Electric Auxiliary Power Unit

ELF Enclosed Laminar Flames

EMA Electromechanical Actuator

EOS Earth Observing System

EPM Enabling Propulsion Materials

EWT Embedded Web Technology

FCF Fluids & Combustion Facility for the International Space Station

FFFT Forced Flow Flame Spreading

FSDC Fiber Supported Droplet Combustion

GAS Get Away Special experiment

GEO Geosynchronous Earth orbit

HDTV High-definition television

HEDS Human Exploration and Development of Space

HEUS High Energy Upper Stage

HISTEC High Stability Engine Control

HITEMP Advanced High Temperature Engine Materials Technology Program

HOST Hubble Space Telescope Orbital System Test

HPCCP High Performance Computing and Communications Program

HREC High Reliability Engine Control

HSCT High-Speed Civil Transport

HSR High-Speed Research

HST Hubble Space Telescope

HTIES High Temperature Integrated Electronics and Sensors Program
191

ICE Interface Configuration Experiment

IMAPS International Microelectronics and Packaging Society

In-STEP In-Space Technology Experiment Program

ISS International Space Station

LCBT Low Cost Booster Technology Program

LEO Low Earth orbit

LSP Laminar Soot Processes Experiment

LTP Learning Technologies Project

MGBX Microgravity (or Middeck) Glovebox

Mir Russian space station

MSL Microgravity Science Laboratory

MSL–1R Microgravity Science Laboratory–First Reflight

MTPE Mission to Planet Earth

NCMR National Center for Microgravity Research

NCP National Cycle Program

NII/GII National and Global Information Infrastructure

NPSS Numerical Propulsion System Simulation

NSTAR NASA Solar Electric Propulsion Technology Application Readiness

OFFS Opposed Forced Flame Spread Experiment

PHSV Propulsion for Highly Survivable Vehicles

POSA Passive Optical Sample Assembly

P&PM Physics and Process Modeling

RITSI Radiative Ignition and Transition to Spread Investigation

RLV Reusable Launch Vehicles

SAL Spread Across Liquids

SAMS Space Acceleration Measurement System

SBIR Small Business Innovation Research

SDFD Solar Dynamic Flight Demonstration project

SD GTD Solardynamic Ground Test Demonstration

SEMAA Science/Engineering/Mathematics & Aerospace Academy

SGE Smart Green Engine

SIBAL Solid Inflammability Boundary at Low Speed

SOMO Space Operations Management Office

SSCE Solid Surface Combustion Experiment

SSP Space Solar Power

STS Space Transportation System

TGDF Turbulent Gas-Jet Diffusion Flames

TIGER–3D Transition From Ignition to Flame Growth Under External Radiation in 3D

TTCP The Technical Cooperative Program

TURBO–AE Three-dimensional viscous propulsion aeroelastic computer code

USML United States Microgravity Laboratory

USMP United States Microgravity Payload

WIF Wire Insulation Flammability experiment

X2000 Deep space development program

X–33 Replacement for space shuttles
192

A

Abdul-Aziz, Ali 117

Abel, Phillip 22

Adamczyk, Dr. John J. 94

Adamovsky, Dr. Grigory 58

Allman, Mark 67a

Altenkirch, Robert A. 149

Anderson, Bernhard H. 105

Ansari, Rafat R. 170

Anzic, Godfrey 84

Arend, David J. 107

Austin, Curt 124

B

Baaklini, George Y. 117, 130

Babula, Maria 183

Bailey, Dr. Sheila G. 34

Bakhle, Milind A. 132, 135b

Banks, Bruce A. 43, 44, 45, 47, 50

Baney-Barton, Elyse A. 47

Barrett, Charles A. 28

Bauer, Robert A. 148

Bauman, Steven W. 182

Bencic, Timothy J. 59

Berg, Robert F. 160

Bhasin, Dr. Kul B. 67b

Bhatacharjee, Subrata 149

Bhatt, Dr. Ramakrishna T. 16

Biaglow, James A. 37b

Binder, Michael P. 4

Bodis, James R. 130

Bollenbacher, Gary 178

Bozzolo, Guillermo 22

Brown, Dr. Gerald V. 133

Button, Robert M. 42b

C

Cairelli, James E. 40

Cannell, David S. 161

Carney, Dorothy V. 130

Caruso, John J. 171

Castelli, Michael G. 119, 120

Chang, Dr. Sin-Chung 114

Chato, David J. 109

Chiaramonte, Dr. Francis D. 172

Choi, Dr. Sung R. 122

Chu, Dr. Pong P. 90

Index of Authors and Contacts

Both authors and contacts are listed in this index. Articles start on the page numbers following the names.

When two articles start on the same page, the first article is indicated by an “a” after the number and the

second article by a “b.”

Chuang, Dr. Kathy C. 25

Cole, Gary L. 6

Costello, Donald E. 186

Curtis, Henry B. 34

D

Davis, Victoria 34

de Groh, Henry C., III 156

de Groh, Kim K. 44, 45, 46, 47, 50

Decker, Dr. Arthur J. 60

DellaCorte, Dr. Christopher 20

Dever, Joyce A. 48, 50

Dhadwal, Dr. Harbans S. 135a

Doherty, Michael P. 164

Draper, Susan L. 124

Dunlap, Patrick H., Jr. 143

Duval, Walter M.B. 158

F

Felder, James L. 4

Ferguson, Dr. Dale C. 35

Fite, E. Brian 60

Follen, Gregory J. 9

Fralick, Gustave C. 57

Fujikawa, Gene 93

Fusaro, Robert L. 138

G

Gaier, Dr. James R. 48, 51

Gardner, Barbara 34

Gaugler, Dr. Raymond E. 97

Geng, Steven M. 40

Gil, Christopher 125

Good, Brian 22

Guo, Dr. Ten-Huei 62

Guptill, James D. 178

Gyekenyesi, Dr. John P. 122

H

Hagedorn, Norman H. 36

Hammoud, Ahmad 53

Handschuh, Dr. Robert F. 21

Hansen, Patricia A. 45, 50

Hathaway, Dr. Michael D. 94

He, Charles 50

Hebsur, Dr. Mohan G. 14

Hegde, Uday 149

Heidmann, Dr. James D. 97

Himansu, Dr. Ananda 114

Hoffmann, Monica I. 170

Honecy, Frank 22

Hopkins, Dale A. 3

Hughes, Christopher E. 137

Hughes, William O. 180

Hunt, Patricia 47

I

Ivancic, William D. 68

J

Jacobson, Dr. Nathan S. 31

Jalics, Alice I. 48

Janosik, Lesley A. 123, 129

Jaworske, Dr. Donald A. 52

Jones, Robert E. 88b, 90

Jones, Dr. William H. 113

Jorgenson, Dr. Philip C.E. 114

Juhasz, Albert J. 42a

K

Kacpura, Thomas J. 168

Kascak, Albert F. 133

Kifle, Muli 91

Kiser, James D. 17

Klimek, Robert B. 174

Kory, Carol L. 71

Koudelka, John M. 151

Krainsky, Dr. Isay L. 73, 76

Krantz, Timothy L. 140

Kudlac, Maureen T. 110

Kurkov, Dr. Anatole P. 135a

Kwatra, Dr. Subhash C. 88b

L

Lambert, Lisa M. 183

Lauver, Richard W. 160

Lavelle, Thomas M. 3

Lee, Dr. Chi-Ming 98

Lee, Dr. Richard Q. 86

Lei, Dr. Jih-Fen 17

Lerch, Dr. Bradley A. 119, 124, 125

Lewandowski, Beth 8

Lewicki, Dr. David G. 141

Liou, Larry C. 99
193

Lissenden, Prof. Clifford J. 125

Litt, Jonathan S. 64

Liu, Dr. Nan-Suey 187b

Lock, James A. 161

Loh, Dr. Ching-Yuen 114

Long, Kenwyn J. 34

M

Malarik, Diane C. 172

Manthey, Lori A. 2

Martin, Lisa C. 17

McNelis, Anne M. 180

Meador, Dr. Michael A. 26

Mehmed, Oral 60

Melis, Matthew E. 187a

Meredith, Roger D. 179

Messenger, Scott 34

Meyer, William V. 161

Miller, Daniel N. 105

Miller, Dr. Robert A. 31

Miranda, Dr. Félix A. 79

Mital, Subodh K. 127

Moldover, Michael R. 160

Morales, Dr. Wilfredo 21

Morscher, Gregory N. 19

Motil, Susan M. 160

Murthy, Dr. Pappu L.N. 126, 127

N

Naiman, Cynthia G. 4, 9

Nathal, Dr. Michael V. 14, 124

Nemeth, Noel N. 129

Nesbitt, Dr. James A. 30

Neudeck, Dr. Philip G. 186

Nguyen, Hung D. 88a

Nguyen, Nam T. 92

Noebe, Ronald D. 22

O

Oberle, Lawrence G. 61

Olson, Sandra L. 149

Ovryn, Dr. Ben 163

P

Pai, Dr. Shantaram S. 126

Palaszewski, Bryan A. 101

Patnaik, Dr. Surya N. 3

Patterson, Michael J. 37a

Patterson, Richard L. 53

Paxson, Dr. Daniel E. 66

Pereira, Dr. J. Michael 124

Perusek, Gail P. 182

Petersen, Ruth A. 8

Plachta, David W. 112

Ponchak, Dr. George E. 74

Ponyik, Joseph G. 183

Potapczuk, Dr. Mark G. 104

Powers, Lynn M. 129

Q

Quintana, Jorge A. 93

R

Ratvasky, Thomas P. 102

Rauser, Richard W. 130

Rawlin, Vincent K. 39

Reehorst, Andrew L. 104

Ritzert, Frank J. 15

Rodgers, Terri D. 172

Rogers, Richard B. 164

Romanovsky, Robert R. 78

Roth, Dr. Don J. 130

Rutledge, Sharon K. 43, 44

S

Sacksteder, Kurt R. 151

Saunders, John D. 107

Schneider, Dr. Steven J. 37b, 38

Sechkar, Edward 47, 54

Shabbir, Aamir 94

Shah, Ashwin R. 127

Shaw, Dr. Robert J. 2

Sicker, Ronald J. 168

Simons, Dr. Rainee N. 86

Singh, Dr. Bhim S. 160

Singh, Dr. Mrityunjay 17

Skarda, Dr. J. Raymond Lee 166

Skoch, Gary J. 95

Smart, Anthony E. 161

Smith, Daniela C. 46

Soeder, James F. 42b

Sovey, James S. 39

Srivastava, Rakesh 132, 135b

Stauber, Laurel J. 187a

Steele, Gynelle C. 187a, 187b

Stefko, George L. 132, 135b

Steinetz, Dr. Bruce M. 143

Stocker, Dennis P. 153, 154

Stueber, Thomas J. 54

Sullivan, Roy M. 131

Suresh, Ambady 6

Sutter, Dr. James K. 120

T

Taylor, Thomas W. 161

Thieme, Lanny G. 55

Thorp, Scott A. 60

Tin, Padetha 161

Tolbert, Carol M. 56

Towne, Dr. Charles E. 108

Townsend, Jacqueline A. 48, 50

Townsend, Scott 6

Tuma, Dr. Margaret L. 61

V

Vaden, Karl R. 76

Valco, Dr. Mark J. 20

Vancil, Bernhard K. 82

Van Keuls, Dr. Frederick W. 78

Van Zante, Judith Foss 102

Vasuvedan, Dr. A.K. 14

Veres, Joseph P. 10

Vergilii, Franklin 151

vonDeak, Thomas C. 69

W

Wallett, Dr. Thomas M. 70

Walters, Robert 34

Wang, Len 50

Wang, Dr. Xiao-Yen 114

Wellborn, Steven R. 94

Wernet, Dr. Mark P. 95

Wey, Dr. Changlie 98

Wilson, Dr. Jeffrey D. 80, 81

Wintucky, Edwin. G. 82

Woodward, Richard P. 137

Wright, Dr. Theodore, W. 174

Y

Yegendu, Ferit 69

York, David W. 183

Yu, Dr. Sheng-Tao 114

Z

Zhu, Dr. Dongming 31

Zimmerli, Gregory A. 160
April 1999

NASA TM—1999-208815

E–11414

None

This publication is available from the NASA Center for AeroSpace Information, (301) 621–0390.

208

A10

REPORT DOCUMENTATION PAGE

2. REPORT DATE

19. SECURITY CLASSIFICATION

 OF ABSTRACT

18. SECURITY CLASSIFICATION

 OF THIS PAGE

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,

gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson

Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102

Form Approved

OMB No. 0704-0188

12b. DISTRIBUTION CODE

8. PERFORMING ORGANIZATION

 REPORT NUMBER

5. FUNDING NUMBERS

3. REPORT TYPE AND DATES COVERED

4. TITLE AND SUBTITLE

6. AUTHOR(S)

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

11. SUPPLEMENTARY NOTES

12a. DISTRIBUTION/AVAILABILITY STATEMENT

13. ABSTRACT (Maximum 200 words)

14. SUBJECT TERMS

17. SECURITY CLASSIFICATION

 OF REPORT

16. PRICE CODE

15. NUMBER OF PAGES

20. LIMITATION OF ABSTRACT

Unclassified Unclassified

Technical Memorandum

Unclassified

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135–3191

1. AGENCY USE ONLY (Leave blank)

10. SPONSORING/MONITORING

 AGENCY REPORT NUMBER

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546–0001

Research & Technology 1998

Aeronautics; Aerospace engineering; Space flight; Space power; Materials; Structures;

Electronics; Space experiments; Technology transfer

Unclassified -Unlimited

Subject Categories: 01 and 31 Distribution: Nonstandard

Responsible person, Walter S. Kim, organization code 9400, (216) 433–3742.

This report selectively summarizes the NASA Lewis Research Center’s research and technology accomplishments for the fiscal

year 1998. It comprises 134 short articles submitted by the staff scientists and engineers. The report is organized into five major

sections: Aeronautics, Research and Technology, Space, Engineering and Technical Services, and Commercial Technology. A

table of contents and an author index have been developed to assist readers in finding articles of special interest. This report is not

intended to be a comprehensive summary of all the research and technology work done over the past fiscal year. Most of the work

is reported in Lewis-published technical reports, journal articles, and presentations prepared by Lewis staff and contractors. In

addition, university grants have enabled faculty members and graduate students to engage in sponsored research that is reported at

technical meetings or in journal articles. For each article in this report, a Lewis contact person has been identified, and where

possible, reference documents are listed so that additional information can be easily obtained. The diversity of topics attests to the

breadth of research and technology being pursued and to the skill mix of the staff that makes it possible. At the time of publication, NASA Lewis was undergoing a name change to the NASA John H. Glenn Research Center at Lewis Field. For more

information about research at NASA Glenn, visit us on the World Wide Web (http://www.grc.nasa.gov). This document is

available on the World Wide Web (http://www.grc.nasa.gov/WWW/RT/). For publicly available reports, visit the Glenn Technical

Report Server (GLTRS) on the World Wide Web (http://gltrs.grc.nasa.gov/GLTRS/).